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The correct OPTION is (c) Panel method and advanced VISCOUS flow solutions

Best 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.

The correct choice is (a) True

Easiest explanation: For the attached flow, the PRESSURE acting at the rear end of the airfoil counteracts the pressure acting on the front edge. But once flow SEPARATION occurs, the rear pressure reduces causing a NET backward pressure – pressure drag.

The correct option is (a) Reduce the camber

To explain I WOULD say: 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.

The CORRECT ANSWER is (a) False

Easiest explanation: The coefficient of moment about the quarter-chord is not zero for a CAMBERED airfoil (cm,l/4=\(\frac {\pi }{4}\)(A2-A1)). The center of pressure is the point about which the total moment is zero. Therefore, this STATEMENT is false.

Right choice is (b) Negative

Easy EXPLANATION: A positive CIRCULATION round an AIRFOIL, this implies that there must be negative circulation round the airfoil. This has been observed experimentally. The vortex shedding continues until the circulation round the airfoil is sufficient to make the main irrotational flow smooth at the TRAILING EDGE.

Right answer is (a) True

To elaborate: A DYED circuit, which is large enough to have been clear of all these regions SINCE the START of the motion. As the original state was ONE of rest, the circulation round that circuit will still be zero for time t.Therefore if we SKETCH in a line an instantaneous line in space at time t such that the curve encloses the airfoil but not the starting vortex.

Right option is (a) There is no BOUNDARY layer FORMATION, hence no vorticity

For explanation I would SAY: 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.

Right ANSWER is (b) False

For explanation I would SAY: The LIFT and drag of an airfoil DEPEND not only on the angle of attack but also on the shape of the airfoil. The lift coefficient and drag coefficient depend on the shape of the airfoil and will alter with changes in the angle of attack and other WING appurtenance.

Correct OPTION is (a) True

Best explanation: Flow over a GOLF ball, CONSIDERING the golf ball to be stationary, with air flowing over it. If the golf ball to be stationary, with air flowing over it. If the golf ball were SMOOTH, the BOUNDARY layer flow over the front sphere would be laminar at the typical condition.

The correct choice is (b) Trailing edge

The explanation: The BOUNDARY layer thickness increases parabolically with the DISTANCE MEASURED from the leading edge (denoted by X) for incompressible, laminar flow. Therefore, it is LARGEST at the trailing edge.

Correct answer is (b) True

The BEST I can EXPLAIN: The assumption of inviscid flow does not affect the lift since the SHEAR force component ALONG the lift direction is negligible. Only the pressure distribution is responsible for CREATING lift.

Correct option is (c) HANDLING flow separation EFFECTS

The explanation is: 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 option is (b) True

Best EXPLANATION: The new low-speed AIRFOILS had considerably higher maximum LIFT coefficients (around 30% higher) when compared to the standard NACA airfoils.

The correct choice is (a) True

Best 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 choice is (a) False

To elaborate: Aerodynamic center is the point where the pitching MOMENT remains constant with changing angle of attack. It is GENERALLY the quarter-chord for an airfoil. Center of pressure is the point where the RESULTANT of FORCES act and the moment at that point will change with the change of angle of attack. Thus, the center of pressure will change and may not be the quarter-chord always.

Right answer is (C) cm,le=-\(\frac {c_l}{2}\)

Explanation: The coefficient of moment about the leading edge is given by cm,le=-π \(\frac {\ALPHA }{2}\). Putting cl = 2πα we GET cm,le=-\(\frac {c_l}{4}\). FINDING the moment coefficient about quarter chord we get,

cm,c/4=cm,le+\(\frac {c_l}{4}\).

The correct OPTION is (d) γ(θ)=0

Best explanation: 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.

The correct choice is (a) True

The explanation: The airfoil with a sharp trailing edge begins to move with a positive angle of attack through air, the two stagnation points are initially located on the underside near the LEADING edge and on the TOPSIDE near the trailing edge, just as with the cylinder, as the air passing the underside of the airfoil REACHES the trailing edge, it must flow AROUND the trailing edge on the top SIDE of the airfoil.

Correct option is (B) False

For explanation I would SAY: 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.

The correct option is (d) 1.33

Explanation: The COEFFICIENT of skin-friction DRAG is related to Reynolds NUMBER as Cf=\(\frac {1.328}{\sqrt{Re_c}}\) where REC is the Reynolds number at the trailing EDGE and the coefficient gives half the total skin-friction drag. The required constant is 1.328 for the case of a laminar flow.

Right CHOICE is (a) True

Explanation: The growth of the BOUNDARY layer the phenomenon of stall and the LOCATION of the separation point THUS play a critical role. This course attempts to develop theories and methods by which these quantities can be computed, given the pressure distribution over the airfoil.

The correct option is (a) True

Easy EXPLANATION: Wind tunnel tests were then conducted to verify the computer designed PROFILES and to obtain the definitive AIRFOIL PROPERTIES, out of this WORK first came the general aviation whit comb airfoil. Which has since been predesignated the LS-0417 airfoil.

Correct OPTION is (d) ∞ (i.e. the point does is not on the AIRFOIL)

To EXPLAIN: The center of pressure of a cambered airfoil will leave the airfoil and move at a very large distance away from the airfoil. This is one reason why the AERODYNAMIC center is preferred over the center of pressure since it is always at the quarter-chord but the LATER keeps on changing with the angle of attack.

Right choice is (c) Γ=cαV∞\(\int_0^c\)(1+cosθ)dθ

The best explanation: Using the transformation ξ=\(\frac {c}{2}\)(1-cosθ), where 0≤θ≤π, CORRESPONDING to 0≤ξ≤c in γ(θ) and INTEGRATING gives the TOTAL circulation Γ.

The correct option is (a) True

For 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.

Right answer is (a) True

The BEST explanation: Viscosity ACTS a smoothing parameter in areal FLUID .these have been extensive studies on a vortex sheet, most of them by discrete or POINT vortex approximation with or WITHOUT using a point vortex approximation and delta-regularization is obtained by smooth roll-up of a vortex sheet into a double branched spiral.

The correct ANSWER is (a) False

Explanation: The lift-drag ratio will reach its maximum at zero DEGREE angle of attack, at this angle we OBTAIN the most lift for less amount of drag.so, that the AIRCRAFT will GET maximum amount of lift at zero angle of attack.

The correct option is (a) allow the wing to operate at high angle of ATTACK

To EXPLAIN I would say: 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.

Right option is (b) to increase maximum lift

The 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.

Right choice is (b) Inversely proportional to \(\sqrt[5]{Re_x}\)

EASY explanation: The boundary layer THICKNESS for an incompressible, turbulent FLOW over a FLAT plate, at a distance x where Reynolds number is Rex is given by δ=\(\frac {0.37x}{\sqrt[5]{Re_c}}\). Thus, the thickness is indirectly proportional to \(\sqrt[5]{Re_x}\).

Correct choice is (b) True

Easy explanation: As the Reynolds number increases, such as by increasing the FLOW RATE of the fluid, the flow will transition from laminar to TURBULENT flow at a specific range of Reynolds number, the laminar-turbulent transition range depending on SMALL disturbance LEVELS in the fluid.

Right option is (B) True

For explanation I would say: Turbulence is commonly observed in everyday phenomena such as surf, fast flowing RIVERS, billowing STORM clouds from a CHIMNEY and most fluid flows occurring in nature and created in engineering APPLICATIONS are turbulent.

Right answer is (a) True

To ELABORATE: The Reynolds number is very small, much less than 1, then the fluid will exhibit stokes, where the viscous FORCES of the fluid DOMINATE the inertial forces. The specific CALCULATION of the Reynolds number, and the values where laminar flow occurs.

The correct OPTION is (b) True

To explain: In laminar flow, the motion of the PARTICLES of the fluid is very orderly with particles close to a SOLID surface moving in straight lines parallel to that surface. Laminar flow is a flow regime characterized by HIGH momentum diffusion and LOW momentum convection.

Right choice is (a) True

For explanation I would say: VISCOUS effect play a major role in other flows, of intersect such as flow through compressors, turbines and diffusers and in NON- AEROSPACE applications as WELL. In all these applications, the raised by the designer and the engineer are often the same.

Right answer is (a) True

The explanation is: During 1970s, NASA designed a series of low-speed airfoils that have performance superior to the EARLIER NACA airfoils. The standard NACA airfoils were based ALMOST EXCLUSIVELY on experimental DATA obtained during 1930s and 1940s.

The correct answer is (C) 50%

For explanation: 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%.