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Fabrication of Wind Tunnel and Testing Report

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Wind tunnel is becoming a tool extensively used for predicting dynamic behavior of objects
when it is submerged in a flow field. Wind tunnel is used for duplicating the flow environment around a body. It is being extensively used for analysis of structures, objects
and bodies when acted upon by flow, particularly air. Relative motion between air and the object is established inside a closed chamber by means of blowing air at high velocity around
object. Object of different shapes when placed at different orientation, different flow pattern are set up around the object. These flow patterns greatly affect the force acting on the body.
Practically encountered flying objects like aeroplane has lot to do with forces acting on it.
Force on the body determines the energy needed to propel it through the flow field. Hence it has become extremely important to know how much force acts on a body submerged in a
flowing fluid and how to reduce the force and the effect of changes in the orientations and shape of the object on the force and hence to find the drag acting on the body. Thus by
knowing the physics of flow over a body one can always put an effort to minimize the energy consumption for propelling a flying object.

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Fabrication of Wind Tunnel and Testing Report

  1. 1. VIVEKANANDA COLLEGE OF ENGINEERING AND TECHNOLOGY (Affiliated to Visvesvaraya Technological University, Belgaum) PUTTUR (DK) – 574 203 Department of Mechanical Engineering CERTIFICATE Certified that the Project Work titled ‘FABRICATION OF WIND TUNNEL AND TESTING’ is carried out by ABY JOYCE 4VP09ME004 AMRITH A V 4VP09ME007 DIKSHITH KUMAR P 4VP09ME024 JAYANTHA M 4VP09ME036 a bonafide student of Vivekananda College of Engineering and Technology, in partial fulfillment for the award of the degree of Bachelor of Engineering in Mechanical Engineering of Visvesvaraya Technological University, Belgaum during the year 2012- 2013. It is certified that all the corrections/ suggestions indicated for Internal Assessment have been incorporated in the report. The report has been approved as it satisfies the academic requirements in respect of Project Work prescribed for the said Degree. Prof. Chandrakantha Bekal Prof. Sudarshan Rao Dr. Ashok Kumar T Guide and Professor Professor and Head Principal Dept. of Mechanical Engineering Dept. of Mechanical Engineering Signature with date and seal: External Viva Name of the Examiners: Signature with Date 1. 2.
  2. 2. VIVEKANANDA COLLEGE OF ENGINEERING AND TECHNOLOGY (Affiliated to Visvesvaraya Technological University, Belgaum) PUTTUR (DK) – 574 203 Department of Mechanical Engineering CERTIFICATE Certified that the Project Work titled ‘Fabrication of Wind Tunnel and Testing’ is carried out by Mr. Jayantha M, USN: 4VP09ME036, a bonafide student of Vivekananda College of Engineering and Technology, in partial fulfillment for the award of the degree of Bachelor of Engineering in Mechanical Engineering of Visvesvaraya Technological University, Belgaum during the year 2012-2013. It is certified that all the corrections/ suggestions indicated for Internal Assessment have been incorporated in the report. The report has been approved as it satisfies the academic requirements in respect of Project Work prescribed for the said Degree. Prof. Chandrakantha Bekal Prof. Sudarshan Rao Dr. Ashok Kumar T Guide and Assistant Professor Professor and Head Principal Dept. of Mechanical Engineering Dept. of Mechanical Engineering Signature with date and seal: External Viva Name of the Examiners: Signature with Date 1. 2.
  3. 3. VIVEKANANDA COLLEGE OF ENGINEERING AND TECHNOLOGY (Affiliated to Visvesvaraya Technological University, Belgaum) PUTTUR (DK) – 574 203 Department of Mechanical Engineering CERTIFICATE Certified that the Project Work titled ‘Fabrication of Wind Tunnel and Testing’ is carried out by Mr. Aby Joyce, USN: 4VP09ME004, a bonafide student of Vivekananda College of Engineering and Technology, in partial fulfillment for the award of the degree of Bachelor of Engineering in Mechanical Engineering of Visvesvaraya Technological University, Belgaum during the year 2012-2013. It is certified that all the corrections/ suggestions indicated for Internal Assessment have been incorporated in the report. The report has been approved as it satisfies the academic requirements in respect of Project Work prescribed for the said Degree. Prof. Chandrakantha Bekal Prof. Sudarshan Rao Dr. Ashok Kumar T Guide and Assistant Professor Professor and Head Principal Dept. of Mechanical Engineering Dept. of Mechanical Engineering Signature with date and seal: External Viva Name of the Examiners: Signature with Date 1. 2.
  4. 4. VIVEKANANDA COLLEGE OF ENGINEERING AND TECHNOLOGY (Affiliated to Visvesvaraya Technological University, Belgaum) PUTTUR (DK) – 574 203 Department of Mechanical Engineering CERTIFICATE Certified that the Project Work titled ‘Fabrication of Wind Tunnel and Testing’ is carried out by Mr. Amrith A V, USN: 4VP09ME007, a bonafide student of Vivekananda College of Engineering and Technology, in partial fulfillment for the award of the degree of Bachelor of Engineering in Mechanical Engineering of Visvesvaraya Technological University, Belgaum during the year 2012-2013. It is certified that all the corrections/ suggestions indicated for Internal Assessment have been incorporated in the report. The report has been approved as it satisfies the academic requirements in respect of Project Work prescribed for the said Degree. Prof. Chandrakantha Bekal Prof. Sudarshan Rao Dr. Ashok Kumar T Guide and Assistant Professor Professor and Head Principal Dept. of Mechanical Engineering Dept. of Mechanical Engineering Signature with date and seal: External Viva Name of the Examiners: Signature with Date 1. 2.
  5. 5. VIVEKANANDA COLLEGE OF ENGINEERING AND TECHNOLOGY (Affiliated to Visvesvaraya Technological University, Belgaum) PUTTUR (DK) – 574 203 Department of Mechanical Engineering CERTIFICATE Certified that the Project Work titled ‘Fabrication of Wind Tunnel and Testing’ is carried out by Mr. Dikshith Kumar P, USN: 4VP09ME024, a bonafide student of Vivekananda College of Engineering and Technology, in partial fulfillment for the award of the degree of Bachelor of Engineering in Mechanical Engineering of Visvesvaraya Technological University, Belgaum during the year 2012-2013. It is certified that all the corrections/ suggestions indicated for Internal Assessment have been incorporated in the report. The report has been approved as it satisfies the academic requirements in respect of Project Work prescribed for the said Degree. Prof. Chandrakantha Bekal Prof. Sudarshan Rao Dr. Ashok Kumar T Guide and Assistant Professor Professor and Head Principal Dept. of Mechanical Engineering Dept. of Mechanical Engineering Signature with date and seal: External Viva Name of the Examiners: Signature with Date 1. 2.
  6. 6. i    Acknowledgements ________________________________________________    It is our pleasure to express our heartfelt thanks to Prof. Chandrakantha Bekal, Assistant Professor, Department of Mechanical Engineering, for his supervision and guidance which enabled us to understand and develop this project. We are indebted to Dr. Ashok Kumar T, Principal, and Prof. Sudarshan Rao, Head of the Department, for their advice and suggestions at various stages of the work. Special thanks go to the Management of VCET for providing us with a good study environment and laboratories facilities. Besides, we appreciate the support and help rendered by the teaching and non-teaching staff of Mechanical Engineering. Lastly, we take this opportunity to offer our regards to all of those who have supported us directly or indirectly in the successful completion of this project work. Aby Joyce Amrith A V Dikshith Kumar P Jayantha M                  
  7. 7. ii      ABSTRACT   Wind tunnel is becoming a tool extensively used for predicting dynamic behavior of objects when it is submerged in a flow field. Wind tunnel is used for duplicating the flow environment around a body. It is being extensively used for analysis of structures, objects and bodies when acted upon by flow, particularly air. Relative motion between air and the object is established inside a closed chamber by means of blowing air at high velocity around object. Object of different shapes when placed at different orientation, different flow pattern are set up around the object. These flow patterns greatly affect the force acting on the body. Practically encountered flying objects like aero plane has lot to do with forces acting on it. Force on the body determines the energy needed to propel it through the flow field. Hence it has become extremely important to know how much force acts on a body submerged in a flowing fluid and how to reduce the force and the effect of changes in the orientations and shape of the object on the force and hence to find the drag acting on the body. Thus by knowing the physics of flow over a body one can always put an effort to minimize the energy consumption for propelling a flying object. Our project aims at developing such a wind tunnel which may help us to analyze the forces generated by the air on various bodies particularly lift. A simple stain gauge is used for the measurement of lift on different shaped and differently oriented objects. Smoke is used to qualitatively study the flow over the body. An open type wind tunnel is fabricated based on available data, using simple fabricating techniques. Testing will be done on different objects. Different variable such as surface roughness, angle of attack, temperature and flow velocity are varied and its effect on lift are studied. Finally the results are plotted on graph to show the variations of lift force with variety of parameters like flow velocity, temperature and surface finish.   
  8. 8. iii    CONTENTS Page No. Acknowledgements i Abstract ii Contents iii List of Figures vi List of Tables viii List of Acronyms and Abbreviations ix Chapter 1 Introduction 01 1.1 Theory of operation 01 1.2 Smoke flow visualization system 02 1.3 Miniature wind tunnel 02 1.4 Classification of wind tunnel 03 1.4.1 Open wind tunnel 03 1.4.2 Closed wind tunnel 04 1.4.3 Subsonic wind tunnel 04 1.4.4 Transonic wind tunnel 04 1.4.5 Supersonic wind tunnel 05 1.4.6 Hypersonic wind tunnel 06 1.4.7 Hot shot wind tunnel 06 1.4.8 High Reynolds’ wind tunnels 07 1.4.9 V/STOL wind tunnel 07 1.4.10 Spin wind tunnel 07 1.4.11 Automobile wind tunnel 07 1.4.12 Aeroacoustic wind tunnel 08
  9. 9. iv    1.5 List of wind tunnels 08 1.6 Types of testing in wind tunnel 1.6.1 Aqua Dynamic Flume 09 1.6.2 Low speed oversize liquid testing 09 1.6.3 Fan testing 09 1.6.4 Wind engineering testing 09 Chapter 2 Literature Review 10 2.1 Origins 10 2.2 World War Two 12 2.3 Post World War Two 12 Chapter 3 Basics of Aerodynamics 15 3.1 Introductory terminology 16 3.1.1 Lift 16 3.1.1.1 Lift coefficient 17 3.1.2 Drag 18 3.1.2.1 Drag coefficient 19 3.1.2.2 Calculation of Drag and Lift force 20 3.1.3 Reynolds’ Number 21 3.1.4 Mach Number 22 3.2 Boundary Layer 23 3.3 Airfoil 23 3.3.1 Angle of attack 25 Chapter 4 Construction and Fabrication 26 4.1 Introduction 26 4.2 Set-up 27 4.3 Description of components 28 4.3.1 Nozzle 28 4.3.2 Diffuser 28
  10. 10. v    4.3.3 Test section 29 4.3.4 Outlet and Inlet fan 30 4.3.5 Step down transformer 30 4.3.6 Honeycomb 31 4.3.7 Heating coil 32 4.3.8 Force measuring sensor 32 4.4 Working of the set-up 33 Chapter 5 Testing and Results 34 5.1 Testing and testing method 34 5.2 Effect of temperature on airfoil lift 35 5.2.1 Effect of temperature on airfoil lift for temperature 32.8ºC 35 5.2.2 Graphs for airfoil models for temperature of 32.8ºC 37 5.2.3 Effect of temperature on airfoil lift for temperature of 51.6ºC 39 5.2.4 Graphs for airfoil models for temperature of 51.6ºC 41 5.3 Effect of velocity of air on airfoil lift 43 5.3.1 Effect of velocity of air on airfoil lift for velocity 6.7m/s 43 5.3.2 Graphs for airfoil models for velocity of 6.7m/s 45 5.3.3 Effect of velocity of airfoil lift for velocity 3.8m/s 47 5.3.4 Graphs for airfoil models for velocity 3.8m/s 49 Chapter 6 Conclusion and scope for further work 51 Reference 53 Cost Details 54 Personal Profile 55
  11. 11. vi    LIST OF FIGURES ________________________________________________  Page No. Figure 1.1 Schematic of Eiffel type open wind tunnel 03 Figure 1.2 Closed circuit or return flow low speed wind tunnel 04 Figure 1.3 Testing at transonic speeds 05 Figure 1.4 Supersonic Wind Tunnel at Lewis Flight Propulsion Laboratory 05 Figure 1.5 NASA Langley's Hypersonic Facilities Complex, 1969 06 Figure 1.6 Vertical wind tunnels T-105 at Central Aero hydrodynamic Institute, Moscow, built in 1941 for aircraft testing 08 Figure 2.1 Replica of Wright Brothers’ wind tunnel 11 Figure 2.2 German aviation laboratories 11 Figure 2.3 Preparing a model in the Kirsten Wind Tunnel, a subsonic wind tunnel at the University. 13 Figure 2.4 Fan blades of Langley Research Center's 16 foot transonic wind tunnel in 1990, before it was mothballed in 2004. 14 Figure 3.1 Forces acting on air foil. 16 Figure 3.2 Longer path theory 17 Figure 3.3 Measured drag coefficient 20 Figure 3.4 An F/A-18 Hornet at transonic speed and displaying a vapor cone just before reaching the speed of sound 22 Figure 3.5 Boundary layer visualization, showing transition from laminar to turbulent condition 23 Figure 3.6 Examples of airfoils in nature and within various vehicles 24 Figure 3.7 Angle of attack 25 Figure 4.1 Schematic diagram of wind tunnel 27 Figure 4.2 Nozzle 28 Figure 4.3 Diffuser 28 Figure 4.4 Test Section 29 Figure 4.5 Fan Housing 30 Figure 4.6 Step down transformer 31
  12. 12. vii    Figure 4.7 Honeycomb structures 31 Figure 4.8 Heating coil 32 Figure 4.9 Strain Gauge 32 Figure 5.1 Test set up 34 Figure 5.2 Symmetric airfoil for temperature 32.8ºC 37 Figure 5.3 Clark ‘Y’ airfoil for temperature 32.8ºC 37 Figure 5.4 Blackbird airfoil for temperature 32.8ºC 38 Figure 5.5 Turbofan airfoil for temperature 32.8ºC 38 Figure 5.6 Symmetric airfoil for temperature 51.6ºC 41 Figure 5.7 Clark ‘Y’ airfoil for temperature 51.6ºC 41 Figure 5.8 Blackbird airfoil for temperature 51.6ºC 42 Figure 5.9 Turbofan airfoil for temperature 51.6ºC 42 Figure 5.10 Symmetric airfoil for velocity of 6.7m/s 45 Figure 5.11 Clark ‘Y’ airfoil for velocity of 6.7m/s 45 Figure 5.12 Blackbird airfoil for velocity of 6.7m/s 46 Figure 5.13 Turbofan airfoil for velocity of 6.7m/s 46 Figure 5.14 Symmetric airfoil for velocity of 3.8m/s 49 Figure 5.15 Clark ‘Y’ airfoil for velocity of 3.8m/s 49 Figure 5.16 Blackbird airfoil for velocity of 3.8m/s 50 Figure 5.17 Turbofan airfoil for velocity of 3.8m/s 50
  13. 13. viii    LIST OF TABLES Page No. Table 3.1 Form drag and Skin friction for different shapes 18 Table 5.1 Effect of temperature on airfoil lift for temperature 32.8o C 35 Table 5.2 Effect of temperature on airfoil lift for temperature 51.6o C 39 Table 5.3 Effect of velocity of air on airfoil lift for velocity 6.7 m/s 43 Table 5.4 Effect of velocity of air on airfoil lift for velocity 3.8m/s 47
  14. 14. ix    LIST OF ACRONYMS AND ABBREVIATIONS AOA Angle Of Attack CFD Computational Fluid Dynamics HWT Hotshot Wind Tunnel LEED Leadership in Energy and Environment Design Ma/M Mach NASA National Aeronautics and Space Administration NASCAR National Association for Stock Car Auto Racing RPM Revolutions Per Minute SUV Sports Utility Vehicle V/STOL Vertical and/or Short Take-Off and Landing
  15. 15. 1    Chapter 1 INTRODUCTION A wind tunnel is a tool used in aerodynamic research to study the effects of air moving past solid objects. A wind tunnel consists of a closed tubular passage with the object under test mounted in the middle. A powerful fan system moves air past the object, the fan must have straightening vanes to smoothen the airflow. The test object is instrumented with a sensitive balance to measure the forces generated by airflow or, the airflow may have smoke or other substances injected to make the flow lines around the object visible. Full-scale aircraft or vehicles are sometimes tested in large wind tunnels, but these facilities are expensive to operate and some of their functions have been taken over by computer modeling. In addition to vehicles, wind tunnels are used to study the airflow around large structures such as bridges or office buildings. The earliest enclosed wind tunnels were invented in 1871[1] ; large wind tunnels were built during the Second World War. 1.1 Theory of operation Wind tunnels were first proposed as a means of studying vehicles (primarily airplanes) in free flight. The wind tunnel was envisioned as a means of reversing the usual paradigm, instead of the air's standing still and the aircraft moving at speed through it, the same effect would be obtained if the aircraft stood still and the air moved at speed past it. In that way a stationary observer could study the aircraft in action, and could measure the aerodynamic forces being imposed on the aircraft. Later on, wind tunnel study came into its own the effects of wind on manmade structures or objects needed to be studied when buildings became tall enough to present large surfaces to the wind, and the resulting forces had to be resisted by the building's internal structure. Determining such forces was required before building codes that could specify the required strength of such buildings and such tests continue to be used for large or unusual buildings. Still later, wind-tunnel testing was applied to automobiles, not so much to determine aerodynamic forces but more to determine ways to reduce the power required to move the vehicle on roadways at a given speed. In these studies, the interaction between the road
  16. 16. 2    and the vehicle plays a significant role, and this interaction must be taken into consideration when interpreting the test results. [4] In an actual situation the roadway is moving relative to the vehicle but the air is stationary relative to the roadway, but in the wind tunnel the air is moving relative to the roadway, while the roadway is stationary relative to the test vehicle. Some automotive-test wind tunnels have incorporated moving belts under the test vehicle in an effort to approximate the actual condition. 1.2 Smoke Flow Visualization System Smoke Flow Visualization has played an important role in understanding the theories and principles of Fluid Dynamics, Heat Transfer, and aerodynamics in general.[2] There have been many developments, and new technologies used to attempt to visualize the flow of air in and around objects. This project focuses on the design and development of a smoke- wire flow visualization wind tunnel. Smoke lines are produced in such a wind tunnel by the vaporization of mineral or other comparable oils on a thin wire. When oil is applied to the wire, it beads up, and when a current is run through the wire, the smoke lines are produced. 1.3 Miniature Wind Tunnel A Wind Tunnel is a tool to study and analyze the aerodynamics of various objects with different shape. Here instead of air standing still and the object is moving with high speed the object is kept stationary and air is allowed to pass over it. The effect would be same in both the cases. When an object is under the influence of air it experiences some kind of forces generated by the object. These forces are the drag, lift, thrust and self weight of the object. Many developments and technologies are developed inorder to measure and analyze the effects of these forces. This project focuses on the development of a miniature wind tunnel to measure and analyze the drag and lift of various objects. The lift and drag experienced by the object will be measured using simple strain gauge. To make the air flow laminar a honeycomb is also used so that accurate result can be obtained. The analysis is carried out taking into consideration of various parameters such as surface finish, angle of attack, flow velocity and temperature.
  17. 17. 1.4 C 1. B 2. B 3. B 1.4.1 In an but th Classificat Based on con i. Open ii. Closed Based on spe i. Low s a) b) ii. High a) b) c) Based on thei i. Ae a) b) c) i. Au ii. Ae Open wind n open-loop w his isn't very tion of Wi nstruction wind tunnel d wind tunn ed of operat speed wind t Subsonic w Transonic speed wind t Supersonic Hypersoni Hot shot w ir use eronautical w High Reyn V/STOL Spin tunne utomobile tu ero acoustic d tunnel Figure 1.1 wind tunnel, y economical nd Tunne l el ion tunnel wind tunnel wind tunnel tunnel c wind tunne ic wind tunn wind tunnel wind tunnel nolds’ numb el unnels tunnels Schematic o , the intake a l. Today, mo 3 el l el nel ber wind tunn of Eiffel type and exhaust ost wind tunn nel e open wind ends of the nels are clos tunnel. tunnel are n sed-loop. [3] not connectedd,
  18. 18. 1.4.2 It doe be re- close the tu return flow 1.4.3 Low the te return increa 1.4.4 High are de are ab the te super due to There walls boun Closed win Figu esn't take a -circulated in d-loop tunne unnel while n duct must in the test se Subsonic w speed wind est section u n flow. The ases the dyn Transonic subsonic w esigned on t ble to achiev est section. rsonic flow r o the reflect efore, perfor s. Since imp dary layer in nd tunnel ure 1.2 Clos rocket scien nto the intak el works. Sp minimizing be properly ection. wind tunnel tunnels are up to 400 km air is move namic pressu wind tunne ind tunnels the same prin ve speeds cl The Mach regions. Tes tion of the sh rated or slot portant visco nteraction) b sed circuit or ntist to see th ke end of the pecial vanes turbulence designed to used for ope m/h (~ 100 m ed with a pr ure to overco el (0.4 < M < nciples as th lose to the s number is a ting at trans hock waves tted walls a ous or invis both Mach a 4 r return flow hat fast-mov e tunnel to he are used to t and power l o reduce the erations at ve m/s, M = 0.3 ropulsion sy ome the visco 0.75) or tran he subsonic w speeds of so approximate onic speeds from the wa are required scid interacti and Reynold w low speed w ving air com elp boost tot turn the airfl loss. In a ret pressure los ery low mac 3).[3] They ar ystem made ous losses. nsonic wind wind tunnels und. The hi ely one with presents add alls of the tes to reduce s ions occur ( s’ number a wind tunnel. ming out of t tal airspeed. flow around t turn-flow wi sses and to e ch number, w re of open-r of a large tunnels (0.7 s. Transonic ghest speed h combined ditional prob st section (se shock reflect (such as sho are important the tunnel ca This is how the corners o ind tunnel th ensure smoot with speeds i return type, o axial fan th 75 < M < 1.2 c wind tunne is reached i subsonic an blems, mainl ee figure 1.3 tion from th ock waves o t and must b an w a of he th in or at 2) els in nd ly 3). he or be
  19. 19. prope used. 1.4.5 A sup The numb a high of 10 super has a opera erly simulate Supersonic Figure 1.4 personic win Mach numb ber is varied h pressure ra 0). [4] Apart rsonic wind a large power ation. ed. Large sca Fig c wind tunn 4 Supersonic nd tunnel is ber and flow d changing th atio is requir from that, tunnel needs r demand so ale facilities gure 1.3 Test el Wind Tunn a wind tun w are determ he density le red (for a su condensatio s a drying or o that most ar 5 and/or pres ting at transo nels at Lewis nnel that pro mined by th evel (pressur upersonic reg on or liquefa r a pre-heati re designed surized or cr onic speeds s Flight Prop oduces super he nozzle g re in the sett gime at M=4 action can o ing facility. A for intermitt ryogenic win pulsion Labo rsonic speed geometry. Th tling chambe 4, this ratio i occur. This A supersoni tent instead nd tunnels ar oratory. ds (1.2<M<5 he Reynold er). Therefor is of the orde means that c wind tunn of continuou re 5). ds’ re er a el us
  20. 20. 6    1.4.6 Hypersonic wind tunnel Figure 1.5 NASA Langley's Hypersonic Facilities Complex, 1969 A hypersonic wind tunnel is designed to generate a hypersonic flow field in the working section. The speed of these tunnels varies from Mach 5 to 15. As with supersonic wind tunnels, these types of tunnels must run intermittently with very high pressure ratios when initializing. Since the temperature drops with the expanding flow, the air inside has the chance of becoming liquefied. For that reason, preheating is particularly critical (the nozzle may require cooling). High pressure and temperature ratios can be produced with a shock tube. 1.4.7 Hot shot wind tunnel One form of HWT is known as a Gun Tunnel or hot shot tunnel (up to M=27), which can be used for analysis of flows past ballistic missiles, space vehicles in atmospheric entry, and plasma physics or heat transfer at high temperatures. [3] It runs intermittently, like other high speed tunnels, but has a very low running time (less than a second). The method of operation is based on a high temperature and pressurized gas (air or nitrogen) produced in an arc-chamber, and a near-vacuum in the remaining part of the tunnel. The arc-chamber can reach several MPa, while pressures in the vacuum chamber can be as low as 0.1 Pa. This means that the pressure ratios of these tunnels are in the order of 10 million. Also, the temperatures of the hot gas are up to 5000 K. The arc chamber is mounted in the gun barrel. The high pressure gas is separated by the vacuum by a diaphragm that breaks down as its resistance is exceeded.
  21. 21. 7    1.4.8 High Reynolds’ number wind tunnels Reynolds’ number is one of the governing similarity parameters for the simulation of flow in a wind tunnel. For Mach number less than 0.3, it is the primary parameter that governs the flow characteristics. There are three main ways to simulate high Reynolds’ number, since it is not practical to obtain full scale Reynolds’ number by use of a full scale vehicle. • Pressurized tunnels - Here test gases are pressurized to increase the Reynolds’ number. • Heavy gas tunnels - Heavier gases like Freon and R134a are used as test gases. The transonic dynamics tunnel at NASA Langley is an example of such a tunnel.[6] • Cryogenic tunnels - Here test gas is cooled down to increase the Reynolds’ number. The European transonic wind tunnel uses this technique. 1.4.9 V/STOL wind tunnels V/STOL tunnels require large cross section area, but only small velocities. Since power varies with the cube of velocity, the power required for the operation is also less. An example for a V/STOL tunnel is the NASA Langley 14' X 22' tunnel. 1.4.10 Spin wind tunnels Aircrafts have a tendency to go to spin when they stall (flight). These tunnels are used to study that phenomenon. 1.4.11 Automobile wind tunnels Automobile tunnels are of two categories; • External flow tunnels - Used to study the external flow through the chassis. • Climatic tunnels - Used to evaluate the performance of door systems, braking systems etc under various climatic conditions. Most of the leading automobile manufacturers have their own climatic wind tunnels.
  22. 22. 8    1.4.12 Aeroacoustic wind tunnels These tunnels are used in the studies of noise generated by flow and its suppression. 1.5 List of wind tunnels Following are the commercially build Wind Tunnel • Modine Wind Tunnel Climatic wind tunnel testing, large truck and automotive • AeroDyn Wind Tunnel, Full scale NASCAR racecars • A2 Wind Tunnel Full scale general purpose • Eight-Foot High Speed Tunnel • Full Scale 30- by 60-Foot Tunnel • Trisonic Wind Tunnel • Unitary Plan Wind Tunnel • Wind Shear's Full Scale, Rolling Road, Automotive Wind Tunnel • Variable Density Tunnel • European transonic wind tunnel[1] Figure 1.6 Vertical wind tunnel T-105 at Central Aero hydrodynamic Institute, Moscow, built in 1941 for aircraft testing
  23. 23. 9    1.6 Types of testing in wind tunnel 1.6.1 Aqua dynamic flume The aerodynamic principles of the wind tunnel work equally on watercraft, except the water is more viscous and so imposes greater forces on the object being tested. A looping flume is typically used for underwater aqua dynamic testing. The interaction between two different types of fluids means that pure wind tunnel testing is only partly relevant. However, a similar sort of research is done in a towing tank. 1.6.2 Low-speed oversize liquid testing Air is not always the best test medium to study small-scale aerodynamic principles, due to the speed of the air flow and airfoil movement. A study of fruit fly wings designed to understand how the wings produce lift was performed using a large tank of mineral oil and wings 100 times larger than actual size, in order to slow down the wing beats and make the vortices generated by the insect wings easier to see and understand.[2] 1.6.3 Fan testing Wind tunnel tests are also performed to measure the air movement of the fans at a specific pressure exactly. By determining the environmental circumstances during the measuring and by revising the air-tightness afterwards, the standardization of the data is warranted. There are two possible ways of measurement: a complete fan or an impeller on a hydraulic installation. Two measuring tubes enable measurements of lower air currents (< 30.000 m³/h) as well as higher air currents (< 60.000 m³/h).[2] The determination of the Q/h curve of the fan is one of the main objectives. 1.6.4 Wind engineering testing In Wind Engineering, wind tunnel tests are used to measure the velocity around, and forces or pressures upon structures. Very tall buildings, buildings with unusual or complicated shapes (such as a tall building with a parabolic or a hyperbolic shape), cable suspension bridges or cable stayed bridges are analyzed in specialized atmospheric boundary layer wind tunnels. These feature a long upwind section to accurately represent the wind speed and turbulence profile acting on the structure.
  24. 24. 10    Chapter 2 LITERATURE REVIEW 2.1 Origins English military engineer and mathematician Benjamin Robins (1707–1751) invented a whirling arm apparatus to determine drag and did some of the first experiments in aviation theory. Sir George Cayley (1773–1857) also used a whirling arm to measure the drag and lift of various airfoils. His whirling arm was 5 feet (1.5 m) long and attained top speeds between 10 and 20 feet per second. [1] However, the whirling arm does not produce a reliable flow of air impacting the test shape at a normal incidence. Centrifugal forces and the fact that the object is moving in its own wake mean that detailed examination of the airflow is difficult. Francis Herbert Wenham (1824–1908), a Council Member of the Aeronautical Society of Great Britain, addressed these issues by inventing, designing and operating the first enclosed wind tunnel in 1871. Once this breakthrough had been achieved, detailed technical data was rapidly extracted by the use of this tool. Wenham and his colleague Browning are credited with many fundamental discoveries, including the measurement of l/d ratios, and the revelation of the beneficial effects of a high aspect ratio. Carl Rickard Nyberg used a wind tunnel when designing his Flugan from 1897 and onwards. [1] In a classic set of experiments, the Englishman Osborne Reynolds (1842–1912) of the University of Manchester demonstrated that the airflow pattern over a scale model would be the same for the full-scale vehicle if a certain flow parameter were the same in both cases. This factor, now known as the Reynolds Number, is a basic parameter in the description of all fluid-flow situations, including the shapes of flow patterns, the ease of heat transfer, and the onset of turbulence. This comprises the central scientific justification for the use of models in wind tunnels to simulate real-life phenomena. However, there are limitations on conditions in which dynamic similarity is based upon the Reynolds number alone.
  25. 25. 11    Figure 2.1 Replica of Wright Brothers’ wind tunnel Figure 2.2 German aviation laboratories, 1935 The Wright brothers' use of a simple wind tunnel in 1901 to study the effects of airflow over various shapes while developing their Wright Flyer was in some ways revolutionary. It can be seen from the above, however, that they were simply using the accepted technology of the day, though this was not yet a common technology in America.[6] Subsequent use of wind tunnels proliferated as the science of aerodynamics and discipline of aeronautical engineering were established and air travel and power were developed.
  26. 26. 12    The US Navy in 1916 built one of the largest wind tunnels in the world at that time at the Washington Navy Yard. The inlet was almost 11 feet (3.4 m) in diameter and the discharge part was 7 feet (2.1 m) in diameter. A 500 hp electric motor drove the paddle type fan blades. Until World War Two, the world's largest wind tunnel was built in 1929 and located in a suburb of Paris, Chalais-Meudon, France. It was designed to test full size aircraft and had six large fans driven by high powered electric motors. 2.2 World War Two In 1941 the US constructed one of the largest wind tunnels at that time at Wright Field in Dayton, Ohio. This wind tunnel starts at 45 feet (14 m) and narrows to 20 feet (6.1 m) in diameter. Two 40-foot (12 m) fans were driven by a 40,000hp electric motor. Large scale aircraft models could be tested at air speeds of 400 mph (640 km/h). The wind tunnel used by German scientists at Peenemunde prior to and during WWII is an interesting example of the difficulties associated with extending the useful range of large wind tunnels. It used some large natural caves which were increased in size by excavation and then sealed to store large volumes of air which could then be routed through the wind tunnels. This innovative approach allowed lab research in high-speed regimes and greatly accelerated the rate of advance of Germany's aeronautical engineering efforts. By the end of the war, Germany had at least three different supersonic wind tunnels, with one capable of Mach 4.4 (heated) airflows. [6] 2.3 Post World War Two Later research into airflows near or above the speed of sound used a related approach. Metal pressure chambers were used to store high-pressure air which was then accelerated through a nozzle designed to provide supersonic flow. The observation or instrumentation chamber ("test section") was then placed at the proper location in the throat or nozzle for the desired airspeed. For limited applications, Computational fluid dynamics (CFD) can augment or possibly replace the use of wind tunnels. For example, the experimental rocket Spaceship One was designed without any use of wind tunnels. However, on one test, flight threads were attached to the surface of the wings, performing a wind tunnel type of test during an actual flight in order to refine the computational model. It should be noted that, for situations where external turbulent flow is present, CFD is not practical due to limitations in present day computing resources. For example, an area that is still much too complex
  27. 27. 13    for the use of CFD is determining the effects of flow on and around structures, bridges, terrain, etc.[6] Figure 2.3 Preparing a model in the Kirsten Wind Tunnel, a subsonic wind tunnel at the University The most effective way to simulative external turbulent flow is through the use of a boundary layer wind tunnel. There are many applications for boundary layer wind tunnel modeling. For example, understanding the impact of wind on high-rise buildings, factories, bridges, etc. can help building designers construct a structure that stands up to wind effects in the most efficient manner possible. Another significant application for boundary layer wind tunnel modeling is for understanding exhaust gas dispersion patterns for hospitals, laboratories, and other emitting sources. Other examples of boundary layer wind tunnel applications are assessments of pedestrian comfort and snow drifting. Wind tunnel modeling is accepted as a method for aiding in Green building design. For instance, the use of boundary layer wind tunnel modeling can be used as a credit for Leadership in Energy and Environmental Design (LEED) certification through the U.S. Green Building Council.[6]
  28. 28. 14    Figure 2.4 Fan blades of Langley Research Center's 16 foot transonic wind tunnel in 1990, before it was mothballed in 2004. Wind tunnel tests in a boundary layer wind tunnel allow for the natural drag of the Earth's surface to be simulated. For accuracy, it is important to simulate the mean wind speed profile and turbulence effects within the atmospheric boundary layer. Most codes and standards recognize that wind tunnel testing can produce reliable information for designers, especially when their projects are in complex terrain or on exposed sites. In the USA many wind tunnels have been decommissioned in the last 20 years, including some historic facilities. Pressure is brought to bear on remaining wind tunnels due to declining or erratic usage, high electricity costs, and in some cases the high value of the real estate upon which the facility sits. On the other hand CFD validation still requires wind-tunnel data, and this is likely to be the case for the foreseeable future. Studies have been conducted and others are under way to assess future military and commercial wind tunnel needs, but the outcome remains uncertain. More recently an increasing use of jet- powered, instrumented unmanned vehicles [“research drones”] has replaced some of the traditional uses of wind tunnels. [6]
  29. 29. 15    Chapter 3 BASICS OF AERODYNAMICS Aerodynamics is a branch of dynamics concerned with studying the motion of air, particularly when it interacts with a solid object. Aerodynamics is a subfield of fluid dynamics and gas dynamics, with much theory shared between them. Aerodynamics is often used synonymously with gas dynamics, with the difference being that gas dynamics applies to all gases. Understanding motion of air (often called a flow field) around an object enables the calculation of forces and moments acting on the object. Typical properties calculated for a flow field include velocity, pressure, density and temperature as a function of spatial position and time. Aerodynamics allows the definition and solution of equations for the conservation of mass, momentum, and energy in air. The use of aerodynamics through mathematical analysis, empirical approximations, wind tunnel experimentation, and computer simulations form the scientific basis for heavier-than-air flight and a number of other technologies. Aerodynamic problems can be classified according to the flow environment. External aerodynamics is the study of flow around solid objects of various shapes. Evaluating the lift and drag on an airplane or the shock waves that form in front of the nose of a rocket are examples of external aerodynamics. Internal aerodynamics is the study of flow through passages in solid objects. For instance, internal aerodynamics encompasses the study of the airflow through a jet engine or through an air conditioning pipe. [4] Aerodynamic problems can also be classified according to whether the flow speed is below, near or above the speed of sound. A problem is called subsonic if all the speeds in the problem are less than the speed of sound, transonic if speeds both below and above the speed of sound are present (normally when the characteristic speed is approximately the speed of sound), supersonic when the characteristic flow speed is greater than the speed of sound, and hypersonic when the flow speed is much greater than the speed of sound. Aerodynamicists disagree over the precise definition of hypersonic flow; minimum Mach numbers for hypersonic flow range from 3 to 12.[1]
  30. 30. 16    The influence of viscosity in the flow dictates a third classification. Some problems may encounter only very small viscous effects on the solution, in which case viscosity can be considered to be negligible. The approximations to these problems are called inviscid flows. Flows for which viscosity cannot be neglected are called viscous flows. Figure 3.1 Forces acting on air foil. 3.1 Introductory terminology • Lift • Drag • Reynolds’ number • Mach number 3.1.1 Lift A fluid flowing past the surface of a body exerts surface force on it. Lift is any component of this force that is perpendicular to the oncoming flow direction. It contrasts with the drag force, which is the component of the surface force parallel to the flow direction. If the fluid is air, the force is called an aerodynamic force. [3] Lift is commonly associated with the wing of a fixed-wing aircraft, although lift is also generated by propellers; kites; helicopter rotors; rudders, sails and keels on sailboats; hydrofoils; wings on auto racing cars; wind turbines and other streamlined objects. While the common meaning of the word "lift" assumes that lift opposes gravity, lift in its
  31. 31. techn flow (cruis desce Lift m racing for in An ai drag. may coeff 3.1.1 The gener refere body Lift dimen Lift c lift pr nical sense c rather than se) most of ending, or ba may also be g car. In this nstance on a irfoil is a str Non-stream also genera ficient domin .1 Lift coeff lift coeffici rated by a lif ence area as such as a fix coefficient nsional foil coefficient m ressure is the can be in any to the direc f the lift o anking in a t entirely dow s last case, th sail on a sai reamlined sh mlined objec ate lift when nated by pre ficient ient ( o fting body, t sociated wit xed-wing air is also use section, whe may be descr e ratio of lift y direction s tion of grav opposes grav turn, for exa wnwards in he term dow lboat. Figure 3.2 hape that is c cts such as b n moving re ssure drag. or ) is a the dynamic th the body. rcraft. ed to refer ereby the ref ribed as the t to reference 17 since it is de vity. When a vity. Howe ample, the lif some aeroba wn force is of Longer path capable of g blunt bodies elative to the a dimension pressure of A lifting bo to the dyn ference area ratio of lift e area. efined with r an aircraft is ver, when ft is tilted w atic manoeu ften used. Li h theory generating si and plates e fluid, but less coeffic the fluid flo dy is a foil o namic lift c is taken as th t pressure to respect to th s flying strai an aircraft ith respect to uvres, or on t ift may also b ignificantly m (not paralle will have a cient that re ow around th or a complet characteristic he foil chord dynamic pr he direction o ight and lev is climbin o the vertica the wing on be horizonta more lift tha l to the flow a higher dra elates the li he body, and te foil-bearin cs of a two d. ressure wher of el g, al. a al, an w) ag ift d a ng o- re
  32. 32. If the a met can b Wher • • • • • 3.1.2 In flu force e lift coeffici thod such as be determine re, L is lift fo ρ is air de v is true a A is plan f is the Reynolds Drag T uid dynamics s which act ient for a win s thin airfoil ed using the f orce, ensity, airspeed, form area, an e lift coeffi ’ number.[6] Table 3.1 For s, drag (som on a solid ob S ng at a speci l theory), th following eq nd icient at the rm drag and metimes calle bject in the d Shape and flo 18 ified angle o hen the lift p quation: e desired a d Skin frictio d air resistan direction of t ow Form drag 0% ~10% ~90% 100% of attack is k produced for angle of atta on for differe nce or fluid r the relative f Skin friction 100% ~90% ~10% 0% known (or es r specific flo ack, Mach ent shapes resistance) r fluid flow ve stimated usin ow condition number, an refers to elocity. ng ns nd
  33. 33. Unlik drag Drag Exam acting for ca the so depen the im 3.1.2 In fl dimen fluid coeff drag The d fluid hydro struct The d ke other resis forces depen forces alwa mples of drag g opposite to ars, aircraft a olid, as for nding on po mmobile pip .1 Drag coe uid dynami nsionless qu environmen ficient indica coefficient i drag coeffici dynamic dra ofoil also inc ture such as drag coeffici Where: is the the flow v is the m is the sp is the r stive forces nd on velocit ays decrease g include th o the directio and boat hul a sails on a ints of sail. e decreases fficient ics, the dra uantity that nt such as air ates the obje is always ass ient of any ag: skin frict cludes the ef an aircraft a ient is def drag force, w velocity, mass density peed of the o eference are such as dry ty. fluid velocit he componen on of the mo lls; or acting a downwind In the case fluid velocit ag coefficien is used to q r or water. It ect will hav sociated with object comp tion and form ffects of lift- also includes fined as: which is by of the fluid, object relativ ea.[6] 19 friction, whi ty relative to nt of the net ovement of th g in the same sail boat, or of viscous d ty relative to nt (common quantify the t is used in t ve less aerod h a particular prises the eff m drag. The -induced dra s the effects definition th ve to the fluid ich is nearly o the solid ob aerodynami he solid obje e geographic r in interme drag of fluid o the pipe.[6] nly denoted drag or res the drag equ dynamic or r surface are ffects of the e drag coeffic ag. The drag of interferen he force com d and independen bject in the f ic or hydrod ect relative t cal direction ediate directi d in a pipe, d as: cd, cx sistance of a ation, where hydrodynam ea. two basic c cient of a lif coefficient nce drag. mponent in th nt of velocity fluid's path. dynamic forc to the Earth a n of motion a ions on a sa drag force o or cw) is an object in e a lower dra mic drag. Th ontributors t fting airfoil o of a complet he direction o y, ce as as ail on a a ag he to or te of
  34. 34. 3.1.2 Exam 20 of the 0.05 r Assum enoug Solut Drag The d L L .2 Calculati mple: The dra C and 96 km e sphere is 2 respectively mptions: the gh to simula tion: force and li density of air =0.47*0.5 =402.4 N L=0.05*0.5*1 L=42.81 N Fig ion of Drag ag and lift fo mph is to be mtsqr. The d y. e flow of air ate free flow ft force are g r at 1atm and 5*1.204*26. 1.204*26.66 gure 3.3 Me and Lift for orces acting determined drag and lift is steady an over the sph given by the d 20 C is 6672 *2 672 *2 20 asured drag rce on a sphere experimenta coefficient u nd incompres here. e formulae, =1.204 kg/ coefficient at the design ally in a wind under lamin ssible; the c/ /m3 . n conditions d tunnel. Th nar condition /s of the tunn s of 1atm, he frontal are n is 0.47 and nel is large ea
  35. 35. T fr dr 3.1.3 Reyn relati densi dimen equal the le intern spher as co rules notab Wher • • • • • The value of rom Goldste rag forces ac Reynolds’ nolds’ numbe ve motion t ity and visc nsion. This d lly valid for ength or wid nal diameter rical objects ompressible apply. The bly stirred ve re, is the m is a ch diameter w is the dy is the ki is the de drag coeffic in’s graph. B cting on diff number er can be de to a surface. cosity, plus dimension is spheres or c dth can be u r is generally have an equ gases or flu e velocity m essels. mean velocity haracteristic when dealin ynamic visc inematic visc ensity of the cient is taken By following ferent object efined for a n . These defi s a velocity s a matter of circles, but o used. For fl y used today uivalent diam uids of variab may also be y of the obje linear dime ng with river osity of the cosity ( fluid (kg/m 21 n from figure g the above p s.[6] number of d initions gene y and a ch f convention one is chosen low in a pip y. Other shap meter define ble viscosity a matter of ct relative to ension, (trav systems) (m fluid (Pa·s o ) (m²/ m³)[1] e 2.3 and the procedure w different situ erally includ haracteristic n – for examp n by convent pe or a sphe pes such as ed. For fluids y such non-N f convention o the fluid (S velled length m) or N·s/m² or k /s) e lift coeffici we can calcul uations wher de the fluid length or ple a radius tion. For air ere moving rectangular s of variable Newtonian f n in some c SI units: m/s) h of the flu kg/(m·s)) ient is taken late lift and re a fluid is i properties o characterist or diameter craft or ship in a fluid th pipes or non e density suc fluids, speci ircumstance ) uid; hydraul in of ic is ps, he n- ch al es, ic
  36. 36. 3.1.4 Figu In flu the sp sound to or Wher Mach condi determ is (qu incom Mach num ure 3.4 An F uid mechanic peed of an d. It is comm above the sp re, is the is the v is the sp h number va itions, espec mine if a flo uasi) steady mpressible fl mber F/A-18 Horne cs, Mach nu object movi monly used peed of soun Mach numb velocity of th peed of soun aries by the cially tempe ow can be tre and isother low model c et at transon reaching t umber ( ing through to represent nd. ber, he source rel nd in the me compositio erature and eated as an i rmal, compr an be used. [ 22 nic speed and the speed of or ) is a air or other the speed o lative to the m edium. n of the sur pressure. T incompressib ressibility ef [7] d displaying sound a dimension r fluid divid of an object medium and rrounding m The Mach ble flow. If M ffects will b a vapor con nless number ded by the lo when it is tr d medium and number can M < 0.2–0.3 be small and ne just before r representin ocal speed o raveling clos also by loc n be used t 3 and the flo d a simplifie e ng of se al to w ed
  37. 37. 3.2 B Fig In ph vicini atmo diurn the b the su Lami circum oscill layer oncom the C heat boun 3.3 A An ai blade Boundary gure 3.5 Bou hysics and fl ity of a boun sphere, the nal heat, moi oundary lay urrounding n inar boundar mstances un lating body refers to th ming unidire Coriolis effec transfer, a t dary layer si Airfoil irfoil (in Am e (of a prope layer undary layer luid mechan nding surfac planetary b isture or mo yer is the par non-viscous ry layers ca nder which t is an examp he well-know ectional flow ct (rather tha thermal bou imultaneous merican Engl ller, rotor or visualizatio c nics, a bound ce where the boundary lay omentum tran rt of the flow flow. an be loosely they are crea ple of a Sto wn similarity w. When a an convectiv undary layer ly. [5] lish) or aero r turbine) or 23 on, showing t condition dary layer is effects of v yer is the ai nsfer to or f w close to th y classified ated. The th okes boundar y solution n fluid rotates e inertia), an r occurs. A ofoil (in Briti sail as seen transition fro s the layer o viscosity are ir layer near from the sur he wing, wh according t hin shear lay ry layer, wh near an attac s and viscou n Ekman lay surface can ish English) in cross-sect om laminar t of fluid in th significant. r the ground rface. On an here viscous  to their stru yer which de hile the Blas ched flat pla us forces are yer forms. In n have mult is the shape tion. to turbulent he immediat In the Earth d affected b aircraft win forces disto ucture and th evelops on a sius boundar ate held in a e balanced b n the theory o tiple types o e of a wing o te h's by ng ort he an ry an by of of or
  38. 38. An a comp comp a cha often work The l at a s in the can b angle attack which press the re than veloc and th Figure airfoil-shape ponent of th ponent parall aracteristic s n with asym king fluid are lift on an airf suitable angl e direction o be resolved i e of attack t k. This "turn h results in ure differen esulting flow on the lower city differen he Kutta-Jou e 3.6 Exampl d body mo his force pe lel to the dir shape with a mmetric  cam e called hydr foil is prima le, the airfoi opposite to th into two com to generate l ning" of the n lower pres nce is accom w field about r surface. Th ce without c ukowski theo les of airfoil ved through erpendicular rection of m a rounded le mber. Foils o rofoils. [1] arily the resu l deflects th he deflection mponents: L lift, but cam e air in the v ssure on on mpanied by a t the airfoil h he lift force computing t orem. [1] 24 ls in nature a h a fluid pr to the dire motion is call eading edge of similar fu ult of its angl e oncoming n. This force Lift and drag mbered airfo vicinity of t ne side and a velocity dif has a higher can be relat the pressure and within va roduces an ection of m led drag. Sub e, followed b function des le of attack a air, resultin e is known a g. Most foil ils can gene the airfoil cr higher pres fference, via average velo ted directly t e by using th arious vehic aerodynami motion is ca bsonic flight by a sharp igned with and shape. W ng in a force as aerodynam shapes requ erate lift at reates curve ssure on th a Bernoulli's ocity on the to the averag he concept cles ic force. Th lled lift. Th t airfoils hav trailing edg water as th When oriente on the airfo mic force an uire a positiv zero angle o ed streamline e other. Th s principle, s upper surfac ge top/bottom of circulatio he he ve ge, he ed oil nd ve of es his so ce m on
  39. 39. 3.3.1 In thi airfoi In flu refere relati attack focus throu In aer a fixe and th be de root o line o do no corre Some Howe angle Angle of at is diagram, il shape. The uid dynamics ence line on ve motion b k is the angl ses on the m ugh air. rodynamics, ed-wing airc he atmosphe efinable, so a of the wing on the fusela ot use an arb sponds to ze e British aut ever, this can e between th ttack the black li e angle α is t s, angle of a a body (ofte between the le between t most common , angle of att craft and the ere. Since a an alternate is chosen as age as the ref bitrary chord ero coefficie thors have u n lead to con e chord of an Figure 3. ines represen the angle of attack (AOA en the chord e body and the body's re n application tack specifie e vector repr wing can ha reference lin s the referen ference line d line but us nt of lift. [3] used the term nfusion with n aerofoil an 25 .7 Angle of a nt the flow attack. , or (Greek d line of an a the fluid thr eference line n, the angle es the angle b resenting the ave twist, a c ne is simply nce line. Ano (and also as se the zero l m angle of h the term rig nd some fixe attack of a fluid a k letter alph airfoil) and th rough which e and the on of attack of between the e relative mo chord line of y defined. Of other alterna s the longitu lift axis inste incidence in ggers' angle ed datum in t round a two a)) is the ang he vector rep h it is movi coming flow f a wing or a chord line o otion betwee f the whole w ften, the cho ative is to use udinal axis). ead - zero an nstead of an of incidence the aeroplan o-dimension gle between presenting th ing. Angle o w. This artic airfoil movin of the wing o en the aircra wing may no ord line of th e a horizont Some author ngle of attac ngle of attack e meaning th ne. al n a he of le ng of aft ot he tal rs ck k. he
  40. 40. 26    Chapter 4 CONSTRUCTION AND FABRICATION 4.1 Introduction This project aims in finding out the lift of various objects for different temperature and velocity of incoming air for various angle of attack and also effect of surface roughness on lift is studied qualitatively. The nozzle and diffuser of the wind tunnel is made of Cold Rolled steel material. The test section is made of acrylic fiber glass. These parts connected by means of nuts and bolts. A DC exhaust fan is kept in the diffuser section to suck the air. To provide proper driving pressure an AC inlet fan is kept in the nozzle section of the wind tunnel which is regulated. To provide a laminar flow a honeycomb made of plastic is kept both at the diffuser and nozzle. The temperature of incoming air is varied with the help of heating coil and provision is made for its regulation. The velocity of the air is changed by varying the speed of the inlet air. The temperature is measured using a digital thermometer and velocity of air is measured using an anemometer. The smoke lines are produced by winding cotton around nichrome wire and lighting arrangement is made for its visualization. For the measurement of lift generated by the object a strain gauge is used which is having least count of 0.01grams. Analysis is conducted on airfoil models taking smooth and rough surface. Lift is measured on parameters such as flow velocity, temperature of incoming air and angle of attack.
  41. 41. 27    4.2 Set-up Figure 4.1 Schematic diagram of wind tunnel Set-up of miniature wind tunnel is as shown in the figure 4.1. It is about 257cm long, and is constructed of Acrylic Fibre glass as test section and the two cones are constructed using CR steel. The exhaust 14cm diameter fan is placed at the outlet, making this wind tunnel work in the suction mode, pulling air through from left to right. To provide proper driving pressure we attached an inlet fan of 14cm diameter. The test object is placed in the centre of the tunnel next to the Nichrome wire. The Nichrome smoke-wire is orientated vertically. A 30*30cm piece of honeycomb is kept before the wire, and makes the flow laminar. Different honeycomb setups were investigated at the inlet, outlet, and both while also changing the distance from the test object and fan.
  42. 42. 28    4.3 Description of components 4.3.1 Nozzle Figure 4.2 Nozzle Nozzle is a part of Wind Tunnel where the air is drawn from the atmosphere to the testing chamber. The air coming through the nozzle will have high velocity and low pressure so that air reaching the test section will have high velocity. It is constructed of Cold Rolled Steel material having a thickness of 0.643mm (22 gauges) 4.3.2 Diffuser Figure 4.3 Diffuser
  43. 43. 29    The air after passing through the testing chamber exits from the Wind tunnel through the diffuser. The air passing through diffuser will have low velocity and high pressure. It is constructed of Cold Rolled Steel material having a thickness of 0.643mm (22 gauges) 4.3.3 Test Section Figure 4.4 Test Section This is the main section of the set up where the actual testing is carried out. It houses the fan at one end and honeycomb/mesh at the other end. It also houses the Nichrome wire set up and the test object. A Digital Thermometer is placed inside the test section to measure the temperature. Two 2mm diameter holes were drilled in the Acrylic Fibre glass to hold the test section without vibration. A door is provided for changing of the test object as well as easy access to cleaning of the tunnel. Stands are provided to support the tunnel. It is a transparent tunnel made of Acrylic fibre glass through which we can visualise the flow pattern easily and study it. Specifications of test section a) 30cm*30cm c/s and 60cm long section is fabricated. The size is arbitrary chosen based on the requirement and also previous work done on wind tunnel by experts suggests the size taken is sufficient to capture the data.
  44. 44. 30      b) Tunnel material (Acrylic Fibre glass) is selected based on the following requirements, Lighter material Transparent material 4.3.4 Outlet and Inlet fan Figure 4.5 Fan Housing An exhaust fan is used at the outlet of the tunnel to accommodate suction of air. It’s about 28cm diameter and is connected to a D C input (220V). Fan is having a maximum speed of 9000rpm. The fan is used to suck the air over the test object, than blowing the air. This is because suction end of the fan causes the air to be more laminar. Inlet fan is not necessary in the Wind Tunnel. In our design it is used to ensure that sufficient amount of air is available for analysis. It is 28cm in diameter and A C input having a maximum speed of 3000rpm 4.3.5 Step down transformer A transformer is a device that transfers electrical energy from one circuit to another through inductively coupled conductors—the transformer's coils.
  45. 45. 31    A step down transformer is a device that has its secondary voltage less than its primary voltage. This transformer reduces voltage and often ranges in sizes from 0.5kva to 500kva. A step down transformer is used to lower voltage input. Figure 4.6 Step down transformer Specification: 12V, 5A But when the AC is supplied, the voltage available is too high than the required voltage. This necessitates us to use step down transformer to lower the voltage from 220V, 230A to 12V, 5A. Diodes and condensers are connected to the circuit to convert AC to DC and to filter the DC. 4.3.6 Honeycomb Figure 4.7 Honeycomb structures
  46. 46. 32    Honeycomb acts as a flow "conditioner" or straightener. The honeycomb provides the area around the test object with straighter, more conditioned flow, which will lead to better which leads the flow laminar. In our project the honey comb is of a simple net like structure made of plastic material 4.3.7 Heating Coil Figure 4.8 Heating coil Heating coil is used in our project to vary the temperature of inlet air. It is made of Nichrome wire. It has a very high melting point of 1673F, highly corrosion resistant and high electrical resistivity. This coil is backed up by a ceramic material which acts as insulator. 4.3.8 Force measuring sensor Figure 4.9 Strain Gauge
  47. 47. 33    To measure the forces generated by the air on the airfoil a simple Stain Gauge is used. The force generated deflects the strain gauge which is calibrated by the microprocessor as force in the reading. Specifications It bears an electronic scale with working voltage of 2.4-3.5V.It has a least count of 0.01g and has a maximum range of 200g. 4.4 Working of the set-up Test object is kept in the testing chamber, by fixing it on a support. This support is in turn connected to the force measuring sensor (Strain Gauge).Cotton thread is wound over the Nichrome wire and grease is applied on it. The wire is connected to transformer which is connected to power supply. Wire gets heated up in few minutes and smoke is generated. Suction fan is switched on which makes the smoke to pass over the object. Honeycomb provided at the entrance makes the flow laminar. Next both inlet and exhaust fans are switched on and air is allowed to pass over the test specimen. The lift generated by the specimen is noted in the Strain Gauge. Next the lift generated is noted for various angles of attacks. The heating coil is switched on to increase the temperature inside the test section and is measure using digital thermometer. Again the lift is noted for various angles of attack. Next the velocity of the air is varied using a fan speed regulator. The effect of temperature and velocity on airfoil lift for different airfoil models for various angle of attack is tabulated and then analyzed with the help of graphs
  48. 48. 34    Chapter 5 TESTING & RESULTS 5.1 Testing and testing method Figure 5.1 Test set up The set up for testing is as shown in the above figure. The strain gauge is connected to the test specimen suitably. The arrangement for varying temperature of incoming air and velocity is made. Our project aims to measure the lift of various airfoil models for different temperature and velocity for various angles of attacks and also the effect of surface finish on lift is studied qualitatively. At first lift is measured for various objects for room temperature and for various angles of attack. Then the temperature is increased by switching the heating coil and lift is measured for various angle of attack. Next the velocity of inlet air is varied and measured with anemometer. The lift generated by the object is measured. Same steps are followed for measuring the lift of rough surface objects for different temperature and velocity of incoming air.
  49. 49. 35    5.2 Effect of temperature on airfoil lift 5.2.1 Effect of temperature on airfoil lift for temperature 32.8o C Table 5.1 Effect of temperature on airfoil lift for temperature 32.8o C TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Symmetric Airfoil Smooth 0 3.48 10 8.65 20 16.82 30 24.24 40 36.25 50 44.87 60 58.44 70 49.67 80 40.17 90 30.02 Rough 0 1.78 10 4.14 20 12.55 30 18.03 40 30.08 50 41.48 60 53.36 70 47.43 80 34.77 90 24.78   Clark Y Airfoil (Subsonic) Smooth 0 14.49 10 21.34 20 28.14 30 37.09 40 44.42 50 56.93 54 64.26 60 60.65 70 44.42 80 39.01 Rough 0 9.41 10 18.11 20 26.08 30 34.96 40 38.33 50 52.29 54 59.93 60 58.66 70 49.44 80 36.61
  50. 50. 36    TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Blackbird Airfoil Smooth 0 30.55 10 36.45 20 45.18 30 51.28 40 66.11 46 84.33 50 77.50 60 58.88 70 - 80 - Rough 0 28.47 10 32.66 20 38.49 30 50.17 40 62.34 46 82.09 50 60.78 60 53.09 70 - 80 -   Turbofan Airfoil Smooth 0 42.22 10 52.96 20 68.09 30 80.64 35 95.54 40 89.96 50 72.65 60 58.69 70 - 80 - Rough 0 30.21 10 48.25 20 64.28 30 70.18 35 90.26 40 73.87 50 54.64 60 - 70 - 80 -
  51. 51. 37    5.2.2 Graphs for airfoil models for temperature of 32.8 °C   Figure 5.2 Symmetric Airfoil for temperature 32.8ºC From the graph it is observed that for peak lift the angle of attack is 60° for smooth and rough surface and then it gradually decreases. The lift for rough surface is slightly lower than that of lower surface.     Figure 5.3 Clark ‘Y’ Airfoil for temperature 32.8ºC For this airfoil model the lift observed for smooth and rough surface is almost similar with slight variations. The peak lift is observed at 55° 0 10 20 30 40 50 60 70 0 20 40 60 80 100 Angle of attack (Degree) Lift (grams) Smooth surface Rough surface 0 10 20 30 40 50 60 70 0 20 40 60 80 100 Angle of attack (Degree) Lift (grams) Smooth Surface Rough Surface
  52. 52. 38      Figure 5.4 Blackbird Airfoil for temperature 32.8ºC For this airfoil its observed that the peak lift is at 43° and there is no much difference for the lift observed for smooth and rough surface. Also considerable lift is observed for 0° angle of attack.     Figure 5.5 Turbofan Airfoil for temperature 32.8ºC For this airfoil it is observed that the peak lift is found to be at 35º angle of attack which is far less when compared to other models. The peak lift obtained is very high and even at 0º angle of attack lift is considerably high. ‐10 0 10 20 30 40 50 60 70 80 90 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface ‐20 0 20 40 60 80 100 120 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  53. 53. 39    5.2.3 Effect of temperature on airfoil lift for temperature of 51.6o C Table 5.2 Effect of temperature on airfoil lift for temperature 51.6o C TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Symmetric Airfoil Smooth 0 1.02 10 5.22 20 9.63 30 22.48 40 35.74 50 43.98 60 53.46 70 46.78 80 38.56 90 26.08 Rough 0 1.55 10 3.94 20 10.64 30 12.94 40 22.37 50 37.11 60 43.02 70 44.29 80 30.99 90 20.18   Clark Y Airfoil (Subsonic) Smooth 0 11.25 10 16.66 20 28.11 30 35.94 40 40.78 50 49.25 54 60.78 60 53.96 70 40.19 80 34.20 Rough 0 7.25 10 15.94 20 26.00 30 33.94 40 37.18 50 48.96 54 54.12 60 50.94 70 45.44 80 30.97
  54. 54. 40    TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Blackbird Airfoil Smooth 0 27.15 10 32.54 20 44.22 30 49.67 40 57.03 46 78.00 50 72.39 60 55.09 70 - 80 - Rough 0 26.17 10 30.88 20 35.96 30 44.28 40 57.36 46 67.00 50 52.94 60 49.38 70 - 80 -   Turbofan Airfoil Smooth 0 38.25 10 50.00 20 62.32 30 69.36 35 90.47 40 85.65 50 69.36 60 60.89 70 - 80 - Rough 0 28.48 10 44.36 20 61.98 30 67.55 35 85.11 40 68.69 50 49.73 60 45.99 70 - 80 -      
  55. 55. 41    5.2.4 Graphs for airfoil models for temperature of 51.6 °C   Figure 5.6 Symmetric Airfoil for temperature 51.6ºC For this airfoil upto 20º angle of attack the lift is considerably same for smooth and rough surface but it varies largely after 20º. The peak lift obtained is less due to increase in temperature.     Figure 5.7 Clark ‘Y’ Airfoil for temperature 51.6ºC For this airfoil model lift for smooth and rough surface are similar but after the peak lift due to increase in angle of attack the lift for rough surface is more and the lift is more than symmetric model. 0 10 20 30 40 50 60 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface 0 10 20 30 40 50 60 70 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  56. 56. 42      Figure 5.8 Blackbird Airfoil for temperature 51.6ºC For this airfoil it’s observed that there is a considerable difference for the lift obtained for smooth and rough surface. The peak lift is observed to be slightly on the higher side when compared to with previous two models and is obtained between 44 to 48º AOA     Figure 5.9 Turbofan Airfoil for temperature 51.6ºC For this airfoil upto 34º angle of attack the lift is considerably same for smooth and rough surface but it varies largely after 34º. The peak lift is obtained at 36º and after the peak lift due to increase in angle of attack the lift for smooth surface is more. ‐10 0 10 20 30 40 50 60 70 80 90 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface ‐20 0 20 40 60 80 100 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  57. 57. 43    5.3 Effect of velocity of air on airfoil lift 5.3.1 Effect of velocity of air on airfoil lift for velocity 6.7m/s Table 5.3 Effect of velocity of air on airfoil lift for velocity 6.7 m/s TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Symmetric Airfoil Smooth 0 4.02 10 8.82 20 17.66 30 26.12 40 36.30 50 46.11 60 60.55 70 51.41 80 43.13 90 30.00 Rough 0 1.92 10 4.68 20 12.98 30 18.45 40 30.78 50 42.12 60 55.16 70 49.11 80 36.31 90 25.83   Clark Y Airfoil (Subsonic) Smooth 0 15.20 10 23.77 20 30.11 30 37.73 40 43.36 50 58.33 54 65.89 60 62.80 70 45.00 80 40.67 Rough 0 9.42 10 18.77 20 27.42 30 35.79 40 41.34 50 53.29 54 60.73 60 59.38 70 50.79 80 38.42
  58. 58. 44    TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Blackbird Airfoil Smooth 0 31.42 10 38.62 20 46.81 30 52.55 40 64.67 46 85.18 50 79.61 60 60.11 70 55.12 80 - Rough 0 30.14 10 33.72 20 40.67 30 52.11 40 60.26 46 83.13 50 61.45 60 55.91 70 51.42 80 -   Turbofan Airfoil Smooth 0 43.60 10 54.81 20 69.72 30 78.42 35 96.84 40 90.54 50 75.66 60 60.85 70 57.74 80 - Rough 0 31.95 10 50.48 20 64.90 30 72.45 35 91.99 40 75.33 50 57.00 60 51.69 70 38.08 80 -
  59. 59. 45    5.3.2 Graphs for airfoil models for velocity of 6.7m/s   Figure 5.10 Symmetric Airfoil for velocity 6.7m/s For this airfoil model the lift increases almost linearly with the angle of attack until the peak lift for smooth and rough surface. The lift at 0º angle of attack is nearly zero.     Figure 5.11 Clark ‘Y’ Airfoil for velocity 6.7m/s For this airfoil model lift for smooth and rough surface are similar but after the peak lift due to increase in angle of attack the lift for rough surface is more and the lift is more than symmetric model.  0 10 20 30 40 50 60 70 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface 0 10 20 30 40 50 60 70 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  60. 60. 46        Figure 5.12 Blackbird Airfoil for velocity 6.7m/s For this airfoil model peak lift for smooth and rough surface is almost equal and the considerable lift is obtained at 0º angle of attack.     Figure 5.13 Turbofan Airfoil for velocity 6.7m/s For this airfoil model initially there is great difference for the lift obtained for rough and smooth surfaces. But during the peak lift there found to be almost equal. After the lift for increasing angle of attack there is a considerable difference for the lift of two surfaces. 0 10 20 30 40 50 60 70 80 90 100 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface 0 20 40 60 80 100 120 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  61. 61. 47    5.3.3 Effect of velocity of air on airfoil lift for velocity 3.8m/s Table 5.4 Effect of velocity of air on airfoil lift for velocity 3.8m/s TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Symmetric Airfoil Smooth 0 2.01 10 4.68 20 9.52 30 14.60 40 20.45 50 23.58 60 28.36 70 35.48 80 26.89 90 16.21 Rough 0 1.02 10 2.78 20 7.50 30 10.30 40 17.80 50 21.36 60 26.54 70 32.14 80 22.87 90 12.20   Clark Y Airfoil (Subsonic) Smooth 0 8.52 10 11.26 20 16.87 30 18.44 40 23.51 50 27.22 54 35.03 60 33.25 70 26.25 80 17.98 Rough 0 5.28 10 10.69 20 13.58 30 18.57 40 21.32 50 26.04 54 36.15 60 29.26 70 22.99 80 16.78
  62. 62. 48    TYPE OF AIRFOIL SURFACE FINISH ANGLE OF ATTACK (Degree) LIFT (grams) Blackbird Airfoil Smooth 0 17.54 10 19.36 20 22.58 30 26.47 40 32.51 46 42.25 50 36.08 60 33.36 70 - 80 - Rough 0 15.96 10 17.73 20 19.22 30 25.63 40 27.85 46 45.05 50 30.26 60 31.00 70 23.50 80 - Turbofan Airfoil Smooth 0 21.20 10 26.85 20 33.56 30 36.69 35 46.97 40 34.02 50 18.22 60 19.25 70 - 80 - Rough 0 14.36 10 23.77 20 32.05 30 36.95 35 43.65 40 37.85 50 22.00 60 20.23 70 - 80 -    
  63. 63. 49    5.3.4 Graphs for airfoil models for velocity of 3.8m/s   Figure 5.14 Symmetric Airfoil for velocity 3.8m/s For this airfoil model the peak lift is found at 70˚ angle of attack and there is a considerable difference in the lift obtained for smooth and rough surface.   Figure 5.15 Clark ‘Y’ Airfoil for velocity 3.8m/s For this airfoil model the lift for smooth surface increases and decreases with increase in angle of attack when compared with rough surface. But during the peak lift the lift for rough surface found to more than smooth surface 0 5 10 15 20 25 30 35 40 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface 0 5 10 15 20 25 30 35 40 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  64. 64. 50        Figure 5.16 Blackbird Airfoil for velocity 3.8m/s For this airfoil model initially lift for smooth surface is higher than rough surface upto 46˚ and peak lift is obtained for rough surface at 46˚ angle of attack. After the peak lift the lift for smooth surface increases and decreases with increase in angle of attack when compared with rough surface.   Figure 5.17 Turbofan Airfoil for velocity 3.8m/s For this airfoil model peak lift is obtained for smooth surface at 35º angle of attack and after that peak value lift for rough surface increases. ‐10 0 10 20 30 40 50 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface ‐10 0 10 20 30 40 50 0 20 40 60 80 100 Angle of attack (Degree) Lift (gram) Smooth Surface Rough Surface
  65. 65. 51    Chapter 6 CONCLUSION AND SCOPE FOR FURTHER STUDY A wind tunnel model is fabricated with the available design. Extensive literature review is carried out and a design which fits best for our limited budget is selected. Suitable modification is done to fit our requirement. The fabrication and assembly is done with simple process usually adopted for the purpose, like welding riveting it. Plexiglas test chamber is prepared with extreme care since it needs a good handling while assembly. It was ensured that the entire basic requirement of a wind tunnel is incorporated. All the measuring instruments like strain gauge, temperature measuring instrument, wind velocity measuring device are properly put in place. Fabrication and assembly is one half of the whole project. After the complete assembly and installation our aim was to conduct series of experiments and analyze the result. Testing was mainly intended to assess the amount of lift produced on different test objects. Test objects were prepared using weightless thermo coal material. Standard shapes of airfoil sections, reviewed from literature, are reproduced. Standard shaped airfoils like Symmetric airfoil, Clark Y airfoil, Blackbird airfoil, Turbofan airfoil are used for testing. Test objects were duplicated for different working conditions like smooth surface and rough surface. Testing is done with simple procedures. Objects are placed in test section and wind is blown over it using blower provided with the setup. Airflow over the object causes force on it. The vertical components of which is called Lift force. Our aim was to quantify this lift force for different working conditions. When the object gets lifted upward force is measured using strain gauge fitted at proper place and is recorded. The procedure is repeated by keeping the object at other angle of attack. Results are tabulated and graphs are plot and suitable interpretations are done based on the results. All these tests are conducted at a particular temperature and sped of air. The testing is repeated at different temperature and speed of wind and different smoothness of the test objects. Results are tabulated and graphs are plot. Tests results clearly show the variations of lift force with variety of parameters like speed, temperature surface finish etc.
  66. 66. 52    Overall objective of our project is to quantify the lift force on different objects using the wind tunnel. Our aim is fulfilled to some extent. Accurate measurement of the force is a great challenge since little variations in the measurement may lead to the wrong conclusions. We sincerely agree that we had bit of problems in proper measurement of force and hence result of our experimentations can be treated as qualitative tool for predicting the performance of airfoils. Present work can be refined with more sensitive measuring instruments and also can be used for testing objects other than aerofoil. Smoke flow visualization can also be incorporated with it to visualize the flow pattern around the object and structure. This project can be used as an educational tool for undergraduate students to understand the basics of aerodynamics.
  67. 67. 53    REFERENCE [1] Mehrdadghods, “Theory of wings And Wind tunnel testing of a NACA 2415 airfoil”, Technical Communication for Engineers, The University of British Columbia, July 23, 2001. [2] Steven Baumgartner, “Design and development of a smoke flow visualization system”, Union College Mechanical Engineering , Winter/Spring 1999. [3] Nathan Tatham, “Wind tunnel design and operation”, University of British Columbia 1998 [4] Bradshaw P, Mehta R D, “Design rules for small low speed wind tunnels”, The Aeronautical Journal of the Royal Aeronautical Society November 1979. [5] Mehta R D, “Turbulent boundary layer perturbed by a screen” , American Institute of Aeronautics and Astronautics Journal Volume, 23, No. 9, September 1985. [6] http://www.nasa.gov/centers/langley/news/factsheets/WindTunnel.html [7] Frank M White, “Fluid Mechanics”, University of Rhode Island, International edition 2005.  
  68. 68. 54    COST DETAILS             Components Quantity Cost Tunnel 1 20500 Fan (Inlet & Outlet) 2 3000 Heating coil 1 180 Step Down Transformer 1 750 Honey Comb 2 70 Nichrome Wire 2 200 Ceramic Plugs 20 300 Miscellaneous - 3550 Airfoil Models 10 500 Transportation 1 4000 Sensors 3 1450 Painting - 2000 TOTAL 36500
  69. 69. 55    Personal Profile Prof. Chandrakanth Bekal Project Guide Prof. Chandrakantha Bekal received the BE degree in Mechanical Engineering from UVCE, Bangalore in the year 2004 and M.Tech in Product Design and Manufacturing from KVGCE, Sullia in 2007 Mr. Aby Joyce 4VP09ME004 B8-3, KSHB Colony, Mysore Road, Sulthan Bathery (PO)-673592 aby_joyce@yahoo.co.in 9480534100 Mr. Amrith A V 4VP09ME007 Vasuki nilaya, Annadka house, Perabe post, Kunthur village, Puttur (D.K)-574285 av4scorpio@gmail.com 9663042175
  70. 70. 56    Mr. Dikshith Kumar P 4VP09ME024 “Ashraya”, Near Cattle Feed Plant, Chemattamvayal, P O Balla-671531, Kasaragod, Kerala. dikshithashes@gmail.com 9743846382 Mr. Jayantha M 4VP09ME036 Madthelu Kadabari house, Mangalapadavu Post, Vittal Kasaba, Bantwal Taluk, D.K. jayantha1311@gmail.com 7760565222  

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