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This section includes InterviewSolutions, each offering curated multiple-choice questions to sharpen your knowledge and support exam preparation. Choose a topic below to get started.

51.

The Kutta Condition according to the thin airfoil theory is_____(a) γ(c)=0(b) γ(x)=0(c) γ(ξ )=0(d) \(\frac {dz}{dx}\)=0This question was addressed to me in a job interview.I'm obligated to ask this question of The Symmetric Airfoil in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right choice is (a) γ(c)=0

The explanation is: 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.

52.

The relation between lift coefficient and moment coefficient about the leading edge is_______(a) cm,le=-(\(\frac {c_l}{2}+\frac {\pi }{4}\)(A1-A2))(b) cm,le=-(\(\frac {c_l}{4}+\frac {\pi }{4}\)(A1-A2))(c) cm,le=-(\(\frac {c_l}{4}\)+(A1-A2))(d) cm,le=(\(\frac {c_l}{2}+\frac {\pi }{4}\)(A1-A2))The question was asked in semester exam.My enquiry is from The Cambered Airfoil in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct choice is (B) cm,LE=-(\(\FRAC {c_l}{4}+\frac {\pi }{4}\)(A1-A2))

To explain I would say: The lift coefficient is cl=π(2A0+A1). When we put this in the formula for moment coefficient about the leading edge, cm,le=-0.5π(A0+A1-0.5A2) we GET the desired relation.

53.

What is the thickness of NACA 0002 airfoil?(a) 3%(b) 1%(c) 0%(d) 2%I had been asked this question in exam.My question is based upon The Symmetric Airfoil topic in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct choice is (d) 2%

Easiest 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.

54.

Generation of lift over an airfoil and formation of starting vortex is correctly explained by which of these?(a) Kutta-Joukowski Theorem(b) Kutta Condition and Kelvin’s Theorem(c) Kutta-Joukowski Theorem and Kelvin’s Theorem(d) Kutta Condition and Helmholtz TheoremI got this question in a job interview.The query is from Kelvin’s Circulation Theorem and the Starting Vortex topic in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct choice is (b) KUTTA CONDITION and Kelvin’s Theorem

To elaborate: 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.

55.

How starting vortex is formed?(a) As the airfoil begins to move vortex are formed(b) As the airflows on the airfoil vortex are formed(c) As the airfoil moves against the relative wind vortex are formed(d) As the airflows on the circular body vortex are formedThe question was posed to me in unit test.This intriguing question comes from The Kutta Condition topic in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct option is (a) 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.

56.

What is a chord?(a) Distance between leading edge and chord(b) Distance between chord and chamber(c) Distance between leading edge and trailing edge(d) Distance between trailing edge and chordThis question was posed to me in an interview for internship.My question is from Airfoil Nomenclature in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right answer is (c) 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.

57.

Is flat vortex sheet with periodic boundaries can be used in high Reynolds number?(a) True(b) FalseThis question was posed to me by my school principal while I was bunking the class.Query is from The Vortex Sheet topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct answer is (a) True

Best EXPLANATION: A flat vortex sheet with periodic boundaries in the stream wise direction can be used to model a temporal free SHEAR layer at high REYNOLDS number. Assume that the INTERVAL between the periodic boundaries is of LENGTH ‘1’, then the flat vortex sheet with periodic boundary can be used in a free shear layer at high Reynolds number.

58.

Is surface is perturb the boundary layer and promote turbulence?(a) True(b) FalseThe question was asked during an interview.I'm obligated to ask this question of Turbulent Flow in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct option is (a) True

Explanation: The surface is dimpled to perturb the boundary layer and PROMOTE a transition to TURBULENCE. This results in higher skin FRICTION, but MOVES the point of boundary layer separation further along, resulting in lower form drag and lower overall drag.

59.

Example of turbulent flow?(a) Smoking rises from cigarette(b) Flow on a symmetric airfoil(c) Laminar flow(d) Turbulent flow on the airfoilI got this question in final exam.The origin of the question is Turbulent Flow topic in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right OPTION is (a) Smoking RISES from cigarette

To explain I would say: 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.

60.

Is Reynolds number is used to predict the turbulence?(a) True(b) FalseThis question was addressed to me in an online quiz.My question is taken from Turbulent Flow topic in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

The CORRECT answer is (a) True

For explanation: The turbulence can be predicted by a dimensionless constant called Reynolds number, which calculates the balance between kinetic energy and viscous DAMPING in a fluid flow. However, turbulence has long resisted detailed PHYSICAL ANALYSIS, and the INTERACTIONS within turbulence create a very complex situation.

61.

Is unsteady vortices appears at turbulent flow?(a) True(b) FalseI got this question during a job interview.I'm obligated to ask this question of Turbulent Flow in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right option is (a) True

The explanation: In turbulent FLOW, unsteady VORTICES appears of MANY sizes which INTERACT with each other, consequently drag due to friction effects increases. This would increase the energy needed to PUMP fluid through a pipe, for instance.

62.

Which of the following is not related to flow separation?(a) Reduction of lift(b) Increased drag(c) Adverse pressure gradients(d) Attached flowThis question was posed to me during a job interview.My enquiry is from Viscous Flow in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

The CORRECT choice is (d) ATTACHED flow

The BEST explanation: The region of adverse pressure GRADIENT near the trailing edge causes flow separation, which consequently causes lesser lift and higher DRAG. The flow is no more attached when flow separation occurs.

63.

Is pressure coefficient distributions are calculated for flow over an airfoil?(a) True(b) FalseI have been asked this question during an online exam.My enquiry is from Modern Low Speed Airfoils in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct choice is (a) True

The best I can EXPLAIN: The unstructured MESH for the numerical calculation of the flow over an airfoil to calculate for pressure COEFFICIENT distributions, airfoil shapes that support the specified pressure distribution is obtained, as given by the circle. The initial airfoil shape is ALSO shown in constant scale.

64.

Which of the following is a new NACA airfoil?(a) NACA 2212(b) NACA LS (1)-0407(c) NACA 0021(d) Both NACA LS (1)-0407 and NACA 2412I had been asked this question in an interview for internship.Query is from Modern Low Speed Airfoils topic in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct OPTION is (b) NACA LS (1)-0407

The BEST 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.

65.

The standard NACA airfoils were based on ____(a) Computer modeling results(b) Numerical techniques and wind- tunnel testing(c) Experiment data(d) Theoretical dataI have been asked this question in class test.Question is from Modern Low Speed Airfoils topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct choice is (c) Experiment data

The EXPLANATION is: 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.

66.

The lift per unit span for a thin, cambered airfoil with α=5°, A0=0.65, A1=1 is____(a) L’ = cπ\(V_∞^2\)ρ∞ (1.15)(b) L’ = cπ\(V_∞^2\)ρ∞ (1.65)(c) L’ = cπ\(V_∞^2\)ρ∞ (-0.35)(d) L’ = cπ\(V_∞^2\)ρ∞ (0.15)This question was addressed to me in a job interview.The question is from The Cambered Airfoil topic in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct choice is (a) L’ = cπ\(V_∞^2\)ρ∞ (1.15)

Easy EXPLANATION: For a thin CAMBERED AIRFOIL lift per unit SPAN is given by L’ = cπ\(V_∞^2\)ρ∞ (A0+\(\frac {1}{2}\)A1). Thus, with the given value of A0 and A1 we get L’ = cπ\(V_∞^2\)ρ∞ (1.15).

67.

The zero-lift angle of attack for a thin, cambered airfoil increases in magnitude with the increasing camber of the airfoil.(a) True(b) FalseThe question was asked by my school principal while I was bunking the class.This key question is from The Cambered Airfoil topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct ANSWER is (a) True

To explain I would say: USING the formula for αL=0=\(\frac {-1}{\pi}\int_0^{\pi}\frac {dz}{dx} \)(cos⁡∅-1)d∅, the DEPENDENCE of the zero-lift angle of ATTACK on camber can be seen. It is clear that for higher cambered AIRFOILS, αL=0 shifts to the negative side.

68.

Is thin sheet of intense vorticity is unstable at downstream?(a) True(b) FalseThis question was posed to me during an online exam.My question is from Kelvin’s Circulation Theorem and the Starting Vortex in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct answer is (a) True

To elaborate: As flow moves to the downstream, this THIN SHEET of intense vorticity is unstable, and it tends to roll up and form a picture SIMILAR to a point vortex. This vortex is called the STARTING vortex and is similar to the point vortex.

69.

Mathematically, what is meant by Kelvin’s Circulation Theorem for an inviscid and incompressible flow? (For the same set of fluid elements moving in a closed curve along with the fluid).(a) DΓ/Dt = 0(b) γ1 ≥ Γ2, where 1 is the upstream direction(c) Γ = -∮C1V.ds(d) Γ = 0The question was asked in homework.The query is from Kelvin’s Circulation Theorem and the Starting Vortex in section Incompressible Flow over Airfoils of Aerodynamics

Answer» CORRECT ANSWER is (a) DΓ/Dt = 0

To explain: Kelvin’s theorem states that the time rate of change of circulation for a SET of fluid elements in a closed curve moving along the fluid is zero. In other WORDS, circulation remains CONSTANT, not zero. Γ = -∮C1V.ds is the definition of circulation.
70.

It is possible to have lift without friction (i.e. in an inviscid medium).(a) True(b) FalseThe question was asked in an internship interview.My enquiry is from Kelvin’s Circulation Theorem and the Starting Vortex topic in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

The CORRECT option is (b) False

To explain I would SAY: Nature enforces KUTTA Condition through the means of the boundary layer (FRICTION). If there were no friction Kutta condition would not be ACHIEVED and there will be no lift.

71.

Turbine blades are used to extract energy from the high temperatures.(a) False(b) TrueThe question was posed to me in examination.Enquiry is from Airfoil Characteristics topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct OPTION is (b) True

To elaborate: The turbine blades are responsible for extracting energy from the HIGH temperature, high pressure gas PRODUCED by the compressor to survive in this difficult environment, turbine blades often use EXOTIC material like super alloy and MANY different methods of cooling. Such as boundary layer cooling and thermal barrier coatings.

72.

Is axial compressor is a compressor that can continuously pressurize gasses?(a) True(b) FalseI have been asked this question in an online quiz.My doubt stems from Airfoil Characteristics in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct answer is (a) True

Explanation: An axial compressor is a compressor that can continuously PRESSURIZE gases. It is a ROTATING airfoil based compressor which the gas principally flows PARALLEL to the axis of rotation .this diffuse from other rotating compressor. Such as centrifugal compressor, axial centrifugal compressor and mixed flow compressors .where the fluid flow will INCLUDE a radial component through the compressor.

73.

What parameter is used to decrease the stall speed?(a) Chord(b) Chord line(c) Camber(d) Camber lineThis question was posed to me in my homework.Question is from Airfoil Nomenclature topic in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct answer is (c) 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.

74.

The transition point for a flow can be found using____(a) Average of Reynolds numbers at leading and trailing edges(b) It is always the mid-chord(c) Hit and trial methods(d) Critical Reynolds numberThe question was posed to me in homework.Origin of the question is Turbulent Flow in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct option is (d) Critical Reynolds number

Best EXPLANATION: The critical Reynolds number is usually found from EXPERIMENTS and then using the free-stream flow parameters we can calculate the critical point, which is where the transition from laminar to TURBULENT flow takes place. All the other options are not the right ways to find the transition point.

75.

A given flow can be laminar all over an airfoil in a real case. Select the correct statement for the assumption given above.(a) Always true(b) Always false(c) True only for symmetric airfoils(d) Depends on the Reynolds numberI have been asked this question in final exam.This key question is from Turbulent Flow topic in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right ANSWER is (b) Always false

The best explanation: The boundary layer over any body has both laminar and TURBULENT parts. For an airfoil, the flow encounters laminar skin- friction when it STARTS near the leading edge and there’s a transition to turbulent skin- friction near the trailing edge in REAL cases. In ideal cases, we sometimes assume a fully laminar or turbulent flow depending on the Reynolds number.

76.

For a Reynolds number Rec=10×10^5 and chord length 1m, what is the turbulent boundary layer thickness at the trailing edge (in cm)?(a) 0.0216(b) 0.037(c) 2.16(d) 0.37I got this question in quiz.My enquiry is from Turbulent Flow in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct ANSWER is (c) 2.16

Easy explanation: The turbulent boundary layer thickness is given by δ=\(\frac {0.37x}{\sqrt[5]{Re_c}}\) where in our QUESTION X is the chord length. SOLVING this we get the thickness to be 2.16 CM or 0.0216m. Note that the answer is not 0.0216 since that is wrong in the required units.

77.

Is laminar flow is used to increase the drag and reduce the lift?(a) True(b) FalseI got this question in class test.My query is from Turbulent Flow in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right option is (a) True

Easiest 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.

78.

The correct formula for the Fourier sine series appearing in the solution of thin airfoil theory is_____(a) An=\(\frac {2}{\pi }\int_0^{\pi }\frac {dz}{dx}\) cos⁡n∅ d∅(b) An=\(\frac {1}{\pi }\int_0^{\pi }\frac {dz}{dx}\) cos⁡n∅ d∅(c) An=\(\frac {2}{\pi }\int_0^{2\pi }\frac {dz}{dx}\) cos⁡n∅ d∅(d) An=α-\(\frac {1}{\pi }\int_0^{\pi }\frac {dz}{dx}\) cos⁡n∅ d∅I had been asked this question in my homework.Query is from The Cambered Airfoil in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

The CORRECT answer is (a) An=\(\frac {2}{\pi }\int_0^{\pi }\frac {dz}{DX}\) cos⁡n∅ d∅

The explanation is: 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.

79.

Is flow separation over the top surface of the airfoil at high angle of attack?(a) True(b) FalseThe question was asked in an interview for internship.This intriguing question comes from Modern Low Speed Airfoils in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct answer is (a) True

Easy explanation: The flow separation over the top surface at a HIGH ANGLE of attack, hence yielding higher values of the MAXIMUM lift coefficient. The lift and MOMENT properties are COMPARED with the NAC 2412 airfoil.

80.

NACA 0023 is______(a) Negatively cambered airfoil(b) Positively cambered airfoil(c) Symmetrical airfoil(d) Thin cambered airfoilThis question was posed to me in my homework.The question is from The Cambered Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics

Answer» CORRECT choice is (c) Symmetrical airfoil

To explain: 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.
81.

Is supercritical airfoil has flattened upper surface?(a) True(b) FalseI had been asked this question during an internship interview.The question is from The Cambered Airfoil topic in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right option is (a) True

Easy explanation: SUPERCRITICAL airfoil employ a flattened upper surface, HIGHLY cambered after section and greater leading EDGE radius as compared to traditional airfoil shapes. These CHANGES DELAY the onset of the wave drag.

82.

For a flat plate, aerodynamic center and center of pressure coincide.(a) True(b) FalseI have been asked this question in an interview for job.This interesting question is from The Symmetric Airfoil topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right ANSWER is (a) True

Easiest explanation: The flat PLATE is a THIN, SYMMETRIC airfoil for which moment about quarter-chord is zero. THUS, quarter-chord acts as both the aerodynamic center and center of pressure.

83.

The Kutta condition is not satisfied at the trailing edge where θ=π in transformed coordinates for a symmetrical airfoil.(a) True(b) FalseThe question was asked in a job interview.I would like to ask this question from The Symmetric Airfoil in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct answer is (b) False

To explain I would SAY: DIRECTLY putting θ=π GIVES an indeterminate form (γ(π)=\(\frac {0}{0}\)), but using L’HOSPITAL’s rule in the solution for γ(θ) gives a finite value of zero. Thus, the Kutta CONDITION is satisfied.

84.

Is symmetric airfoil produce positive lift?(a) False(b) TrueThe question was posed to me in a national level competition.The question is from The Symmetric Airfoil in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct option is (b) True

Best EXPLANATION: A symmetrical airfoil only produces positive LIFT, when it is at a positive ANGLE of attack that is leading edge is higher than the TRAILING edge. The air stream splits between top and bottom at the STAGNATION point which is in front of and below the Centre of nose radius.

85.

Is wake behind the airfoil exhibits a continuous vortex?(a) False(b) TrueI got this question in an online interview.I need to ask this question from The Symmetric Airfoil in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct CHOICE is (b) True

Easy explanation: The wake BEHIND the airfoil exhibits a continuous vortex shedding PATTERN below 8 degrees INCIDENCE ANGLE of NACA 0002 and below 7 degrees incidence angle for NACA 0012 at remolds number 1000 it exhibits the continuous vortex.

86.

How lift is calculated?(a) Perpendicular to the direction of motion(b) Parallel to the direction of motion(c) opposite to the direction of motion(d) relative to the direction of motionThe question was asked in semester exam.This is a very interesting question from The Symmetric Airfoil in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct answer is (a) PERPENDICULAR to the direction of motion

The explanation is: 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.

87.

Is fluid flows across the airfoil with a backward facing step?(a) False(b) TrueThe question was asked in final exam.This intriguing question originated from The Symmetric Airfoil topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct option is (b) True

To elaborate: The fluid flows across the AIRFOIL with a backward facing step, a vortex region which has a lower PRESSURE will be CREATED. This will REDUCE the wake region of the symmetric airfoil. The idea of the backward facing step airfoil originates from the idea of dimples on the golf ball where the dimples induce the CREATION of a small vortex region.

88.

Is kelvin established his result subject to weaker conditions?(a) True(b) FalseThe question was asked by my school principal while I was bunking the class.My enquiry is from Kelvin’s Circulation Theorem and the Starting Vortex in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct OPTION is (a) True

The explanation: The condition of incompressibility and CONSTANT DENSITY are not essential to unstable the intense vorticity over the sheet, hence kelvin established his results subject to a weaker condition. To which vortex are FORMED.

89.

The skin-friction drag coefficient for turbulent flow on the bottom surface of an airfoil, with chord length being 1m is _____ (Reynolds number at the trailing edge Rec=32×10^5)(a) 0.074(b) 0.0037(c) 0.027(d) 0.0074I have been asked this question in semester exam.This intriguing question comes from Turbulent Flow in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct choice is (B) 0.0037

To explain: The turbulent skin-friction coefficient, Cf=\(\FRAC {0.074}{\SQRT[5]{Re_c}}\). Putting the given values, we get the answer as 0.01328. This is the coefficient for any one side of the AIRFOIL and is the same for both top and bottom surfaces.

90.

Is dimensionless Reynolds number is an important parameter?(a) False(b) TrueThe question was posed to me in class test.My question comes from Laminar Flow topic in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct choice is (b) True

Easiest explanation: The flow occurring in a fluid in a channel is important in fluid- dynamics problems and SUBSEQUENTLY affects heat and mass TRANSFER in a fluid SYSTEM. The dimensionless Reynolds number is an important PARAMETER in the equations that describe whether FULLY developed flow conditions lead to laminar.

91.

The extra term appearing in the thin airfoil theory solution for a cambered airfoil is a______(a) Full Fourier series(b) Fourier sine series(c) Fourier cosine series(d) ConstantThe question was asked in homework.This intriguing question originated from The Cambered Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct answer is (b) Fourier SINE SERIES

The best I can explain: 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.

92.

Select the unmatched pair.(a) Pressure distribution: lift(b) Shear stress: skin-friction drag(c) Flow separation: form drag(d) Air pressure: pressure dragThis question was posed to me by my college professor while I was bunking the class.This key question is from Viscous Flow in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right OPTION is (d) Air pressure: pressure drag

The explanation: Lift is caused by the pressure distribution on the airfoil. The TWO WAYS drag is exerted is due to shear stress on the surface of the airfoil (SKIN – friction drag) and due to the flow separation (FORM drag, also called pressure drag).

93.

Is positive camber upper surface is more convex?(a) True(b) FalseI had been asked this question in a job interview.This intriguing question comes from The Cambered Airfoil topic in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right choice is (a) True

The explanation is: An airfoil is said to have POSITIVE camber if, as is commonly the case, its upper SURFACE or in the case of a propeller or TURBINE blade its forward surface is more convex. But camber is a complex PROPERTY that can be FULLY characterized by an airfoil camber line.

94.

Select the incorrect statement for a thin, symmetric airfoil out of the following.(a) Quarter-chord is the aerodynamic center(b) Quarter-chord is the center of pressure(c) Moment about quarter-chord depends on the angle of attack(d) Moment about quarter-chord is zeroThis question was addressed to me in an international level competition.Origin of the question is The Symmetric Airfoil topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The correct answer is (c) MOMENT about quarter-chord depends on the angle of attack

To ELABORATE: The coefficient of moment about the quarter chord is zero, THEREBY making it the aerodynamic center (moment coefficient independent of angle of attack) and center of pressure (moment coefficient is zero) for a THIN symmetric airfoil.

95.

How lift and drag ratio can be expressed in a relation?(a) Dividing the lift coefficient by the drag coefficient(b) Dividing the lift coefficient by the moment coefficient(c) Dividing the drag coefficient by the lift coefficient(d) Dividing the drag coefficient by the moment coefficientI had been asked this question by my college director while I was bunking the class.My question is from Airfoil Characteristics topic in chapter Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct option is (a) Dividing the LIFT coefficient by the drag coefficient

The explanation is: 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.

96.

Is flat bottomed wing is more efficient than fully-symmetrical airfoil?(a) True(b) FalseI had been asked this question by my college director while I was bunking the class.The origin of the question is The Symmetric Airfoil topic in section Incompressible Flow over Airfoils of Aerodynamics

Answer» RIGHT answer is (a) True

Explanation: The flat bottomed wing is more EFFICIENT at low SPEEDS than a fully-symmetrical AIRFOIL with an angle of attack LESS, so that the reason most of the flat bottomed, under cambered wing uses less power and drag for the same lift.
97.

For which of the followingKelvin’s theorem is applicable?(a) Flow with Viscous Stresses(b) Compressible Flow(c) Inviscid, Compressible Barotropic Flow(d) Flow with Non-Conservative Body ForcesI had been asked this question in examination.I'd like to ask this question from Kelvin’s Circulation Theorem and the Starting Vortex in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

The CORRECT CHOICE is (c) INVISCID, Compressible Barotropic Flow

The explanation: Kelvin’s Theorem is APPLICABLE for the special CASE of barotropic flow while dealing with inviscid, compressible flows.

98.

What is the thickness in NACA 2412 airfoil?(a) 12%(b) 24%(c) 41%(d) 2%I had been asked this question in exam.The origin of the question is Airfoil Nomenclature topic in section Incompressible Flow over Airfoils of Aerodynamics

Answer»

Right choice is (a) 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.

99.

Who designed the nomenclature of airfoil?(a) NACA series(b) EULAR series(c) EPPLER series(d) CLARK seriesI had been asked this question by my college professor while I was bunking the class.Asked question is from Airfoil Nomenclature in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct option is (a) NACA series

To ELABORATE: 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.

100.

The skin-friction coefficient for turbulent flow Cf is proportional to Rec. The proportionality constant is_____(a) 0.074(b) 0.64(c) 0.33(d) 1.328I had been asked this question in an international level competition.The question is from Turbulent Flow topic in division Incompressible Flow over Airfoils of Aerodynamics

Answer»

Correct choice is (a) 0.074

The best explanation: The coefficient of skin-friction drag for TURBULENT flow is GIVEN as Cf=\(\FRAC {0.074}{\sqrt[5]{Re_c}}\) where Rec is the REYNOLDS number at the trailing edge. The required CONSTANT is 0.074. For laminar flows, the constant of proportionality is different.