# تصنيع جناح لسرعة أقل من سرعة الصوت ودارسة توزيع الضغط عليه



## سامح الفيومى (29 يونيو 2014)

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 Manufacturing of a subsonic wing and study the pressure distribution around it​ ​  Contents:-​ 1) INTRODUCTION​ 2) CH. (1)………………. Airfoil and gumetry ​ selection​ 3) ch. (2)………………. wing structure​ 4) ch. (3)………………. pressure distribution ​ 5) theoretical and experimental side ​ 6) side of some information
​ Introduction​ The wing may be considered as the most important component of an aircraft, since a fixed-wing aircraft is not able to fly without it. Since the wing geometry and its features are influencing all other aircraft components, we begin the detail design process by wing design. The primary function of the wing is to generate sufficient lift force or simply lift (L). However, the wing has two other productions, namely drag force or drag (D) and nose-down pitching moment (M). While a wing designer is looking to maximize the lift, the other two (drag and pitching moment) must be minimized. In fact, a wing is considered as a lifting surface that lift is produced due to the pressure difference between lower and upper surfaces. Aerodynamics textbooks are a good source to consult for information about mathematical techniques for calculating the pressure distribution over the wing and for determining the flow variables.​ The particular design of a wing depends on many factors such as the size, weight, speed, rate of climb, and use of the aircraft. The wing must be constructed so that it holds its aerodynamics shape under the extreme stresses of combat maneuvers or wing loading.​ The steps you should follow for designing are :​ 1. Understand wing’s application.​ 2.  Select proper wing design.​ 3.  Define parameters.​ Type of wing:​ 
 Low Wing​ These types of aircraft have their wings connected to the lower bottom part of the fuselage, thus the definition for low wing aircraft. Flying enthusiasts have different likes for the type of aircraft they enjoy. One of Tom’s favorite low wing aircraft is the F4U Corsair. When you look yourself at the characteristics here are some to consider:

Low wing advantages:
-Looks cooler (to some).
-Takes better advantage of ground effect.
-Better top visibility (very handy for entering the runway at an uncontrolled airport).
-Better crash performance. 
-Easier to make cantilever (strut less) without sacrificing head room.​
 Figure (1) low wing​ High Wing​ These types of aircraft have their wings connected to the top part of the fuselage, thus the definition of high wing aircraft. Flying enthusiasts have different likes for the type of aircraft they enjoy. In addition to Tom training in a Cessna 172, one of his other favorite high wing aircraft is the Cessna 210 Centurion. When you look yourself at the characteristics here are some to consider. ​ High wing advantages:
-Better lift (due to top of the fuselage effectively adding to wing area).
-Better downward visibility.
-More shade.
-Shelter from rain when entering/exiting.
-Easier entry/exit.
-Larger doors possible (because sealing/securing is not as difficult in the high pressure area below the wing).
-Larger flaps can be installed.
-Greater wing clearance allows for clearing obstacles (on the taxiway/runway or in the hangar).
-Fuel doesn't require boost pumps to move to engine(s).
Easier to mount struts (for lighter weight but more parasite drag).
Doesn't require as much dihedral (angled up) for lateral (roll) stability.​ Chapter One​ Airfoil and Geometry Selection​ Airfoil and geometry selection:-​ The airfoil is the heart of the airplane because it is the shape of a wing or blade (of a propeller, rotor or turbine), That is effects the cruise speed, takeoff and landing distances and overall aerodynamic efficiency during all phases of flight.​ *Airfoil geometry:-*

figure(1-1)​ The figure gives technical definitions of wing’s geometry which is one of the chief factors affecting airplane lift and drag. The terminology is used throughout the airplane industry.​ *The chord length: -* isa straight line connecting the leading edge and trailing edge of the airfoil. So all airfoils dimensions are measured in terms of the chord.​ *Mean chamber line*:-is a line drawn halfway between the upper surface and the lower surface.​ *Maximum chamber*:- is the maximum distance between the mean chamber line and the chord line .​ *The maximum thickness*: - is the maximum distance between the upper and lower surfaces. The location of the maximum thickness isimportant. The airfoil thickness ratio (t/c) refers to the maximum thickness of the airfoil divided by its chord.​ *The leading edge*: - is the point at the front of the airfoil that has maximum curvature.​ *The leading edge radius*: - is the measure of sharpness of the leading edge it varied from zero for knife edge supersonic airfoil to about 2% of chord for the blunt leading edge airfoil.​ *The trailing edge*:- is defined as the point of maximum curvature at the rear of the airfoil​ figure (1-2)​ *Aspect Ratio:-* of awing is essentially the ratio of its breadth(chord). A high aspect ratio indicates long, narrow wings, whereas a low aspect ratio aspect ratio indicates short, stubby wings.​ *The pitching moment*:- it is the that aerodynamic moment (or torque) produced by the aerodynamic force on the airfoil if force is considered to be applied, not at the center of pressure, but at the aerodynamic center of the airfoil. The pitching moment on the wing of an airplane is part of the total moment that must be balanced using the lift on the horizontal stabilizer.​ *Aerodynamic center:-* it is that point about which the aerodynamically generated moment is independent of angle of attack​ *The center of pressure* :- it is the location where the resultant of a distributed load effectively acts on the body or the point on the body about which the pitching moment is zero​ *Airfoil families*:-​ figure(1-3)​ *The NACA airfoil series:-*

A variety of airfoils is shown in figure(1-3),​ The early NACA airfoil series, the 4- digit, 5-digit and six-digit series, where generated using analytical equations that describe the camber (curvature) of the mean –line (geometric centerline) of the airfoil section as well as the section’s thickness distribution along the length of the airfoil .​ *NACA four-digit series:- *​ The first family of airfoils designed using the approach become known as the NACA four-digit series. The first digit specifies the maximum camber (m) in percentage of the chord (airfoil length), the second indicates the position of the maximum camber (p) in tenths of chord, and the last two numbers provide the maximum thickness (t) of the airfoil in percentage of chord.​ For example, the NACA2415 airfoil:-​ has a maximum thickness of 15% with a camber of 2%located 40% back from the airfoil leading edge (or 0.4c).​ *NACA five- digit series*:-​ *Consider NACA 23012 airfoil*​ · 2 : multiplied by 3/2 gives the design lift coefficient = 0.3​ · 30 : divided by 2 give the location of maximum camber along the chord from the leading edge =0.15 c​ · 12 : the maximum thickness =0.12 c​ *NACA six- digit series*:-the 6-series was derived using an improved theoretical method that, like the 1-series, relied on specifying the desired pressure distribution and employed advanced mathematics to derive the required geometrical shape. For example NACA 65-218:-​ 6…identify the series​ 5*(c/10) location of minimum pressure (0.5c)​ 2/10 design lift coefficient (0.2)​ 18%c maximum thickness (0.18c)​ *Airfoil lift and drag*​ An airfoil-shaped body moved through a fluid produces an aerodynamic force. The component of this force perpendicular to the direction of motion is called lift. The component parallel to the direction of motion is called drag. Subsonic flight airfoil have a characteristic shape with a rounded leading edge, followed by a sharp trailing edge, often with a symmetric camber.​ The lift on an airfoil is primarily the result of its angle of attack and shape. When oriented at a suitable angle, the airfoil deflects the oncoming air, resulting in a force on the airfoil in the direction opposite to the deflection. This force is known as aerodynamic force and can be resolved into two components: lift and drag. Most foil shapes requirea positive angle of attack to generate lift, but cambered airfoils can generate lift at zero angle of attack. This “turning” of the air in the vicinity of the airfoil creates curved streamlines which results in lower pressure on one side and higher pressure on the other. This pressure difference is accompanied by a velocity difference, via Bernoulli’s principle, so the resulting flow field about the airfoil has a higher average velocity on the upper surface than on the lower surface. The lift force can be related directly to the average top/ bottom velocity difference without computing the pressure by using the concept of Circulation and the Kutta -Joukowskitheorm​ Streamlines around a NACA 0012 airfoil​ *The lift equation :-*

Lift =​ *​ Where :-​ CL= lift coefficient.​ L= lift .​ = density of air .​ A=wing area.​ V= velocity of air.​ Lift coefficient is a number that aerodynamicists use to model all of a complex dependencies of shape, inclination and some flow condition on lift.​ The drag equation:-​ Drag =​ For an aircraft:​ The drag coefficient is number that aerodynamicists use to model all the complex dependencies of shape, inclination and flow condition on aircraft drag.​ When q​ The drag coefficient which includes the effects of skin friction and shape [form] and an additional drag coefficient related aircraft lift. Because of pressure difference above and below the wing the air on the bottom of the wing is drawn on to near the wing tips. This creates a swirling flow which changes the effective angle of attack along the wing and induces a drag on the wing the induced drag coefficient Cdi is equal to square of the lift coefficient CL divided by the quantity π times the aspect ratio AR​ Where:-​ AR=​ The total drag is:-​ Cd= cdo+ cdin​ *Mach number:-*

Is a dimensionless quantity representing the ratio of speed of an object moving through a fluid and the local speed of sound​ M=​ Where:-​ M= the Mach number .​ V= the velocity of the source relative to the medium.​ V sound= the speed of sound in the medium​ There are three important regions of Mach number:-​ 1. M<1: subsonic flow.​ 2. M=1: sonic flow.​ 3. M>1: supersonic flow.​ These regions can be split up into more detailed, but less important regions, as can be seen in table (1).​ 
 *Mach number:*​ *M< 0.3*​ *0.3<M<1*​ *M=1*​ *M*​ *1<M<5*​ *5<M*​ *Name:*​ Low subsonic​ High subsonic​ Sonic​ transonic​ supersonic​ hypersonic​
 Table (1)​ Chapter Two​ Wing Structure​ ​ *Wing structure :-*

Providing lift is the main function of the wing of an aircraft . The wings consist of two essential parts* . The internal wing structure *, consisting of spars,ribs and stringers,* and the external wing ,*which is the skin . The internal wing structure is made up of spars , stringers and ribs.​ *Ribs :-*​ Give the shape to the wing section, support the skin (prevent buckling) and act to prevent the fuel surging around as the aircraft maneuveres . They serve as attachment points for the control surfaces, flaps, undercarriage and engines. The ribs need to support the wing-panels, achieve the desired aerodynamic shape and keep it provide points for conducting large forces, add strength, prevent buckling, and separate the individual fuel tanks within the wing.​ There are many kinds of ribs:-​ · Form ribs:- consist of a sheet of metal, bent into shape.​ · Plate-type ribs:- consist of sheet-metal, which has upturned edges and weight-saving holes cut out into it. These ribs are used in conditions of light to medium loading.​ · Truss ribs :- consist of profiles that are joined together . These ribs may be suitable for a wide range of load-types.​ figure(2-3)Truss ribs​ · Closed ribs :-are constructed from profiles and sheet-metal, and are suitable for closing off sections of the wing. This rib is also suitable for a variety of loading conditions.​ figure (2-1) closed ribs​ · Forged ribs :- are manufactured using heavy press-machinery, and are used for sections where very high loads apply.​ · Milled ribs :- are solid structures, manufactured by milling away excess material from asolid block of metal, and are also used where very high loads apply.​ figure (2-2) milling ribs​ figure(2-4) types of ribs​ *stringers :-*​ the stringers on the skin panels run in the length of the wing, and so usually need to bridge the ribs. There are several methods for dealing with this problem. The stringers and ribs can both be un interrupted. The stringers now run over the rib. Leaving a gap between rib and skin . Rib and skin are indirectly connected, resulting in a bad shear load transfer between rib and skin. The stringers can be interrupted at the rib .interrupting the stringers in this way certainly weakens the structure, and therefore extra strengthening material, called a *doubler* , is usually added . Naturally, the stringers can also interrupt the rib . The stringers now run through holes cut into the rib, which also causes inevitable weakening of the structure.​ figure(2-5)​ The ribs also need to be supported, which is done by the:-​ *Spars:-*​ These are simple beams that usually have a cross-section similar to an I-beam. The spars are the most heavily loaded parts of an aircraft. They carry much more force at its root, than at the tip. Since wings will bend upwards, spars usually carry shear forces and bending moments.​ Aerodynamic forces not only bend the wing, they also twist it. To prevent this, the introduction of a second seems logical. Torsion now induces bending of the two spars, which is termed differential-bending. Modern commercial aircrafts often use two-spar wings where the spars are joined by a strengthened section of skin, forming the so-called torsion-box structure .The skin in the torsion-box structure serves both as a spar-cap (to resist bending), as part of the torsion box (to resist torsion) and to transmit aerodynamic forces.​ figure(2-6) types of spars​ Chapter three​ The pressure distribution​ Pressure distribution:-​ From experiments conducted on wind tunnel models and on full size airplanes, it has been determined that as air flows along the surface of a wing at different angles of attack there are regions along the surface where the pressure is negative, or less than atmospheric, and regions where the pressure is positive, or greater than atmospheric. This negative pressure on the upper surface creates arelatively larger force on the wing than is caused by the positive pressure resulting from the air striking the lower wing surface. Figure (3-1) shows the pressure distribution along an airfoil at three different angles of attack. In general, at high angles of attack the center of pressure moves forward, while at low angles of attack the center of pressure moves aft. In the design of wing structures this center of pressure travel is very important, since it affects the position of the air loads imposed on the wing structure in low angle of attack conditions and high angle of attack conditions. The airplane’s aerodynamic balance and controllability are governed by changes in the center of pressure.​ The center of pressure is determined through calculation and wind tunnel tests by varying the airfoil’s angle of attack through normal operating extremes. As the angle of attack through normal operating extremes. As the angle of attack is changed, so are the various pressure distribution characteristics fig. (3-1). Positive (+) and negative(-) pressure forces are totaled for each angle of attack and the resultant force is obtained. The total resultant pressure is represented by the resultant force vector shown in fig.(3-2)​ *figure (3-1) *​ figure(3-2)​ the point of application of this force vector is termed the “center of pressure” (cp). For any given angle of attack , the center of pressure is the point where the resultant force crosses the chord line. This point is expressed as a percentage of the chord of the airfoil. If the designer would place the wing so that its center of pressure was at the airplane’s center of gravity, the airplane would always balance. The difficulty arises, however, that the location of the center of pressure changes with change in the airfoil’s angle of attack fig.(3-3).​ figure(3-3)​ in the airplane’s normal range of flight attitudes, if the angle of attack is increased, the center of pressure moves forward , and if decreased, it moves rearward. Since the center of gravity is fixed at one point, it is evident that as the angle of attack increases, the center of lift (CP) moves ahead of the center of gravity , creating a force which tends to raise the nose of the airplane or tends to increase the angle of attack still more. On the other hand, if the angle of attack is decreases, the center of lift (CP) moves aft and tends to decrease the angle of greater amount. It is seen then, that the ordinary airfoil is inherently unstable, and that an auxiliary device, such as the horizontal tail surface, must be added to make the airplane balance longitudinally. The balance of an airplane in flight depends , therefore, on the relative position of the center of gravity (CG) and the center of pressure (CP) of the airfoil. Experience has shown that an airplane with the center of gravity in the vicinity of 20 percent of the wing chord can be madeto balance and fly satisfactorily.​ The tapered wing presents a variety of wing chords throughout the span of the wing. It becomes necessary then, to specify some chord about which the point of balance can be expressed. This chord , known as the mean aerodynamic chord(MAC), usually is defined as the chord of an imaginary untapered wing which would have the same center of pressure characteristics as the wing in question.​ Theoretical and experimental side​ Introduction:-​ The design of the airfoil is a complex and time consuming process and needs expertise in fundamentals of aerodynamics at graduate level. Since the airfoil needs to be verified by testing it in a wind tunnel, it is expensive too. Large aircraft production companies such as Boeing and Airbus have sufficient human experts (aerodynamicists) and budget to design their own airfoil for every aircraft, but small aircraft companies, experimental aircraft producers and homebuilt manufacturers do not afford to design their airfoils. Instead they select the best airfoils among the current available airfoils that are found in several books or websites.​ Steps of design the wing :-​ 1. Selecting the NACA of an airfoil (0012)​ 2. Using a profile program to calculate the coordinate of the airfoil and draw it ​ 3. By select the length of the chord (30 cm) and semi span (40 cm) of the airfoil we can draw the wing​ 4. Go with this data to the Carpentry workshop and manufacture the wing​ 5. In the Carpentry workshop making a taps (5 taps) on the upper and lower surface of the wing​ The coordinate of NACA 0012:-​ NACA 0012 AIRFOILS​ x/c y/c​  0.0000000 0.0000000​  0.0005839 0.0042603​  0.0023342 0.0084289​  0.0052468 0.0125011​  0.0093149 0.0164706​  0.0145291 0.0203300​  0.0208771 0.0240706​  0.0283441 0.0276827​  0.0369127 0.0311559​  0.0465628 0.0344792​  0.0572720 0.0376414​  0.0690152 0.0406310​  0.0817649 0.0434371​  0.0954915 0.0460489​  0.1101628 0.0484567​  0.1257446 0.0506513​  0.1422005 0.0526251​  0.1594921 0.0543715​  0.1775789 0.0558856​  0.1964187 0.0571640​  0.2159676 0.0582048​  0.2361799 0.0590081​  0.2570083 0.0595755​  0.2784042 0.0599102​  0.3003177 0.0600172​  0.3226976 0.0599028​  0.3454915 0.0595747​  0.3686463 0.0590419​  0.3921079 0.0583145​  0.4158215 0.0574033​  0.4397317 0.0563200​  0.4637826 0.0550769​  0.4879181 0.0536866​  0.5120819 0.0521620​  0.5362174 0.0505161​  0.5602683 0.0487619​  0.5841786 0.0469124​  0.6078921 0.0449802​  0.6313537 0.0429778​  0.6545085 0.0409174​  0.6773025 0.0388109​  0.6996823 0.0366700​  0.7215958 0.0345058​  0.7429917 0.0323294​  0.7638202 0.0301515​  0.7840324 0.0279828​  0.8035813 0.0258337​  0.8224211 0.0237142​  0.8405079 0.0216347​  0.8577995 0.0196051​  0.8742554 0.0176353​  0.8898372 0.0157351​  0.9045085 0.0139143​  0.9182351 0.0121823​  0.9309849 0.0105485​  0.9427280 0.0090217​  0.9534372 0.0076108​  0.9630873 0.0063238​  0.9716559 0.0051685​  0.9791229 0.0041519​  0.9854709 0.0032804​  0.9906850 0.0025595​  0.9947532 0.0019938​  0.9976658 0.0015870​  0.9994161 0.0013419​  1.0000000 0.0012600​  0.0000000 0.0000000​  0.0005839 -.0042603​  0.0023342 -.0084289​  0.0052468 -.0125011​  0.0093149 -.0164706​  0.0145291 -.0203300​  0.0208771 -.0240706​  0.0283441 -.0276827​  0.0369127 -.0311559​  0.0465628 -.0344792​  0.0572720 -.0376414​  0.0690152 -.0406310​  0.0817649 -.0434371​  0.0954915 -.0460489​  0.1101628 -.0484567​  0.1257446 -.0506513​  0.1422005 -.0526251​  0.1594921 -.0543715​  0.1775789 -.0558856​  0.1964187 -.0571640​  0.2159676 -.0582048​  0.2361799 -.0590081​  0.2570083 -.0595755​  0.2784042 -.0599102​  0.3003177 -.0600172​  0.3226976 -.0599028​  0.3454915 -.0595747​  0.3686463 -.0590419​  0.3921079 -.0583145​  0.4158215 -.0574033​  0.4397317 -.0563200​  0.4637826 -.0550769​  0.4879181 -.0536866​  0.5120819 -.0521620​  0.5362174 -.0505161​  0.5602683 -.0487619​  0.5841786 -.0469124​  0.6078921 -.0449802​  0.6313537 -.0429778​  0.6545085 -.0409174​  0.6773025 -.0388109​  0.6996823 -.0366700​  0.7215958 -.0345058​  0.7429917 -.0323294​  0.7638202 -.0301515​  0.7840324 -.0279828​  0.8035813 -.0258337​  0.8224211 -.0237142​  0.8405079 -.0216347​  0.8577995 -.0196051​  0.8742554 -.0176353​  0.8898372 -.0157351​  0.9045085 -.0139143​  0.9182351 -.0121823​  0.9309849 -.0105485​  0.9427280 -.0090217​  0.9534372 -.0076108​  0.9630873 -.0063238​  0.9716559 -.0051685​  0.9791229 -.0041519​  0.9854709 -.0032804​  0.9906850 -.0025595​  0.9947532 -.0019938​  0.9976658 -.0015870​  0.9994161 -.0013419​  1.0000000 -.0012600​ Pressure coefficient:-​


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## جنان حمزة (29 يونيو 2014)

شكرا لك اخ سامح


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## abo.alnoor.tec (28 يونيو 2015)

ملف رائع جزاكم الله خيراً
أتمنى وجود ملفات في كل جزء من الطائرة


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