Correct Answer : Distance between leading edge and trailing edge
Explanation : Distance between leading edge and the trailing edge is called chord. Chord refers to the imaginary straight line joining the leading edge and trailing edge of an airfoil. The chord length is the distance between the trailing edge and the point on the leading edge, where the chord intersects the leading edge.
Correct Answer : allow the wing to operate at high angle of attack
Explanation : Leading edge is a part of the airfoil. Leading edge allows the wing to operate at a high angle of attack. Slats are placed at the leading edge. These are aerodynamic surfaces on the leading edge of the wing. These are high lift devices used for short takeoff.
Correct Answer : NACA series
Explanation : The nomenclature of airfoil is designed by the NACA series. The shape of the NACA airfoil is described using a series of digits following the word NACA. The NACA identified different airfoils shape with a logical numbering system, such as symmetric airfoil and cambered airfoil.
Correct Answer : to increase maximum lift
Explanation : The main purpose of camber is to increase the maximum lift in an airfoil. The maximum lift coefficient can get by increasing the camber in an airfoil. Some recent design use negative camber. That airfoil is called the supercritical airfoil. This type of airfoil is used in the supersonic flight and to produce a higher lift to drag ratio.
Correct Answer : Because of no camber
Explanation : The NACA 0012 airfoil indicates the symmetric airfoil, because an airfoil with no camber, that is with the camber line and chord line co-incident in an airfoil is called a symmetric airfoil. For the NACA 0012 airfoil is with a maximum thickness of 12%.
Correct Answer : 15%
Explanation : In NACA 747A315 airfoil. The 7 denotes the series, 4 provides the location of minimum pressure on the upper surface in a tenth of chord, 7 provides the location of the minimum pressure on the lower surface in tenth of chord, the 4th character i.e. A indicates the thickness distribution and mean line form used. Again 5th digit 3 indicates the design lift coefficient in the tenth and the final two digits indicate the thickness of the chord 15%.
Correct Answer : Airflow rejoins
Explanation : Trailing edge is a part of airfoil. Trailing edge of an aerodynamic surface such as awing is its rear edge. Where the airflow will separate at the leading edge and rejoins at the trailing edge. Ailerons are placed at the trailing edge of the wing and rudder, elevator is placed at the trailing edge of tail.
Correct Answer : Circular
Explanation : The shape of the airfoil at the leading edge is usually circular with a leading edge radius of approximately 0.02c. The shape of all standard NACA airfoils are generated by specifying the shape of the mean chamber line and then wrapping specified symmetric thickness distribution around the mean camber line.
Correct Answer : 12%
Explanation : The NACA 2412 airfoil has a maximum camber of 2% located at 40% from the leading edge with a maximum thickness of 12%. It was the first 4-digit NACA series and it was developed in 1930’s because of the 12% thickness the shape of the airfoil is symmetric.
Correct Answer : Camber
Explanation : Camber is generally used to increase the maximum lift coefficient. Which in turn it decreases the stall speed of the aircraft. Due to decrease in the stall there will be an increase in the lift. In supercritical airfoil. There will be a highly cambered curve after section which used in supersonic flight for their maximum lift.
Correct Answer : Dividing the lift coefficient by the drag coefficient
Explanation : The lift-drag ratio is used to express the relation between lift and drag, and is obtained by dividing the lift coefficient by the drag coefficient. We can get the maximum lift coefficient by the drag coefficient for a given airfoil. The characteristics of any airfoil section can conveniently be represented by a graph showing the lift-drag ratio.
Correct Answer : Coefficient of lift-coefficient of drag ratio
Explanation : The characteristics of any particular airfoil section can conveniently be represented by a graph, showing the amount of lift and drag obtained at a various angle of attack, the lift-drag ratio, and the movement of the center of pressure. With these graphs, we can choose suitable airfoil for the aircraft.
Correct Answer : The curve increases maximum at a particular section and decreases rapidly in the same section
Explanation : The lift curve reaches its maximum for any wing section at certain degree of angle of attack and then rapidly decreases at some degree of angle of attack is therefore the stalling angle. Stalling will occurs at particular degree were the maximum lift will be achieved by the aircraft at that point stalling will takes place.
Correct Answer : Drag will increase rapidly at particular degree of angle of attack and overcomes the lift curve at particular degree of angle of attack
Explanation : The drag curve increases very rapidly at a certain degree of angle of attack and completely overcomes the lift curve at some degree of angle of attack, at this point maximum drag will be caused.so, at this point lift will drastically decrease and hence producing the drag over the body.
Correct Answer : With supplied power
Explanation : A propeller creates a thrust force out of the supplied power. The magnitude of this force is not for a given propeller but depends on the velocity of the incoming air and the rotational velocity of the propeller and helps the propeller to create the forward thrust.
Correct Answer : True
Explanation : An airfoil is the shape of a wing, blade of a propeller, rotor, and blades of a turbine. A propeller is a type of fan that transmit power by converting rotational motion in its thrust.A pressure difference is produced between the forward and rear surface of the airfoil-shape and fluid is accelerated behind the blade.
Correct Answer : Side by side
Explanation : The flow in the planes perpendicular to the vortex filament at zero are identical to the flow induced by the point vortex of strength. An infinite number of straight vortex filaments side by side. Where the strength of each filament is infinite small. These vortex filament are arranged side by side from the vortex sheet.
Correct Answer : Zero
Explanation : Kelvin’s circulation theorem, is the absence of external forces on the sheet, the circulation between any two material points in the sheet remains conserved at zero. The equation of motion of the sheet can be rewritten in terms of circulation and by a change of variable.
Correct Answer : To reduce the dimensionality of aerodynamic model
Explanation : Vortex method is extensively applied to reduce the dimensionality of these aerodynamic models based on the proper estimation of the strength and distribution of the vortices in the wake this condition can be readily applied to the flat plate or an airfoil with a cusped trailing edge were the flow will leave smoothly.
Correct Answer : False
Explanation : The vortex sheet solution as given by the birkoff-roff equation cannot go beyond the critical time. The spontaneous loss of analyticity in a vortex sheet is a consequence of mathematical modeling since a real fluid with viscosity, however, small will never develop singularity.
Explanation : The kutta condition is a principle in steady-flow fluid dynamics, especially aerodynamics that is applicable to solid bodies with sharp corners, such as the trailing edge of the airfoil. It is named for German mathematician and aerodynamicist martin Wilhelm kutta.
Explanation : When a smooth symmetric body, such as a cylinder with oval cross-section moves with zero angle of attack through a fluid it generates no lift. There are two stagnation points on the body one at the front and the other at the back. Since no lift will be generated by the cylinder at zero angle of attack.
Correct Answer : Trailing edge
Explanation : Vortex flow occurs at the trailing edge and because the radius of the sharp trailing edge is zero, the speed of the air around the trailing edge should be infinitely fast, real fluid cannot move at infinite speed, they can move extremely fast finite speed.
Correct Answer : Due to viscous forces
Explanation : The stagnation point on the topside of the airfoil then moves until it reaches the trailing edge. The starting vortex eventually dissipates due to viscous forces. As the airfoil continues on its way, there is a stagnation point at the trailing edge. The flow over the topside of the airfoil conforms to the upper surface of the airfoil.
Correct Answer : Flow over the topside
Explanation : As the vorticity increases the bound vortex and also progressively increases and causes the flow over the topside of the airfoil to increase in speed. The starting vortex is soon cast-off the airfoil and is left behind, spinning in the air, where the airfoil left it.
Correct Answer : As the airfoil begins to move vortex are formed
Explanation : As the airfoil begins to move it carries this vortex, known as the starting vortex along with it, pioneering aerodynamicists were able to photograph starting vortices in liquids to confirm their existence. The vorticity in the starting vortex is matched by the vorticity in the bound vortex in the airfoil.
Correct Answer : Kutta Condition
Explanation : According to the Kutta Condition, for a given angle of attack, the value of circulation around the airfoil is such that the flow leaves the trailing edge smoothly. The velocity at the trailing edge is dependent on the shape of the airfoil.
Correct Answer : Kelvin’s Circulation Theorem
Explaination : For the same fluid elements in closed curves C1 and C2, the circulation remains constant with time as the fluid elements move downstream. This is essentially Kelvin’s circulation theorem.
Correct Answer : Inviscid, Compressible Barotropic Flow
Explanation : Kelvin’s Theorem is applicable for the special case of barotropic flow while dealing with inviscid, compressible flows.
Correct Answer : There is no boundary layer formation, hence no vorticity
Explaination : For inviscid flows, the boundary layer is not formed. Therefore, in the regions of high velocity, high viscosity is not there and hence no vortex can form. Thus, there is no lift produced. Starting vortex cannot form in inviscid medium and in the viscous medium it dies due to viscosity.
Correct Answer : Frozen Vortex Lines
Explanation : Kelvin’s theorem can be used to prove Helmholtz theorems, one of which says ‘vortex lines move with the fluid’ which is what is known as “frozen vortex lines”.
Correct Answer : Barotropic Flow
Explanation : A barotropic flow is a fluid where density is a function of pressure only, i.e. ρ = ρ(p). Baroclinic flow is the fluid which is not only dependent on the pressure but on other factors also. Viscous and inviscid flows are not necessarily dependent on pressure only.
Correct Answer : generation of lift and circulation
Explanation : From Kelvin’s Theorem, circulation remains constant with time. So for initial zero circulation, the formation of starting vortex means there has to be equal and opposite circulation in the form of lift.
Correct Answer : Kutta Condition and Kelvin’s Theorem
Explanation : Kutta condition enforces smooth flow at the trailing edge. In doing so high velocity gradients formed at the trailing edge generates vorticity and hence circulation is there. From Kelvin’s circulation theorem starting vortex is formed to conserve circulation.
Correct Answer : Due to Viscosity
Explanation : Starting vortex cannot form in inviscid medium. It can form only in a viscous medium. In a viscous medium, it dies instantly due to viscous effects.
Explanation : The symmetric airfoil with good lift to drag ratio is used for an aircraft wing, this will be translated into lower fuel consumption, shorter take-off and landing times, and shorter runways. Airfoil With backward facing will get high lift coefficient.
Explanation : The NACA 0012 is symmetric airfoil because the mean camber line and chord line intersect in the same line.so, there is no camber in the NACA 0012 airfoil. That the reason the first and second digits are will become zero, and the thickness of NACA 0012 airfoil is 12 percent.
Correct Answer : 2%
Explanation : The thickness of the NACA 0002 airfoil is 2 percent. There is no camber in the NACA 0002 airfoil because of the mean camber line and chord line intersect in the same line, so the position maximum camber in the NACA 0002 airfoil is zero.
Correct Answer : 5
Explanation : The wake structure is classified into five different modes according to their pattern obtained from instantaneous and mean vorticity fields by also taking into account the amplitude spectrum of the lift coefficient.
Correct Answer : Perpendicular to the direction of motion
Explanation : 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 lift will oppose the motion and it produces the lift.
Correct Answer : Thin Airfoil
Explanation : Thin airfoils can be simulated as a vortex sheet kept along the camber line. Thick airfoils cannot be represented like this. Negatively cambered airfoils may or may not be thin and in that case, this representation might be incorrect. NACA airfoils have all types of airfoils, not thin airfoils only.
Correct Answer : γ(TE) = 0
Explaination : According to the Kutta condition, the condition for flow leaving smoothly is satisfied always by γ(TE) = 0. The other options may not be true as per the Kutta Condition.
Correct Answer : 15+4=0
Explaination : In thin airfoil theory, the camber line has to be a streamline. Thus, the component of velocity in the normal direction to the camber line is zero at any given point. This means the sum of velocity components by the free stream velocity and the induced velocity is zero.
Correct Answer : Streamline
Explanation : For thin airfoils, camber and chord lines are very close and the vortex sheet can be assumed to be placed at the chord line. And the camber line becomes a streamline.
Correct Answer : tan-1?-\(\frac {dz}{dx}\)
Explaination : For a camber distribution where z=z(x), at a given point on the camber line the missing angle is the slope at that point, which is tan-1?-\(\frac {dz}{dx}\) which is independent of α.
Correct Answer : Camber and Chord Line
Explaination : For a symmetric airfoil(\(\frac {dz}{dx}\)=0), camber line is the same as chord line. Therefore, the best option is Camber and Chord Line. Only Camber or Chord Line is an incomplete answer.
Correct Answer : π
Explaination : This is a standard integral which is used many times in the study of airfoils. This is used to solve the transformed fundamental equation in thin airfoil theory where the chord length is expressed as θ=0 to θ=π.
Correct Answer : γ(c)=0
Explaination : The Kutta Condition says the flow leaves the trailing edge smoothly. The trailing edge is at chord length c for a thin airfoil which gives γ(c)=0. Also, γ(x)=0 or γ(ξ)=0 is not true for all x or ξ. \(\frac {dz}{dx}\)=0 is the definition of the symmetric airfoil, not Kutta condition.
Correct Answer : γ(θ)=0
Explaination : The transformation for thin airfoil theory uses ξ=\(\frac {c}{2}\)(1-cosθ) to make the coordinate transformation, where 0≤θ≤π. For a symmetrical airfoil (\(\frac {dz}{dx}\)=0), while γ(θ)=0 is not true for all values of θ.
Correct Answer : Kutta condition is satisfied at the trailing edge
Explanation : Kutta condition is always satisfied at the trailing edge. This is the boundary condition for the camber line to be a streamline also and is not an approximation. While the other three statements are valid approximations, which have been made under the assumption of a thin airfoil where camber and chord line is very close.
Correct Answer : ii
Explaination : V∞ is the free-stream velocity. Here i is not a component of free-stream velocity, iii is the parallel component and ii is the normal component for the camber line.
Explanation : Camber is usually designed into an airfoil is to increase the maximum lift coefficient. This minimizes the stalling speed of aircraft using airfoil. Aircraft with wings based on cambered airfoils usually have low lower stalling speeds than similar aircraft with wings based on symmetric airfoils.
Correct Answer : Reduce the camber
Explanation : An aircraft designer may also reduce the camber of the outboard section of the wings to increase the critical angle of attack at the wing tips. When the wing approaches the stall angle this will ensure that the wing root stalls before the tip, giving the aircraft resistance to spinning and maintaining ailerons effectiveness close to the stall.
Correct Answer : Camber line curves backup near the trailing edge
Explanation : An airfoil where the camber line curves back up near the trailing edge is called a reflexed camber airfoil. Such an airfoil is useful in certain situations, such as with tailless aircraft, because the moment about the aerodynamic center of the airfoil is zero.
Explanation : The camber of an airfoil causes an increase in velocity and a consequent decrease in pressure of the stream of air moving over it due to the increase in velocity it gets a maximum coefficient of lift. Hence it is used in military aircraft.
Correct Answer : Always false
Explanation : The thin airfoil theory solution when subjected to the Kutta condition makes the camber line as a streamline of the flow, irrespective of the airfoil being symmetrical or cambered.
Correct Answer : The slope of the camber line is zero
Explanation : Slope of the camber line \(\frac {dz}{dx}\) is not zero for a cambered airfoil but is some finite value. All the other statements are valid assumptions of thin airfoil theory.
Correct Answer : Symmetric airfoils only
Explaination : The original fundamental equation of thin airfoil theory is \(\frac {1}{2\pi } \int_0^c \frac {\gamma(\xi)d\xi}{x-\xi}\)=V∞(α-\(\frac {dz}{dx}\)). For the symmetric airfoils, \(\frac {dz}{dx}\)=0 and so \(\frac {1}{2\pi } \int_0^c \frac {\gamma(\xi)d\xi}{x-\xi}\)=V∞α is valid. While for the cambered airfoils \(\frac {dz}{dx}\) is non-zero.
Correct Answer : Fourier sine series
Explanation : The cambered airfoil solution of the thin airfoil theory is different from that of symmetric airfoils with the addition of a Fourier sine series term.
Correct Answer : 0.04c, 0.4c
Explanation : For NACA 4 digit airfoils the first digit gives the maximum camber in 100th parts of the chord length c and the second digit gives the position of maximum camber in 10th parts of the chord length from the leading edge.
Correct Answer : Symmetrical airfoil
Explanation : The first two digits in the NACA nomenclature give the maximum camber and position of maximum camber. For a symmetric airfoil, both of these are zero.
Correct Answer : An depends on chord length of the airfoil
Explaination : The given equation is the fundamental equation of thin airfoil theory. For a cambered airfoil, the solution is in the form γ(θ)=2V∞(A0\(\frac {1+cos\theta }{sin?\theta }\) + Σ\(_{n=1}^∞\)sin? nθ An) where An depends on the slope of camber line and A0 depend both on the slope of camber line and angle of attack.
Correct Answer : An=\(\frac {2}{\pi }\int_0^{\pi }\frac {dz}{dx}\) cos?n∅ d∅
Explaination : From the general solution of thin airfoil theory we have An=\(\frac {2}{\pi }\int_0^{\pi }\frac {dz}{dx}\) cos?n∅ d∅ and A0=α-\(\frac {1}{\pi }\int_0^{\pi }\frac {dz}{dx}\) cos?n∅ d∅ where the limits are 0≤∅≤π.
Correct Answer : 0
Explaination : The lift per unit span is given by the formula L’=Γρ∞V∞ by Kutta-Joukowski Theorem. Putting the respective values given in the question, L’=102.55\(\frac {N}{m}\) (unit is N/m, not N).
Correct Answer : Experiment data
Explanation : The earlier standard NACA airfoils were based exclusively on the experimental results from 1930s-40s. Later on, numerical techniques using the computer were used followed by wind-tunnel testing to develop modern airfoils.
Correct Answer : Partially true
Explanation : The modern low-speed airfoils were developed using numerical techniques on the computer which was followed by wind- tunnel testing to confirm the computer results. This gave the definite airfoil properties for the new airfoils.
Correct Answer : Handling flow separation effects
Explanation : The use of computers led to the design of better airfoils since it made possible to get the definitive properties of the airfoils. This had many advantages like a higher coefficient of lift and shape to tackle the flow separation effects at high angles of attack.
Correct Answer : NACA LS (1)-0407
Explanation : The new airfoils are the low-speed airfoils (designated by LS). So NACA LS (1)-04XX are the new airfoils while the NACA XXXX are the standard airfoils.
Correct Answer : 50%
Explaination : The lift coefficient of 1 is vital for the aviation sector. The new low-speed airfoils developed had higher L/D ratios. For a lift coefficient equal to 1, the increase was about 50%.
Correct Answer : Supercritical airfoil
Explanation : The GA (W)-1 (also known as Whitcomb airfoil) airfoil was the first low-speed airfoil obtained under the new airfoils. It led to the development of supercritical airfoils which had an almost similar shape. The supercritical airfoils had lesser drag at high subsonic speeds, which was a major performance improvement.
Correct Answer : Higher symmetry
Explanation : The larger leading edge radius gave a reduced peak in pressure coefficient at the leading edge. The trailing edge was cusped which increased camber, thus decreasing symmetry. These features reduced flow separation and gave a higher value of maximum lift coefficient.
Correct Answer : Panel method and advanced viscous flow solutions
Explanation : The numerical methods used were like source and vortex panel methods and numerical predictions of the viscous flow behavior, to analyze skin friction and flow separation effects. This was followed by experimental testing for verification of computer results.
Explanation : At low angle of attack, a thin viscous region forms over the airfoil, and grows from the leading edge to the trailing. On the upper surface, where adverse pressure gradients exists the boundary layer grows more rapidly.
Explanation : At higher angles, very large adverse pressure gradients that develop on the upper side as the airfoil attempts to generate more lift causes the boundary layer to separate, leading to a major disruption of the flow over the airfoil and the wing stalls.
Correct Answer : Smoking rises from cigarette
Explanation : Smoke rising from a cigarette is mostly turbulent flow. However, for the first few centimeters, the flow is laminar. The smoke plume becomes turbulent as its Reynolds number increases, due to its flow velocity and the characteristic length.
Explanation : The effect can also be exploited by devices such as aerodynamic spoilers on aircraft, which deliberately spoil the laminar flow to increase the drag and reduce the lift on the aircraft, which also used as a break to stop the aircraft
