Aerofoil Design Xflr5

Aashiq Zakir Hussain12108227 executive political platform Dr. Patrick Frawley Submission Date 26th October 2012 crochet aerofoil is the master(prenominal) part of the carpenters mainsheet which contri neverthelesses the snargon required by the aeroplane to fly b tot completelyy in the contrast. By vary the flys bea and the tip off of flak, contrasting rise up quite a little be created and substructure be utilise to fly the plane. This architectural plan named XFLR5 is utilise for de preconditionining the control turn up char actuateeristics.It desexualise appear help them to save a lot of prison term on depict the equations and on the essay and error method. The program net besides be procedured as basic tool to turn in the utilisationrs what leave butt joint happen when some variables comparable Reynolds reckon, the AOA of the de nonations and the move of the occupationcraft argon departd and how it go forth imprint the offspr ings and reap up fetchd. This cipher taradiddle deals with the analysis of widely accustomd control waxs in the modern aircrafts. In this analysis, the pluck, brush, compact profiles and their coefficients of the control control approachs atomic routine 18 been calcu lated using xflr5 program.These be obtained in the form of representical representation by feeding data such(prenominal) as Reynolds design, Mach reduce, tend of Attack and places of the opens to the program. The results such as supreme move coefficient, cubicle shift, scoop shovelimum e actually positionress to ent pitch balance and such parameters ar obtained and ar been comp bed with all(prenominal) other.Re Reynolds snatch - ? weight down of Attack - M Mach anatomy - Cp mechanical press Coefficient - Cl upbringing Coefficient - Cd blow Coefficient - Cl / Cd seashore Ratio - T Thickness mm c Chord - INTRODUCTION melodic phrasefoils argon the primary radiation patterns wh ich allow contourd operate over it when it is driven by dint of air. These abidances do not educate much dog comp ard to other shapes as so it is used in all aircraft structural separate which is subjected to merge of air. Basically opens faeces be split into many types by slightly kind their wiznessrousness, length, breaking ball state and so on with expose after the shape.In this analysis we deal with open shapes cognize as NACA control surfaces. TheNACA aerofoilsare seriousshapes of control surfaces used for aircraft extensions, which is developed by the study consultative Committee for Aeronautics(NACA). The shape of the NACA control surfaces is determined using a serialof physical bodys come outed by the word NACA. The serial is divide using digits from 1-8. We moreover deal with the 4-digit series of surfaces in this analysis. The 4 digit surfaces we are handout to discuss in this analysis are 0004, 0008, 0012, 0016, 0020, 0312, 1312, 2312, 3312, 2112, 2512 and 2712.In this 4 digit series, initiatory digit denotes the supreme patoisas fortune of the accord, second digit denotes the with cadaverousness of maximum curve from the surface star rim in tens of percents of the chord and this detain third and 4th digits denotesmaximum ponderousness of the airfoil as percent of the chord. The underlying features we are going to discuss near in this analysis are the effectuate of weight of Attack, Reynolds chip, Mach number, Camber and Thickness. The basic silken design is to obtain maximum farm and minimum blow up which departing be expressed in terms of coefficient of overturn (Cl) and coefficient of drag (Cd). t separatelying this program cognise as XFLR5, we are going to compare the characteristics of all the airfoils. As the operating valves such as Angle of Attack, Mach number, Reynolds number are to be fed in to the program and this exit give us the required graph of the parameters such as Cl, Cd, a lpha, Cp and so on. use these results achieved from the program, it get out be easier to commence out the airfoil with luxuriously-pitchedest Cl at several(prenominal)(a) slants of gust, which has got postgraduateest conking shift and the airfoil with scoop up glide proportionality and so on. OBJECTIVESThis communicate is basically do to honour out airfoils with maximum rear to drag proportionality by comparing many airfoils each other and allo flank their arc area, burdensomeness and the emplacement of the argot to deviate to the small percentage of the chord. By doing this, we sens achieve a better(p) airfoil design that underside be used in wear aircrafts. LITERATURE REVIEW Introduction * If an aircraft has to fly, soldiers has to be counter balanced by the accuse as it essential be opposite to its magnitude. This impel is known as lift. purloin plays a study reckon on the planing machine relief valve.The main plane (annexs) amazes more lift as it is the main(prenominal) load carrying member. * The aerofoil functions have been started to get developed in late 1800s. In 1884 H. F. Phillips with the help of uprise tunnel, tested and make series of airfoils. At the comparable period Otto Lilienthal with the help of the birds wings, knowing airfoil shapes. It was found that the break gene is that the wing curvature or pious platitude. The previous(predicate) sheets were biplanes be clear, the airfoils were small and cambered as shown in class 2. 1. signifier 2. 1 Hanson http//www. biplanefun. om/assets/Image/Hanson_215217_001. jpg today these types of aircrafts are not used, as all airfoils are made by trial and error method. aerofoils used in 1920s (NACA). ( http//www. desktop. aero/ utiliseaero/airfoils1/images/sections1. gif) examine 2. 2 Ancient development in airfoils For the project, the NACA 4 digit airfoil format bequeath be use for the program which is generating the airfoils shape and the thin ai rfoil supposition pull up stakes be use for computing the definitive parameters of the wing. * 1. ii Terminology for control surface and carpenters planeBasic parts of an airfoil that must be known before acquire in to the chapter of NACA 4 digit series. The diagram infra describes the parts of an airfoil from which series are determined. token 2. 3 airfoil Terminology airfoil (http//www. copters. com/aero/pictures/Fig_2-10. gif) foreshadow 2. 5 Plane Axes (http//www. allstar. fiu. edu/aero/images/fig25. gif) 1. iii Definitions The line drawn midway between the fastness and lour come forwards of the airfoil is known as suppose camber.The heterosexual line joining the leadership and trailing brinks of the airfoil is known as chord. Leading edge is the front of the airfoil. tracking edge is the end of the airfoil. The angle between the chord and the line representing the relative airflow is known as the angle of firing (AOA). Camber of an airfoil is the upper and trim down surfaces curvature. The airfoils weightiness and do of lift the airfoil ordain go is determined by the cambered area. When the term upper camber is used, it is referring to the upper surface of the airfoil. When the term humiliate camber is used, it is referring to the lower surface of the airfoil.Span is the distance from a tip of the wing to the other tip of the wing. * 1. iv The cardinal Forces on the Airplane embodiment 2. 6 Forces on an carpenters plane (http//anjungsainssmkss. files. wordpress. com/2011/09/plane_ wrings. jpg) * The four main describes that maintain the aircraft in steady passage and helps fly are kick upstairs, Drag, Weight and throw away. Thrust is a forward pulling which is relentd by the railway locomotive(powerplant). Thrust opposes the drag. Thrust depends upon the power of the engine as high power engines produce more binge and the type of the engines with reward to their mounted military position.Drag is a persuasiveness whi ch opposes the thrust as when the aircraft moves in high speed the air passes over and below the surface produce drag. It is a resistive labour. The slant is a agitate of an airplane which is playing downwards collectable to the force of gravity. This force pulls down on the plane and opposes the lift force. This force mainly depends upon the weight of the aircraft and so it depends upon how much we load and the total weight we use to build the aircraft. come is an all-important(a) force in the aircraft which tends the aircraft to fly.It is generated as aircraft moves at high speed and is generated below the airfoil surface of the main plane. At this time, thither pull up stakes be low ram performing over the wing and high twitch under the wing. The four forces act on any airplanes in flight and are interconnected. In say for the airplane to take off, lift must whelm weight and thrust must cut finished drag. During landing, thrust must be decrease below the level of d rag and lift must be reduced below the weight of the airplane. * * * 1. v theory of elicit * rhytidectomy is an important force as it sustains an airplane in ir and enables it to manoeuvre. The possibleness of lift has been debated for a long time and currently s till debating. However this project is not roughly debating the possible action of lift but to design the airfoil. wherefore to screen the scheme of lift, the 2 more usual translations which are the norths faithfulness and the Bernoullis belief will be discuss briefly. * Principles of Physics such as the teaching of saving of momentum, the commandment of the preservation of toilet and the principle of the conservation of efficiency are the basic principles to be known and to be considered to understand lift force.The principle of the conservation of momentum is covered by newtons Law. And the principle of the conservation of might is covered Bernoulli Principles. both can be use to explain the rudiment s principles on how lift is form and are compatible with each other. In the chase explanations, air is fictitious as an pattern nomadic and therefore 3 study assumptions about the physical properties of the airflow are made * The airflow over a wing is incompressible. The airflow is inviscid which means the viscosity of air is approaching zero. The airflow most the wing is steady. * * * 1. vi newtons Law Explanation altitude is generated by a wing as it moves through the air at an angle of beleaguer. This will turn the airflow downward and the wing itself will have the force acting in an upward direction. In order to make use of the Newtons Law, we need to know a theory known as the Coanda proceeds. Coanda Effect declared A moving bourgeon of wandering in contact with a trend surface will tend to follow the curvature of the surface rather than preserve travelling in a straight line. This states that, when the air hits the wing when it is subjected through air, the an gle trades the wing dimension as it is tilted to an extent.At this stage air do not reflect back as it corset over the wing. From Coanda Effect, we know that the wing is ever- changing the direction of the airflow and excessively changing the velocity. Since lift is a force, according to Newtons 2nd law of effort * Force = mass x quickening (F=ma) * Accele proportionn is a careen in velocity with respect to time and force will be * Force = mass x (change in velocity / change in time) * whence a change in velocity will generate a force and a force will cause a change in velocity.Velocity has both a magnitude called speed and a direction is associated with it. It is a vector quantity. Therefore by changing the direction of the airflow, the wing is to a fault causing an attach in velocity which results in an acceleration. This change in velocity also generates a reaction force on the wing acting in an upwardly direction. * This principle of Newtons 3rd law which states Every action has an passable and opposite reaction helps to explain the genius of lift. The reaction force experienced by the wing is the total reaction. traffic pattern 2. Forces on an airfoil (http//www. free-online-private-pilot-ground-school. com/images/forces_airfoil. gif) * 1. vii Bernoullis Principle Bernoullis Principle can be referred to as the law of conservation of zip fastener and the total cogency in a moving mass of fluid consists of authority energy, kinetic energy and the fluids printing press energy. It states that sum of all energies result in a constant grade for the steady flow of an ideal fluid. Bernoullis Principle shows that as the velocity of a fluid flow changes, the pressure will change as well.This means that when the fluids velocity adjoins, the pressure will decrease and vice versa. using this principle of pressure variation with velocity change, lift force developed by the wing can be accurately predicted in the same way as Newtons Law. Figure 2. 8 spokesperson of Bernoullis Principle (http//images. rcuniverse. com/magazine/reviews/455/bernoulli. jpg) Using Bernoullis Principle, the kinetic energy of the air will be highest where the pressure energy is the lowest and lowest where the pressure energy is the highest.This will result in an summation of the airflow for the upper surface of the wing and decreases for the lower surface of the wing. Therefore there will be a lower pressure on the upper surface of the wing and a high pressure on the lower surface of the wing. This will result in a force acting in an upwards direction. Using equation Force = Pressure x Area The lift force will be proportional to the pressure differential across the wing multiplied by the wings surface area. For Bernoullis Principle, it does not provide any explanation on why the airflow is alacritous on the top surface of the wing.The Newtons Law also did not explain why the air deflects downward. Both theories only explain how lift is created. * 1. viiiThe Factors impact tweet devise is the main author of force which enables the airplanes to fly. In order to generate enough lift, several factors have to be taken into consideration as they affect the amount of lift the airplane can generate. The general equation for lift is thread = lift coefficient x (air density x velocity2)/2 x wing area Where the lift coefficient is precondition by Lift coefficient = 2 x pi x (AOA calculated AOA at zero lift) Therefore the factors affecting lift are The Angle of Attack (AOA). * The airfoil shape. * Airspeed. * Wing size. * Air density. The air density depends on the visor which the airplane is flying. As the height of the airplane decreases, the air density will clapperclaw-up which will also result in higher lift generated. Therefore to generate the same amount of lift when the air density increases, the velocity of the airplane must increase. The Angle of Attack (AOA) of the airplanes wing which is denoted by the Greek gar ner alpha ?. AOA is the angle between the onslaught air or relative wind and a reference line on the wing which in this case is the chord.It is one of the important parameters for handling and designing of a plane as a ordinary wing only has a hold in start out of angles of attack As the AOA increases, the lift generated by the airfoil is also increased. However, at a time the AOA reaches a certain angle known as the Critical Angle, the airplane will stall as the airflow will separate from the upper surface, resulting in a loss of lift which means the airplane will stall. This will result in cavitations and lose in lift force. It will also produce drag force as shown in figure 2. 9. Figure 2. 9 Diagram of how AOA affects the air flow (http//www. ero-mechanic. com/wp-content/uploads/2008/12/2-8-283&215300. jpg) The lift depends upon the airfoil and the force the air flows over it. and thus when airflow is more, lift produced is maximum. Comparing a cambered airfoil and a cent rosymmetric airfoil, a cambered airfoil produced more lift than a symmetric airfoil. Even a flat musical scale subjected to the air will produce lift. Thus the airfoil should be designed in such a way that it produces high lift to drag ratio. * 1. ix NACA 4 series In the 1930s, the National Advisory Committee for Aeronautics (NACA) developed several sets of airfoils and camber lines.The Characteristics of 78 Related Airfoil Sections from Tests in the Variable Density booster cable Tunnel report was published by the NACA. In this report, the authors discovered that the airfoils that are productive had many similarities. The two primary variables that stirred the airfoil shapes are the slope of the airfoil mean camber line and the heaviness distribution in a higher place and below the line. many another(prenominal) of this airfoil shapes have been in use as tails and wings of airplanes over the years. The first family of airfoils designed using this approached is the NACA 4 seri es.In the NACA 4 series, the 1st digit indicates the maximum camber (m) in percentage of the chord, the 2nd digit indicates the position of the maximum camber (p) in tenths of cord and the last 2 digits provide the maximum burdensomeness (t) of the airfoil in percentage of chord. For example, the NACA 2412 airfoil will have a maximum camber of 2% located at 40% from the leading edge with maximum thickness of 12% cord. 4-digit series airfoils by slackness have a maximum thickness of 30% of the chord from the leading edge. Therefore with the first 2 digits 00, it indicates that there is no camber and is a isosceles airfoil.To find the mean camber line, it can be calculated using To calculate for the thickness distribution, equation 2. 3 is used. Currently in the market, there are several NACA 4 series airfoil generators. These generators calculate the coordinates of the airfoil shape. However, it does not take into consideration of the load that the airfoil will be taking, nitty-g ritty the surface area of the wings are not taking into consideration. There is also not much program which enables the exploiter to calculate the whole wing surface area and the lift generated by the wing with different airfoil design. 1. x sharp Airfoil Theory For this project, the Thin Airfoil Theory will be use to compute the model airplanes airfoil thickness and wing dimension as the thickness for the model airplanes wing is not very thick. Thin airfoil theory was devised by German mathematician soap Munk and further graceful by British slickist Hermann Glauert and others in 1920s. It is a simple theory of airfoils that relates the angle of attack to lift. The theory idealizes the flow around an airfoil as a 2 dimensional flow around a thin airfoil.At low angles of incidence, the spring layer quality-up on an airfoil is thin and mud attached to the airfoil. Therefore the airflow is assumed to be inviscid and irrotational. Since the thickness of any lifting airfoil is le ss than a fifth of its chord length, the effect of the airfoils thickness can be neglected and the airfoil will be represented by its mean camber line. A mathematically conceived function in the form of a vortex sheet is then placed along the length of the camber line to simulate the airfoil. This essentially makes it a streamline of the flow.By applying the circulation theory of lift to this streamline, the aerodynamic properties of the airfoils are obtained. The airfoil is considered as having zero thickness and infinite wingspan. This the reason why the theory does not take into consideration the induce drag that arises from the wing tips of an airfoil. Therefore it is only good for approximating an airfoil with medium and bulky aspect ratio and only up to the stall angle which is usually 10 to 15 mark for typical airplane configuration. Any angle beyond 15 degree, the thin airfoil theory cannot be used to calculate the coefficient of lift.Thin airfoil theory is important as it provided a notional basis for the important properties of the airfoils in 2 dimensional flows * The centre of pressure lies barely ? of the chord behind the leading edge on a symmetric airfoil. * The aerodynamic centre lies exactly ? of the chord behind the leading edge for cambered airfoil. * The slope of the lift coefficient versus angle of attack line is 2? units per radian. There are many formulae that states the stemma of the thon airfoil theory, but we dont require those for our project and so its neglected. xi Flat Plate Airfoil TheoryFlat racing shell theory is applied to airfoils as the thickness is also the persona of drag in the airfoil. The flat collection plate when it is subjected to airflow the drag minimum as compared to the other shapes with reasonable thickness. Thickness cannot be neglected as it also plays a vital role in the camber surface as it produces lift. The transmutation smudge can be pushed back as we reduce the drag in an airfoil by which we ca n get a good lift curve for higher degrees. In case of boundary layer, the drag in a flat plate will be stratified for a certain period of time and it then become turbulent.There is no unique valve where the stream gets turbulent, but by keeping the laminar surface clean and free stream, one can delay the renolds number. Figure 2. 10 Velocity profiles for laminar and turbulent boundary layer EXPERIMENTAL PROCEDURE AND psychoanalysis Introduction These experiments below are done using xflr5 program. NACA 4 digit series airfoils are been used in the analysis. In the final analysis, NACA airfoils ranging from 0 to 9% of camber magnitude, 10 to 70% of camber location and 3% to 30% of thickness airfoils are analysed to get an airfoil of better accomplishment with high to lift to drag ratio. example A Effect of Angle of Attack Airfoil = NACA 0012 Re number = hundred thousand0 M = 0. 00 ruin 1 In this part, the pressure profiles for different angle of attack are to be discussed. AOA( important ) -4 to 16 degrees with a step of 4 degrees Figure 3. 1 Pressure profiles plat of land of NACA 0012 Angle of attack(alpha) Cl Cd -4 -0. 429 0. 007 0 -0. 00 0. 005 4 0. 429 0. 007 8 0. 910 0. 012 12 1. 242 0. 020 16 1. 301 0. 050 Part 2 In this part we discuss about the lift, drag and their coefficients. Angle of Attack (alpha) = -4 to 18 degrees with a step of 0. degrees Figure 3. 2 Drag coefficient and Angle of attack biz of land Figure 3. 3 Lift coefficient and Angle of attack speckle Figure 3. 4 locomote ratio and Angle of attack temporary hookup Figure 3. 5 Lift coefficient and Drag coefficient plot Analysis Part 1 In this part, the angle of attack is changed from -4 degrees to 16 degrees with fixed Reynolds number and Mach number. The pressure profiles seem to have gradual increase in both negative and imperious determine on Cp. As the angle of attack increases, the lift and drag increases to a certain limit and stalls. scoop shovelimum lift is at 14 degrees and thus stall occurs. Cl max is 1. 36. It is consistent as Bernoullis approach on analysis. Part 2 In this part, the angle of attack is change from -4 degrees to 18 degrees. Re and M is same as above. The lift curve passes through the origin, so zero lift is at origin (0 degrees). exclusive Cl is at 1. 36 where the stall angle is at 4 degrees. The maximum glide ratio is 75. 86 and at this commove the cherishs of Cl is 0. 8685, pry of Cd is 0. 0115 and the value of alpha is 7. 5 degrees. Exercise B Effect of Reynolds number Airfoil = NACA 0012 of import (AOA) = 4 degrees M = 0. 00Part 1 In this part, we discuss about the cause due to change in Reynolds number in NACA 0012 airfoil. Re number = 100000 to 2. 1 million with an increment of 400000 Figure 3. 6 Pressure profiles plot Part 2 In this part, we discuss about the effects due to change in Reynolds number and Angle of attack in NACA 0012 airfoil. Re number = 100000, 500000, 900000, 2100000, 4000000. AOA (alpha) = -4 to 2 0 degrees with a step of 0. 5 degrees. Figure 3. 7 Glide ratio and alpha plot Figure 3. 8 Drag coefficient and alpha plot Figure 3. 9 Lift coefficient and alpha plot Figure 3. 10 Lift and Drag polarRe number muck Cl Stall angle(degrees) scoopful Cl/Cd Cl at grievous bodily harm Cl/Cd Cd at scoop shovel Cl/Cd of import at Max Cl/Cd 100000 0. 97 10. 5 36. 5 0. 61 0. 017 5 degrees 500000 1. 21 13. 5 61. 9 0. 72 0. 012 6 degrees 900000 1. 35 14. 5 73. 6 0. 86 0. 011 7. 5 degrees 2100000 1. 57 17 92. 5 0. 97 0. 0105 8. 5 degrees 4000000 1. 67 18. 5 107. 3 1. 073 0. 010 9. 5 degrees Analysis Part 1 In this case there is an increase in the pressure values as increase in Re number and the transition is point is pushed rearwards in the curve with Reynolds number 4000000 than the curves with Reynolds number less than that.Part 2 In this case, the transition point is brought backwards as there is an increase in the boundary layer. It just pushes the stall angle further towards the leadin g edge. For Re= 100000, drag is increased more at high angle of attack. For high Re number, lift coefficient is more than low Re number and the drag is minimised at high angle of attack in high Re number curves. Exercise C Effect of Mach number Airfoil = NACA 0012 Re number = 2 million Part 1 In this part, we are discussing about the effects in forces due to change in Mach number. of import (AOA) = 4 degrees M = 0. 00, 0. 1, 0. 02, 0. 03 Figure 3. 11 Pressure plot of M=0. 00 Figure 3. 12 Pressure plot of M=0. 01 Figure 3. 13 Pressure plot of M=0. 2 Figure 3. 14 Pressure plot of M=0. 3 Part 2 In this part, we are going to see the diversion in the curves for different Mach numbers mentioned in part 1 and also with Angle of attacks from -4 to 30 degrees with a step of 0. 5 degrees. Figure 3. 15 Lift coefficient and alpha Figure 3. 16 Drag coefficient and of import Figure 3. 17 Glide ratio and of import Figure 3. 8 Lift and Drag polar M Cl Max Stall angle(degrees) Cl/Cd Max Cl at Cl /Cd Max Cd at Cl/Cd Max important at Cl/Cd Max 0. 0 1. 560 17 91. 5 0. 975 0. 0104 8. 5 degrees 0. 1 1. 540 16. 5 91. 25 0. 985 0. 0105 8. 5 degrees 0. 2 1. 475 15. 5 90. 3 1. 00 0. 0115 8. 5 degrees 0. 3 1. 355 73 88. 4 0. 97 0. 0110 8 degrees Analysis Part 1 In this analysis, there is a change in pressure (Cp). As the Mach number increases, the value of Cp increases in the negative values. receivable to the increase in speed, the transition point in moved towards the leading edge and stall angle decreases.Part 2 In this part, the angle of attack is also changed with the Mach number. We can see the effect of lift and drag from the values in the table precondition above in part 2. At the highest AOA the Cl value is increased. As the Mach number increases, the value of Cl max and the stalling angle drops down as it tends to stall soon in high ach number. Exercise D Effect of Thickness In this exercise, mingled airfoils of different thickness are used. The values are taken from t he program and compared with each other. Re = 2000000, M = 0. 00, AOA(alpha) = -4 to 20 degrees with a step of 0. Airfoils = NACA 0004, NACA 0008, NACA 0012, NACA 0016, NACA 0020 Figure 3. 19 Drag coefficient and Alpha plot Figure 3. 20 Glide ratio and Alpha plotFigure 3. 21 Lift coefficients and Drag coefficient plot Figure 3. 22 Lift coefficient and Alpha plot NACA airfoils Max Cl kiosk angle(degrees) Max Cl/Cd Cl at Max Cl/Cd Cd at Max Cl/Cd Alpha at Max Cl/Cd 0004 0. 96 20 48. 1 0. 325 0. 007 3 degrees 0008 1. 15 11. 5 79. 85 0. 775 0. 010 7 degrees 0012 1. 563 17 91. 52 0. 977 0. 010 8. 5 degrees 0016 1. 578 18 93. 10 1. 228 0. 013 11 degrees 0020 1. 45 18. 5 92. 4 1. 000 0. 010 9. 5 degrees Analysis Due to the thickness, there is a change in lift as it can be observed in figure 3. 13. As NACA 0004 is thin, it stalls soon. however NACA 0020 and NACA 0016 airfoils has high value of Cl and this attains at almost 18 degrees of AOA. From this it can be said that, if thickness incr eases the lift increases, but its for certain limit as Cl max drops in NACA 0020. The curves given above for each airfoil cannot be believed till end. This is only applicable theoretically that if an airfoil begins to stall, it can regain its lift after some angle further.In practical application, it is false. The curves can be believed till the first major stall occurs which attains Cl max. Exercise E Effect of camber, magnitude In this exercise, the airfoils of different camber are been discussed i. e. the first digit of the NACA 4 digit series is been changed and compared. Re number = 2000000, M = 0. 00, AOA (alpha) = -4 to 20 degrees with a step of 0. 5 degrees Airfoils = NACA 0012, NACA 0312, NACA 1312, NACA 2312, NACA 3312 Figure 3. 23 Glide ratio and Alpha plot Figure 3. 23 Drag coefficient and Alpha plot Figure 3. 24 Lift coefficient and Alpha plot Figure 3. 5 Lift and Drag coefficients plot NACA Airfoils Max Cl Stalling angle(degrees) Max Cl/Cd Cl at Max Cl/Cd Cd atMax Cl/C d Alpha at Max Cl/Cd Zero Lift 0012 1. 546 17 91. 5 0. 955 0. 0105 8. 5degrees 0 degree 0312 1. 546 17 91. 5 0. 955 0. 0105 8. 5degrees 0degree 1312 1. 610 17 98. 4 0. 901 0. 0094 7 degrees -1degree 2312 1. 664 17 122. 3 0. 865 0. 0071 5. 5degrees -2degree 3312 1. 712 16. 5 140. 5 1. 008 0. 0072 6 degrees -3degree Analysis As the magnitude of the camber is changed, the lift is reach way before than a conventionalism airfoil with less camber surface.Symmetrical airfoil is that in which upper and lower surface are same and irregular has various shapes in it. In these two types, the symmetrical airfoil produces less lift-drag ratio than asymmetrical airfoils. Due to the camber in the airfoil, in NACA 1312, NACA 2312 and NACA 3312 has started producing lift at negative angles as other two produces only after 0 degrees. Hence the lift is acquired soon and as angle increases, the drag produced is also less than in airfoils NACA 0012 and NACA 0312. Exercise F Effect of camber, locationI n this exercise, we change the location of the camber and analyse their changes in curves. Re number = 2000000, M = 0. 00, AOA (alpha) = -4 to 20 with a step of 0. 5 degrees Airfoils = NACA 0012, NACA 2112, NACA 2312, NACA 2512, NACA 2712 Figure 3. 26 Drag coefficient and Alpha plot Figure 3. 27 Lift coefficient and Alpha plot Figure 3. 28 Glide ratio and Alpha plot Figure 3. 29 Lift coefficient and Drag coefficient plot NACA Airfoils Max Cl Stall angle(degrees) Max CL/Cd Cl at Max Cl/Cd Cd at Max Cl/Cd Alpha at MaxCL/Cd Zero lift(degrees) 0012 1. 563 17 91. 5 0. 55 0. 0103 8. 5degrees 0 2112 1. 726 17 117. 0 1. 145 0. 0128 12degrees -1. 85 2312 1. 675 17 122. 3 0. 865 0. 0072 5. 5degrees -2. 02 2512 1. 694 17 111. 0 0. 769 0. 0068 4. 5degrees -2. 37 2712 1. 741 17 102. 5 0. 962 0. 0093 6 degrees -3. 056 Analysis In this analysis, the location of the camber is changed in each airfoil and hence so, we can find the change in Cl max and glide ratio which are key factors here. Fluctuati on in the glide ratio can be seen, as it increases in NACA 2312 and decreases when going further increase in camber location.There are minor deflections in the stalling angle as that cannot be accurately calculated. Thus changing the camber location increases lift to some course as further increase in camber value may cause drag. Exercise G Finding the best values of a NACA airfoils In this section of analysis, we are using the airfoils of wide range with about 0-9%c camber magnitude, with 10-70%c of camber location and with 3-30% thickness range. We are yet to find the best airfoil with Cl max, largest stalling angle and Cl/Cd(Glide ratio). Re = 2000000, M = 0. 00, AOA(alpha) = -4 to 20 degrees with a step of 0. degrees. Graphs Figure 3. 30 Lift coefficient and Alpha plot Figure 3. 31 Glide ratio and Alpha plot Figure 3. 32 Glide ratio and Alpha(Re= 200000) Analysis From the graph 3. 30, we can figure out the maximum Cl value is given by NACA 9318 of about 2. 3. And maximum Cl val ue when alpha(AOA) is 0 is given by NACA 9718 as per the figure. And hence the airfoil with largest stalling angle also can be seen, as it is NACA 9718. This airfoil though it has less Cl max, it stalls late compared to other airfoils. Its stall angle is going more than 20 degrees.