

InterviewSolution
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.
101. |
Is viscosity effects the pitching moments?(a) True(b) FalseThe question was posed to me during an online exam.I'm obligated to ask this question of Viscous Flow in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct choice is (a) True |
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102. |
Is lift loss with increase in angle of attack?(a) True(b) FalseThis question was addressed to me in class test.I want to ask this question from Viscous Flow in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right choice is (a) True |
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103. |
Is initial distribution is given by solid curves?(a) True(b) FalseThis question was addressed to me in unit test.Question is from Modern Low Speed Airfoils topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The CORRECT choice is (a) True |
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104. |
For a cambered airfoil, the center of pressure does not vary with_____(a) Chord length(b) Angle of attack(c) The shape of the airfoil(d) Location of the aerodynamic centerThe question was asked in an interview.I'd like to ask this question from The Cambered Airfoil topic in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct option is (d) LOCATION of the AERODYNAMIC center |
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105. |
What is the leading edge radius of LS-0417 airfoil?(a) 0.07c(b) 0.05c(c) 0.08c(d) 0.09cThe question was posed to me during an interview for a job.Asked question is from Modern Low Speed Airfoils topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct choice is (c) 0.08c |
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106. |
The fundamental equation for the thin airfoil theory does not encompass which of the following approximations?(a) The angle of attack and slope of the camber line is small(b) Camber line induced velocity distribution is the same for chord line(c) Vortex sheet is placed at the chord line(d) Kutta condition is satisfied at the trailing edgeThis question was posed to me during an internship interview.The above asked question is from The Symmetric Airfoil topic in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right option is (d) Kutta CONDITION is satisfied at the trailing edge |
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107. |
A vortex sheet placed along the chord line gives the best representation for______(a) Thick Airfoil(b) NACA Airfoils(c) Negatively Cambered Airfoil(d) Thin AirfoilI got this question during a job interview.I need to ask this question from The Symmetric Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» The CORRECT option is (d) Thin Airfoil |
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108. |
In reality, the starting vortex dies out. Why?(a) Lift becomes zero(b) At later times, Kelvin’s theorem is not applicable(c) Due to Viscosity(d) This assumption is wrong. Starting vortex never diesI got this question in examination.This interesting question is from Kelvin’s Circulation Theorem and the Starting Vortex topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct choice is (c) DUE to Viscosity |
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109. |
Is kutta condition apply to oval shaped body?(a) True(b) FalseI had been asked this question in homework.My question is from The Kutta Condition topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct ANSWER is (a) True |
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110. |
Is airfoil can be characterized by the relation between the angle of attack, lift coefficient and drag coefficient?(a) True(b) FalseThe question was asked in final exam.My question is taken from Airfoil Characteristics topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct option is (a) True |
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111. |
Is lift is created on the airfoil using kutta condition?(a) True(b) FalseI got this question during an internship interview.This question is from The Kutta Condition in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right choice is (a) True |
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112. |
Define drag curve.(a) Drag will increase rapidly at particular degree of angle of attack and overcomes the lift curve at particular degree of angle of attack(b) Drag curve will decrease at particular degree of angle of attack and lift curve increasing at particular degree of angle of attack(c) Drag curve remains constant and lift curve will be increasing(d) Lift curve remains constant and drag curve will be increasingI have been asked this question by my college director while I was bunking the class.The doubt is from Airfoil Characteristics in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct answer is (a) Drag will increase RAPIDLY at PARTICULAR DEGREE of angle of attack and overcomes the lift curve at particular degree of angle of attack |
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113. |
What is the purpose of trailing edge?(a) Airflow rejoins(b) Airflow separated(c) Vortex are created(d) Stalling will createdThe question was asked in homework.Query is from Airfoil Nomenclature topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right choice is (a) Airflow rejoins |
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114. |
We can get the skin-friction drag coefficient for a laminar flow for a flat plate by using x as chord length in local skin- friction drag coefficient calculation.(a) True(b) FalseI had been asked this question in examination.Enquiry is from Laminar Flow in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct choice is (b) False |
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115. |
Is thin viscous region forms over the airfoil at low angle of attack?(a) True(b) FalseThis question was posed to me in final exam.I would like to ask this question from Viscous Flow in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right option is (a) True |
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116. |
Is high adverse pressure gradient that occurs at a sharp leading edge?(a) True(b) FalseI got this question during an online exam.My question is from Viscous Flow topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct answer is (a) True |
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117. |
The coefficient of lift for a thin, cambered airfoil with A0=0.2 and A1=1.12 is____(a) π1.52(b) π1(c) π0.52(d) π2The question was posed to me in my homework.My question comes from The Cambered Airfoil topic in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct CHOICE is (a) π1.52 |
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118. |
For α=5°, A0=1 and A1=-2 total circulation Γ for a thin cambered airfoil equals______(a) 0(b) 2πcV∞(c) πcV∞(d) 2.5πcV∞The question was posed to me during an interview for a job.Question is from The Cambered Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right choice is (a) 0 |
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119. |
For NACA4313 what is the maximum camber and the position of maximum camber from the leading edge respectively is______(a) 0.04c, 0.4c(b) 0.4c, 0.03c(c) 0.13c, 0.4c(d) 0.04c, 0.03cI had been asked this question in an internship interview.Origin of the question is The Cambered Airfoil topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct choice is (a) 0.04c, 0.4c |
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120. |
Is upper camber refers to the camber of the upper surface?(a) True(b) FalseThe question was asked in examination.Asked question is from The Cambered Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» The CORRECT option is (a) True |
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121. |
What is the Kutta Condition in terms of strength (γ) of the thin airfoil?(a) γ = 0(b) γ(LE) = 0(c) \(\frac {D\gamma }{Dt}\)=0(d) γ(TE) = 0I have been asked this question in an international level competition.Query is from The Symmetric Airfoil in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right answer is (d) γ(TE) = 0 |
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122. |
Is high airspeed around the trailing edge causes strong viscous forces?(a) False(b) TrueI got this question in examination.The doubt is from The Kutta Condition in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» RIGHT OPTION is (a) False To explain: The high airspeed AROUND the trailing edge causes strong viscous forces to act on the air adjacent to the trailing edge of the airfoil and the result is that a strong vortex accumulates on the TOPSIDE of the airfoil, NEAR the trailing edge. |
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123. |
The value of integral ∫\(_0^?\frac {cosn\theta d\theta }{cos\theta-cos∅}\)=π\(\frac {sin n∅}{sin ∅}\) is valid for the limit 0 to_____(a) π(b) 2π(c) –π(d) nπThis question was addressed to me in examination.This key question is from The Symmetric Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» RIGHT option is (a) π The best I can explain: 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. |
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124. |
Purpose of vortex method is ____________(a) To reduce the dimensionality of aerodynamic model(b) To increase the dimensionality of aerodynamic model(c) To reduce the lift of aerodynamic model(d) To increase the lift of aerodynamic modelThe question was asked during an interview.My question comes from The Vortex Sheet topic in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct option is (a) To reduce the dimensionality of aerodynamic model |
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125. |
How propeller creates a thrust force?(a) With transmitted power(b) With own power(c) With the existing power(d) With supplied powerThis question was addressed to me in homework.My question is from Airfoil Characteristics in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct choice is (d) With supplied POWER |
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126. |
What is responsible for creating the aerodynamic drag on the airfoil?(a) Pressure distribution(b) Friction(c) Thrust(d) GravityThis question was posed to me during a job interview.I need to ask this question from Viscous Flow in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct CHOICE is (B) Friction |
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127. |
The quarter-chord moment coefficient for a cambered airfoil depends on______(a) Chord length(b) Angle of attack(c) The shape of the airfoil(d) Center of pressureI have been asked this question by my college professor while I was bunking the class.My doubt is from The Cambered Airfoil topic in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» RIGHT choice is (c) The shape of the airfoil Explanation: The coefficient of moment about the quarter-chord is independent of angle of ATTACK and chord length. It depends on the shape of the cambered airfoil (slope of the camber LINE (\(\FRAC {dz}{dx}\))). |
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128. |
The lift curve slope for a thin, cambered airfoil is 2π.(a) True(b) FalseI got this question during an interview.This intriguing question originated from The Cambered Airfoil topic in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right choice is (a) True |
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129. |
Modern low-speed airfoils were developed using numerical methods using the computer directly. The given statement is_____(a) Partially true(b) False(c) True(d) IncompleteThe question was posed to me during an interview.The doubt is from Modern Low Speed Airfoils topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right option is (a) PARTIALLY true |
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130. |
For γ(θ)=2V∞(A0\(\frac {1+cos\theta }{sin\theta }\) + Σ\(_{n=1}^∞\)sin nθ An) select the statement which is invalid.(a) The solution is valid only for cambered airfoils(b) The solution is valid for all thin airfoils(c) A0 Is the n=0^th term for the Fourier series(d) Kutta condition is satisfied at the trailing edge i.e. θ=π.This question was posed to me in quiz.My doubt stems from The Cambered Airfoil topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct choice is (a) The solution is valid only for cambered AIRFOILS |
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131. |
The equation \(\frac {1}{2\pi } \int_0^c \frac {\gamma(\xi)d\xi}{x-\xi}\)=V∞α is called the fundamental equation of thin airfoil theory for______(a) Cambered airfoils only(b) Symmetric airfoils only(c) All thin airfoils(d) Symmetric and positively cambered airfoilsI got this question by my college director while I was bunking the class.I need to ask this question from The Cambered Airfoil in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct CHOICE is (b) Symmetric airfoils only |
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132. |
What is the total circulation around the symmetric airfoil according to the thin airfoil theory?(a) Γ=πα^2cV∞(b) Γ=π^2αcV∞(c) Γ=2παcV∞(d) Γ=παcV∞I had been asked this question at a job interview.Question is from The Symmetric Airfoil in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct answer is (d) Γ=παcV∞ |
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133. |
Find the induced velocity in the direction normal to the camber line a point on where the slope is 0.087. Given the angle of attack is 5° and free-stream velocity is 20 units. Use thin airfoil approximation.(a) 0(b) 101.74 units(c) 98.26 units(d) 3.48 unitsThis question was posed to me by my school principal while I was bunking the class.The query is from The Symmetric Airfoil topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct answer is (a) 0 |
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134. |
Is inviscid incompressible fluid has constant density?(a) False(b) TrueThis question was addressed to me in an online interview.The query is from Kelvin’s Circulation Theorem and the Starting Vortex topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct option is (a) False |
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135. |
At what point circulation between two material points in the sheet remains conserved.(a) Zero(b) One(c) Greater than zero(d) Less than zeroI have been asked this question in a job interview.I'm obligated to ask this question of The Vortex Sheet topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» CORRECT option is (a) 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. |
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136. |
Is airfoil shape similar to blades of the propeller?(a) True(b) FalseI had been asked this question during an interview.This is a very interesting question from Airfoil Characteristics topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct CHOICE is (a) True |
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137. |
The boundary layer thickness for an incompressible, laminar flow at a distance x with Reynolds number Rex is δ. Which is____(a) Directly proportional to √x(b) Inversely proportional to x(c) Directly proportional to Rex(d) Inversely proportional to Rex^2The question was asked in an online quiz.My enquiry is from Laminar Flow topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right option is (a) Directly PROPORTIONAL to √x |
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138. |
Is laminar flow used in the smooth flow of a viscous liquid through a tube?(a) True(b) FalseI had been asked this question in unit test.This key question is from Laminar Flow in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct CHOICE is (a) True |
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139. |
The major breakthrough in high- speed airfoil industry was_____(a) GA (W)-1 airfoil(b) Supercritical airfoil(c) Standard NACA airfoils(d) Symmetrical airfoilsThe question was posed to me in a job interview.This interesting question is from Modern Low Speed Airfoils in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct CHOICE is (b) Supercritical airfoil |
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140. |
Given an angle of attack 5°, A1=1.5, A2=2 and the chord length is 5m, the moment coefficient about the quarter-chord is_____(a) \(\frac {\pi }{8}\)(b) \(\frac {\pi }{4}\)(c) \(\frac {-\pi }{8}\)(d) –\(\frac {\pi }{4}\)The question was asked during an internship interview.This interesting question is from The Cambered Airfoil topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» Correct option is (a) \(\frac {\pi }{8}\) |
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141. |
Which of the following is the correct solution of the transformed fundamental equation of aerodynamics for a symmetrical airfoil?(a) γ(θ)=2αV∞\(\frac {sin\theta }{1+cos\theta }\)(b) γ(θ)=2αV∞\(\frac {1+cos\theta }{sin\theta }\)(c) γ(θ)=2αV∞\(\frac {1-cos\theta }{sin\theta }\)(d) γ(θ)=2αV∞\(\frac {cos\theta }{sin\theta }\)The question was posed to me during an internship interview.This is a very interesting question from The Symmetric Airfoil topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» RIGHT choice is (b) γ(θ)=2αV∞\(\frac {1+cos\theta }{sin\theta }\) The explanation is: The solution of the fundamental equation of THIN airfoil THEORY is obtained using the transformation of coordinates. We have α and V∞ and using the standard integrals we can find a solution for γ(x) as γ(θ)=2αV∞\(\frac {1+cos\theta }{sin\theta }\) where 0≤θ≤π for 0≤x≤C. |
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142. |
For an arbitrary inviscid and incompressible flow, with all the body forces zero, what is best described by the given sketch?(a) Boundary Layer Formation(b) Kelvin’s Circulation Theorem(c) Kutta Condition(d) Generation of LiftI have been asked this question during a job interview.Query is from Kelvin’s Circulation Theorem and the Starting Vortex in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right answer is (B) Kelvin’s Circulation Theorem |
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143. |
Is vortex sheet is unstable at high Reynolds number?(a) True(b) FalseI got this question in an online quiz.The above asked question is from The Vortex Sheet in portion Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct CHOICE is (a) True |
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144. |
For a real flow over an airfoil, with Reynolds number in the laminar region as it starts, select the incorrect choice.(a) Flow at trailing edge is laminar(b) Flow is both turbulent and laminar at different places(c) There is a transition region between leading and trailing edge(d) Flow at trailing edge is turbulentThe question was asked in exam.This key question is from Turbulent Flow topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct option is (a) Flow at trailing edge is LAMINAR |
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145. |
When the free-stream velocity is made one-third, the transition point moves downstream.(a) The given statement is true or false?(b) True(c) FalseThis question was posed to me by my college director while I was bunking the class.Question is taken from Turbulent Flow in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» Right answer is (a) The given statement is true or FALSE? |
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146. |
Is outer flow over an airfoil has less camber than the original airfoil?(a) False(b) TrueThe question was asked in an international level competition.This question is from Viscous Flow in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» CORRECT option is (b) True Explanation: The outer flow SEES an equivalent airfoil that has less camber than the original airfoil, due to the disproportionate growth of the boundary LAYER on the two side and that has an open trailing EDGE. Such an airfoil produce less lift than the original airfoil. |
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147. |
For finding the skin- friction drag we need to only measure shear-stress at the top or bottom surface.(a) Always true(b) Always false(c) True for flat plate, which is a symmetric airfoil(d) Depends on the Reynolds numberThis question was addressed to me in an internship interview.Origin of the question is Laminar Flow topic in chapter Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct answer is (c) True for flat plate, which is a SYMMETRIC AIRFOIL |
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148. |
Is 17% is the thickness of NASA LS-0417 airfoil?(a) True(b) FalseThis question was posed to me in an online quiz.My question is based upon Modern Low Speed Airfoils in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct OPTION is (a) True |
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149. |
The lift per unit span for a thin, cambered airfoil with Γ=10\(\frac {m^2}{s}\), ρ∞=1.0255\(\frac {kg}{m^3}\), V∞=10\(\frac {m}{s}\) is____(a) 0(b) 102.55\(\frac {N}{m}\)(c) 102.55N(d) 55\(\frac {N}{m}\)This question was posed to me during an online exam.My query is from The Cambered Airfoil topic in section Incompressible Flow over Airfoils of Aerodynamics |
Answer» The correct OPTION is (a) 0 |
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150. |
Is camber is the asymmetry between the two acting surfaces of an airfoil?(a) True(b) FalseI got this question during an interview for a job.My query is from The Cambered Airfoil in division Incompressible Flow over Airfoils of Aerodynamics |
Answer» CORRECT option is (a) True Easiest explanation: Camber is the asymmetry between the two acting SURFACES of an airfoil, with the top SURFACE of a wing or corresponding the front surface of a propeller BLADE, COMMONLY being more convex, positive camber. |
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