74 + Interview Questions in Airfoil And Geometry Selection in Aircraft Design Page 1 InterviewSolution

1.	A cambered airfoil is operating with 200N of lift. The velocity of the flow is 500m/s. Determine lift-induced drag by this camber airfoil.(a) 0N(b) 0(c) 0.5(d) 0.5This question was addressed to me during an online interview.Question is taken from Airfoil Selection-3 in division Airfoil and Geometry Selection of Aircraft Design
Answer» The correct answer is (a) 0N The best explanation: Given, Lift=200N, VELOCITY = 500m/s. For cambered AIRFOIL the DOWNWASH ANGLE = 0°. Hence, there will be no downwash ACTING on the flow. Hence, there will be no induced drag. Hence, for any airfoil induced drag = 0N.

Discussion

2.	Let’s consider we need to design supersonic aircraft then, which type of airfoil should be adopted?(a) Supersonic airfoil(b) Symmetric airfoil(c) Bluff shape(d) Thicker airfoilI got this question by my school teacher while I was bunking the class.The doubt is from Airfoil Selection-3 topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» The correct answer is (a) SUPERSONIC airfoil For explanation: Supersonic aircraft needs to FLY at much higher speed. Hence, supersonic airfoils are used to provide required velocity. Thicker airfoils cannot generate required amount of FLOW characteristics such as CRITICAL mach number.

Discussion

3.	If taper ratio is T and span is b with surface area as S then the root chord is _____(a) Croot = 2S / (b(1+T))(b) Croot = 2S / (b(3+T))(c) Croot = 2S / (2b(1+T))(d) Croot = 2*S / (1+T)This question was addressed to me during an internship interview.Asked question is from Wing Geometry-1 in section Airfoil and Geometry Selection of Aircraft Design
Answer» Correct option is (a) Croot = 2S / (B(1+T)) The best explanation: Taper ratio represents how much WING is tapered. For any given area S, span b and taper ratio T, root CHORD is given by, Croot = 2S / (b(1+T)). The value of root chord will be different for different aircrafts.

Discussion

4.	A typical wing has aspect ratio of 7 and span of 2m. If taper ratio T = 0.2 then, find the value of root chord.(a) 0.4761m(b) 0.61m(c) 0.81m(d) 0.71cmThis question was posed to me in homework.Question is taken from Wing Geometry-1 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Correct ANSWER is (a) 0.4761m Explanation: Given, SPAN = 2m, aspect ratio = 7 and taper ratio T = 0.2 The root chord is given by, Croot = 2S / (b(1+T)) Here, S is not given but aspect ratio is given. Aspect ratio AR = Span square / S = 22 / S 7 = 4/ S S = 0.5714 m^2 Hence, Croot = 2S / (b(1+T)) = 20.5714 / (2*(1+0.2)) = 0.4761m.

Discussion

5.	Which type of wing planform is shown below?(a) Tapered wing(b) Elliptic wing(c) Rectangular wing(d) Delta wingThis question was posed to me in an internship interview.My doubt is from Wing Geometry-1 topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» Correct option is (a) Tapered wing For explanation I would say: Tapered wings have finite value of taper RATIO. In taper wing root and tip chord will have different values of chord. ELLIPTIC wings are of elliptic shape. RECTANGULAR wing has taper ratio of unity which shows a rectangular profile.

Discussion

6.	A thin airfoil is defined as NACA 0009 is at ɑ = 5°. What is the value of lift curve slope?(a) 2ᴨ per rad(b) 2ᴨ per degree(c) 4 per degree(d) 0.4 per degreeThis question was posed to me during an interview for a job.My query is from Airfoil Selection-2 topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» Right answer is (a) 2ᴨ per rad Best explanation: Given, NACA0009 airfoil, AOA ɑ = 5° As mentioned the airfoil is thin. HENCE, thin airfoil THEORY can be used. According to thin airfoil theory, for thin airfoil, LIFT CURVE slope = 2π 1/rad = 0.11 1/degree.

Discussion

7.	For high aspect ratio strength of the wing tip vortex is low.(a) True(b) FalseThis question was addressed to me during an online exam.My doubt stems from Wing Geometry-2 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Correct CHOICE is (a) True Easiest explanation: Wing tip of high aspect ratio wing is far from root as compared to low aspect ratio wing. The wing affected by HIGHER ratio will be less and the STRENGTH of tip vortex will be less as WELL.

Discussion

8.	Wing sweep is used to ____(a) increase critical mach number(b) decrease critical mach number(c) increase lofting(d) increase draftingThe question was asked during an interview.This is a very interesting question from Wing Geometry-2 topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» Correct choice is (a) increase critical mach number The best EXPLANATION: If wing is at some finite angle from fuselage REFERENCE line then it is called wing is sweep by that much degree. Main function of sweep is to increase critical mach number. LOFTING is SKIN modelling and drafting is drawing phenomena.

Discussion

9.	Canard pusher aircraft uses wing sweep to change location of aerodynamic centre.(a) True(b) FalseThe question was asked in an interview.This key question is from Wing Geometry-2 topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» CORRECT answer is (a) True Easiest EXPLANATION: CANARD pusher aircrafts are heavier at rearward part. At tail weight is more. By sweeping the wing we can alter the LOCATION of aerodynamic center for balance.

Discussion

10.	An aerofoil is operating with flow velocity of 100m/s. Aerofoil is symmetric and is at 0° AOA. Determine lift produced by this aerofoil.(a) 10.5 N(b) 0 N(c) 50 N(d) 100NI have been asked this question during an interview.The question is from Airfoil Selection-1 topic in portion Airfoil and Geometry Selection of Aircraft Design
Answer» Right answer is (b) 0 N The BEST I can explain: LIFT is an upward FORCE PRODUCED by pressure difference. Given, velocity V=100m/s, AOA = 0°. For symmetric airfoil at 0° AOA pressure will be same as upper and lower surface. Hence, lift = 0N.

Discussion

11.	Find location of aerodynamic centre from leading edge if, chord is 10m and airfoil has flow velocity of 0.2 M.(a) 2.5 m(b) 0.2 m(c) 5 m(d) 10 mThe question was posed to me during an online exam.My question comes from Airfoil Selection-1 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» The correct CHOICE is (a) 2.5 m Best explanation: Given, VELOCITY = 0.2M = 0.2340 = 68m/s CHORD = 10m Given velocity is less than the sonic speed. Hence, location of aerodynamic center is = 25% of chord = 0.2510 = 2.5m from leading edge.

Discussion

12.	Pitching moment co-efficient for an airfoil is given by _____(a) liftarea(b) dragarea(c) section pitching moment/dynamic pitching moment(d) lift/dynamic momentI got this question during an online exam.My question is from Airfoil Selection-1 topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» Correct option is (c) section pitching moment/DYNAMIC pitching moment The BEST explanation: LIFTarea will give moment. DRAGarea will also provide moment but will be less. The ratio of SECTIONAL pitching moment and dynamic pitching moment is called pitching moment co-efficient.

Discussion

13.	If I want to locate my minimum pressure at 0.5 of chord c then _____________(a) NACA 6512(b) NACA 652132(c) NACA 65211(d) NACA 0009This question was posed to me in my homework.Question is taken from Airfoil Selection-3 in section Airfoil and Geometry Selection of Aircraft Design
Answer» Correct answer is (B) NACA 652132 For explanation I would say: NACA series is used to identify AIRFOILS as per our requirement. NACA -4 and -5 digit series are not used to provide a MINIMUM pressure point. NACA 6 digit series is used for such operations.

Discussion

14.	A twin turboprop engine is flying with 120m/s. Wing of the aircraft has span of 2m and lift co-efficient as 0.15 with aspect ratio as 9.2. Find the lift produced by the wing.(a) 574.18 N(b) 574.18(c) 120 N(d) 620 NThe question was posed to me in my homework.My query is from Wing Geometry-1 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» The correct OPTION is (a) 574.18 N For EXPLANATION: Given, a twin turboprop Velocity V = 120m/s, CL = 0.15, Aspect ratio = 9.2, span b=2m Aspect ratio = bb/S 9.2 = 22/s S = 0.434m^2 Now, LIFT is given by, L = (1/2)ρV^2SCL = 0.5* 1.2251201200.4340.15 = 574.18N.

Discussion

15.	Typically tail provides ______(a) stability(b) control(c) stability and control(d) accelerating forceI had been asked this question during an interview.Enquiry is from Tail Geometry and Arrangement in section Airfoil and Geometry Selection of Aircraft Design
Answer» Right choice is (c) stability and control The best I can explain: TYPICALLY tail is used to provide stability and control. Accelerating FORCE will be provided from engines. One of the main function of the tail is to provide enough stability and controllability to make aircraft fly SMOOTHER at the TIME of any DISTURBANCE.

Discussion

16.	Which tail profile is shown in below diagram?(a) Cruciform tail(b) T – tail(c) H – tail(d) Conventional tailThe question was posed to me in an international level competition.I'd like to ask this question from Tail Geometry and Arrangement topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» The correct answer is (a) Cruciform tail To EXPLAIN I would SAY: Cruciform tail viewed from front will look LIKE a ‘plus’ as shown in figure. T – tail will look like letter T and SIMILARLY H tail has SHAPE of letter H.

Discussion

17.	What is the use of V type tail?(a) To reduce wetted area(b) To increase wetted area(c) To prevent adverse yaw roll(d) To produce thrustThis question was posed to me during an interview for a job.This interesting question is from Tail Geometry and Arrangement topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» Correct choice is (a) To reduce wetted area Easiest explanation: V tail arrangement is used to combine HORIZONTAL and vertical tail. It combines in such WAY that the wetted area reduces. However, it generates adverse yaw roll MOTION. To prevent such EFFECT INVERTED V tail can be used.

Discussion

18.	If root chord is 2m and tip chord is 0.9m then, find taper ratio.(a) 0.45(b) 0.55(c) 0.65(d) 0.25I have been asked this question in an online quiz.My query is from Wing Geometry-2 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» RIGHT answer is (a) 0.45 Easiest EXPLANATION: TAPER RATIO is DEFINED as, Taper ratio = tip chord/root chord = 0.9/2 = 0.45.

Discussion

19.	An airfoil with symmetric profile has pitching moment about aerodynamic centre as 13 Nm at 4° AOA. Determine Pitching moment at same point if AOA is increased by 8°.(a) 13Nm(b) 8Nm(c) 5Nm(d) 13NcmThe question was posed to me by my college director while I was bunking the class.The query is from Airfoil Selection-2 topic in portion Airfoil and Geometry Selection of Aircraft Design
Answer» Right option is (a) 13Nm To elaborate: Given, a symmetric airfoil Pitching moment PM1 = 13Nm at AOA = 4°. Now AOA is INCREASED by 8°. HENCE new AOA = 4+8 = 12°. Now PM1 is acting through aerodynamic centre. PM will be independent of AOA at aerodynamic centre. Hence, at 12° pitching moment will not change. Hence, at same POINT at 12° PM2 = PM1 = 13Nm.

Discussion

20.	The straight line joining L.E. and T.E. of an airfoil is _______(a) chord line(b) camber(c) camber line(d) radiiThe question was posed to me in final exam.This question is from Airfoil Selection-2 in portion Airfoil and Geometry Selection of Aircraft Design
Answer» RIGHT answer is (a) CHORD line To elaborate: Chord line is defined as the straight line which connects the leading and TRAILING edge of an airfoil. Camber line or MEAN camber line is line passing from locus of mid points. Almost every parameter is defined with RESPECT to chord.

Discussion

21.	What do you mean by double cambered airfoil?(a) Airfoil with flat lower surface(b) Airfoil with flat upper and lower surface(c) Airfoil with curved upper and lower surface(d) Only upper surface is curvedThe question was posed to me during an online exam.I need to ask this question from Airfoil Selection-3 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» The correct option is (C) Airfoil with curved upper and lower surface For explanation: Traditionally, airfoils with only curved upper surface and flat bottom was termed as camber airfoil. However, if both surfaces are curved then, it is CALLED a double cambered airfoil. In now days not most PEOPLE use this terms but at some PLACES it is STILL in use.

Discussion

22.	If aspect ratio of wing is 8 and S=0.1m^2 then, what will be the span of wing?(a) 0.89m(b) 0.89cm(c) 0.89(d) 0.89inchI have been asked this question by my school principal while I was bunking the class.The doubt is from Wing Geometry-2 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Right option is (a) 0.89m Easiest EXPLANATION: Given, WING Aspect ratio AR=8, S=0.1m^2 Aspect ratio = span SQUARE/ reference area 8 = span square / 0.1 Span square = 8*0.1 = 0.8 Hence, span = 0.89m.

Discussion

23.	What will be the location of maximum camber for NACA 13250 with chord length of 2m?(a) 0.32m from leading edge(b) 0.32 from trailing edge(c) 0.32m from trailing edge(d) 0.32 from leading edgeThis question was addressed to me in final exam.Enquiry is from Airfoil Selection-3 in section Airfoil and Geometry Selection of Aircraft Design
Answer» Correct OPTION is (a) 0.32m from leading edge For explanation: GIVEN, NACA 13250 airfoil Based on NACA -5 digit series, LOCATION of maximum CAMBER = (second and THIRD digit)/2 in hundredth of chord from leading edge = (32/2) * 0.01 * chord = (32/2)0.012 = 0.32m from leading edge.

Discussion

24.	Excessive dihedral can produce _____(a) dutch roll(b) thrust(c) pitch(d) deflection at noseThe question was posed to me during an interview.The query is from Wing Geometry-2 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» The correct CHOICE is (a) dutch roll Best explanation: Dihedral is upward deflection of wing. It is used to provide roll stability. If excessive dihedral is PROVIDED then, it will LEAD to dutch roll; continuous side to side MOTION involving YAW and roll.

Discussion

25.	Which of the following is correct?(a) Tail sweep is generally more than wing sweep(b) Tail sweep is generally same as wing sweep(c) Tail sweep is always ten times of wing sweep(d) Tail sweep is generally less than wing sweepThe question was posed to me by my school principal while I was bunking the class.Enquiry is from Tail Geometry and Arrangement in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Right choice is (a) TAIL sweep is generally more than wing sweep For explanation I WOULD say: Tail sweep is the sweep angle provided to tail. In GENERAL, tail sweep is more than wing sweep. Higher sweep at tail than wing sweep will increase critical mach NUMBER of the tail. Hence, it will delay SHOCK formation at tail.

Discussion

26.	Glider with span of 2m is flying at 10000ft. If aspect ratio of tail is 10 then, determine reference area of the tail.(a) 0.4m^2(b) 0.4(c) 0.8(d) 0.8m^2The question was posed to me during an internship interview.My doubt is from Tail Geometry and Arrangement topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» RIGHT option is (a) 0.4m^2 The explanation is: Aspect RATIO of TAIL = SPAN square / AREA of tail 10 = 2*2 / area Area = 4/10 = 0.4m^2.

Discussion

27.	Which of these is not part of an airfoil?(a) Leading edge(b) Chord(c) Camber(d) FinsThe question was posed to me by my college professor while I was bunking the class.The doubt is from Airfoil Selection-1 topic in portion Airfoil and Geometry Selection of Aircraft Design
Answer» The CORRECT answer is (a) Leading edge Easiest explanation: Fins are extended SURFACES which are used to IMPROVE heat TRANSFER. Leading edge is foremost PART of an airfoil. Camber is curve of an airfoil.

Discussion

28.	From following diagram body-1 is not streamline?(a) True(b) FalseI got this question in a job interview.I would like to ask this question from Airfoil Selection-2 in portion Airfoil and Geometry Selection of Aircraft Design
Answer» The correct choice is (a) True The explanation is: Body-1 is TYPICAL AIRFOIL shape. An airfoil is STREAMLINE body. Body-2 is not a streamline body as compare to body-1. Streamline body will provide streamline FLOW.

Discussion

29.	Determine the value for maximum camber for NACA1235 airfoil. Given chord is 1m.(a) 0.01m(b) 0.02m(c) 0.021(d) 0.012The question was posed to me in examination.My question comes from Airfoil Selection-2 topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» The correct option is (a) 0.01m The explanation: GIVEN, NACA 4 digit AIRFOIL as NACA 1235. CHORD c=1m. Based on NACA 4 digit, FIRST digit is the maximum camber in hundreds of chord. Hence, maximum camber for NACA 1235 = 0.01chord = 0.011 = 0.01m.

Discussion

30.	An airfoil is such that the drag produced by airfoil is 25N. Airfoil has chord of 2m and is experiencing dynamic pressure of 10Pa. Find the value of drag co-efficient for airfoil.(a) 1.25(b) 1.6(c) 2.5(d) 2.51I have been asked this question in a job interview.Asked question is from Airfoil Selection-3 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Right choice is (a) 1.25 The best I can EXPLAIN: Given, Drag = 25N, DYNAMIC pressure q = 10Pa, chord C = 2m. Now, drag co-efficient Cd = Drag / (qc) = 25 / (102) = 1.25.

Discussion

31.	A NACA 22104 airfoil is operating at AOA = 3°. The lift-coefficient at this AOA is 0.15. Determine the value of angle of attack at designed lift co-efficient.(a) 4.36°(b) 4.36rad(c) 0.05 rad(d) 0.05°I have been asked this question in an interview for job.Enquiry is from Airfoil Selection-3 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» The correct choice is (a) 4.36° For explanation: Given NACA 22104 airfoil. AOA = 3°, Cl = 0.15 at 3° Based on NACA -5 digit SERIES DESIGN Cl = FIRST digit0.15 = 20.15 = 0.3 Now, for thin airfoil LIFT curve slope = 0.11 per degree 0.11 = (design Cl – Cl at 3°) / (AOA –3°) 0.11 = (0.3-0.15) / (AOA -3°) 0.11*(AOA -3°) = 0.15 Hence, at designed Cl, AOA = (0.15/0.11) + 3° = 4.36°.

Discussion

32.	Wing planform is affected by _____(a) sweep only(b) aspect ratio, sweep angle, taper ratio(c) aspect ratio only(d) taper ratio onlyI had been asked this question by my college director while I was bunking the class.My question comes from Wing Geometry-1 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Correct option is (B) aspect ratio, sweep angle, taper ratio Easy explanation: WING PLANFORM is shape of the wing when viewed from TOP. Aspect ratio will affect SPAN and area. Overall planform will be affected by aspect ratio, taper ratio, sweep etc.

Discussion

33.	Which of the following is not part of tail?(a) Aileron(b) Rudder(c) Elevator(d) Horizontal stabilizerI had been asked this question during an online interview.I need to ask this question from Tail Geometry and Arrangement in portion Airfoil and Geometry Selection of Aircraft Design
Answer» The correct option is (a) Aileron The BEST I can explain: Aileron is part of the wing. Rudder is part of vertical tail. ELEVATOR is part of the HORIZONTAL tail. Horizontal tail is also called Horizontal STABILIZER.

Discussion

34.	Airfoil is a streamline shape.(a) True(b) FalseThis question was addressed to me in my homework.This is a very interesting question from Airfoil Selection-1 topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» Correct choice is (a) True The best I can explain: Streamline SHAPE is that shape which GIVES streamlined FLOW. An airfoil has that characteristic. An airfoil is able to provide much smoother flow than non-streamlined BODY.

Discussion

35.	Which of the following is not an application of an airfoil?(a) To provide thrust for acceleration(b) To provide lift(c) To provide lift co-efficient(d) To provide streamlined flowI had been asked this question in quiz.This interesting question is from Airfoil Selection-2 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» The CORRECT choice is (a) To provide THRUST for acceleration To EXPLAIN: In an aircraft, engines are used to provide thrust for acceleration. Airfoil is used to provide lift, to improve certain flow characteristics. Streamlined flow is ALSO produced by an airfoil.

Discussion

36.	A symmetric airfoil is operating with flow velocity of 350m/s. The lift produced by the airfoil is 21 N at 0.008 rad AOA. If chord is c m then, what will be the pressure difference across the airfoil?(a) 21/c Pa(b) 21c/2 N(c) 21 Pa(d) 21c NThis question was addressed to me at a job interview.I'd like to ask this question from Airfoil Selection-1 topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» Right option is (a) 21/C Pa The best I can explain: Given, symmetric airfoil Flow VELOCITY V=350m/s, Lift L=21N, AOA = 0.008rad Chord = c m, Span of airfoil = 1 unit Pressure difference = Lift/Area of airfoil = 21/c*1 = 21/c Pa.

Discussion

37.	Which of the following is correct?(a) Weight will be more for higher aspect ratio wing(b) Military aircraft have much higher aspect ratio(c) Civil aircraft are flying at hypersonic speed(d) Lofting is a conceptual design of wingThe question was posed to me in a job interview.Query is from Wing Geometry-1 in section Airfoil and Geometry Selection of Aircraft Design
Answer» Correct answer is (a) Weight will be more for higher aspect ratio WING Best explanation: Aspect ratio affects span and reference area. For same area if aspect ratio is more then, the corresponding VALUE of span will be more as well. This will increase MATERIAL requirements and which INCREASES weight of the wing.

Discussion

38.	The ratio of tip and root chord is called ____(a) taper ratio(b) sweep(c) aspect ratio(d) slopeThe question was asked in examination.I want to ask this question from Wing Geometry-2 topic in portion Airfoil and Geometry Selection of Aircraft Design
Answer» The CORRECT OPTION is (a) taper ratio Easy EXPLANATION: Taper ratio is the ratio of tip chord of WING to root chord of the wing. Aspect ratio is SQUARE of span divided by area. Sweep is angle between wing and fuselage reference line.

Discussion

39.	The cross-sectional shape of wing is called _____(a) airfoil(b) circle(c) camber(d) chordI had been asked this question during an interview for a job.My question is based upon Airfoil Selection-1 topic in portion Airfoil and Geometry Selection of Aircraft Design
Answer» CORRECT OPTION is (a) AIRFOIL Explanation: Wings are DESIGNED with airfoil cross-section. Engine inlets are typically circular in cross-sections. Camber and chord are PARAMETERS of an airfoil.

Discussion

40.	What is represented by ‘?’ in following diagram of an airfoil?(a) Thickness(b) Chord(c) Leading edge(d) Trailing edgeI got this question in final exam.Asked question is from Airfoil Selection-2 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» RIGHT choice is (a) Thickness To EXPLAIN: Thickness is distance between upper and lower SURFACE of the airfoil. LEADING and trailing edge are foremost and most rear parts of airfoil. Thickness will be different at each section of an airfoil.

Discussion

41.	An airfoil is operated with flow velocity such that the pressure on upper and lower surface are 30 Pa and 15 Pa respectively. If chord is 1m then, what will be the lift produced by this airfoil?(a) 15N(b) 15KN(c) 25N(d) 30 PaI got this question during an interview.I would like to ask this question from Airfoil Selection-3 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Right answer is (a) 15N For EXPLANATION I WOULD say: Given, Pressure on upper surface P1 = 30Pa, Pressure on LOWER surface P2 = 15Pa, chord = 1m. Now LIFT = (P1- P2)Chord = (30-15)1 = 15N.

Discussion

42.	Location of aerodynamic centre for low speed airfoil is ____(a) at quarter chord point(b) at quarter chord from leading edge(c) there is no such place(d) at half of chordThis question was posed to me in semester exam.My doubt stems from Airfoil Selection-3 in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» Right option is (a) at QUARTER chord point For explanation: AERODYNAMIC center is that point at which pitching moment curve slope becomes zero. Hence, it is the point at where pitching moment is almost INDEPENDENT of AOA. For, low speed airfoils it is LOCATED at 25% of chord from leading edge.

Discussion

43.	Which control surface is primarily used for spin recovery?(a) Rudder(b) Elevator(c) Aileron(d) FlapI got this question in a national level competition.My question is from Tail Geometry and Arrangement topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» Correct option is (a) Rudder The explanation is: During spin aircrafts is not STABLE. AIRCRAFT performs unstable yaw motion. For STABILIZING such EVENT adequate amount of rudder control is required. Flaps are used to increase lift and drag.

Discussion

44.	Pressure on the upper surface is _________(a) lower than lower surface(b) higher than lower surface(c) always 30.25 times lower surface(d) always sameThe question was asked in an internship interview.This is a very interesting question from Airfoil Selection-3 in division Airfoil and Geometry Selection of Aircraft Design
Answer» The correct CHOICE is (a) lower than lower surface The EXPLANATION: Pressure on the UPPER surface of an airfoil is lower than that of the lower surface. This will provide pressure difference which will generate lift force. Pressure will be the same if AOA is 0° in case of SYMMETRIC airfoil.

Discussion

45.	Camber represents the curvature of an airfoil.(a) True(b) FalseI have been asked this question during an online exam.I would like to ask this question from Airfoil Selection-2 in section Airfoil and Geometry Selection of Aircraft Design
Answer» Right choice is (a) True To elaborate: CAMBER of an airfoil is nothing but the curvature of it. Camber is used to produce LIFT at even ZERO angle of ATTACK. The curvature allows improvement in flow characteristics. Flow will be ATTACHED to airfoil at forward section due to the camber.

Discussion

46.	An airfoil is subjected to the flow with temperature of 288.16K and flow is ideal. The velocity at the leading edge is ____(a) 0m/s(b) 250m/s(c) 288.16m/s(d) 16.882m/sThe question was asked by my college professor while I was bunking the class.This intriguing question comes from Airfoil Selection-1 topic in division Airfoil and Geometry Selection of Aircraft Design
Answer» The correct option is (a) 0m/s Best explanation: GIVEN, TEMPERATURE T=288.16K For IDLE flow, at leading EDGE flow is brought to rest isentropically. Hence, velocity at leading edge = velocity of flow at the impact to leading edge = 0m/s.

Discussion

47.	We are using the airfoil in the aircraft as _____(a) it is a streamline shape which provides better aerodynamics(b) it is a random shape can be changed to rectangular as well(c) it is easy to make than rectangular(d) it is not a streamline bodyI have been asked this question in an interview for internship.This question is from Airfoil Selection-1 in section Airfoil and Geometry Selection of Aircraft Design
Answer» RIGHT CHOICE is (a) it is a streamline shape which provides better aerodynamics Best explanation: Airfoil is a STREAMLINED body which provides MUCH smoother flow than non-streamlined body. Airfoil is DESIGNED in such a way that it can provide better aerodynamics than any other shape. Hence, in aircraft we use airfoils.

Discussion

48.	NACA stands for _____(a) National Advisory Committee for Aeronautics(b) National Authorized Committee for Aeronautics(c) National Advisory Committee for Aero models(d) National Advisory Committee for ArithmeticThis question was posed to me in a national level competition.This interesting question is from Airfoil Selection-2 topic in chapter Airfoil and Geometry Selection of Aircraft Design
Answer» RIGHT answer is (a) NATIONAL Advisory Committee for Aeronautics To explain I WOULD say: NACA is a National Advisory Committee for Aeronautics. NACA airfoils are SIMPLE to adopt. NACA series is used to identify the airfoil.

Discussion

49.	If a sailplane is to be designed then, the aspect ratio of the sailplane will be ______(a) higher than military aircraft(b) same as military aircraft(c) lower than military aircraft(d) exactly half of the military aircraftI have been asked this question in examination.My enquiry is from Wing Geometry-1 topic in section Airfoil and Geometry Selection of Aircraft Design
Answer» Right option is (a) higher than military aircraft Easiest EXPLANATION: Sailplanes are different from military aircrafts. Sailplanes are designed to GLIDE through AIR. For such REQUIREMENTS, it will require higher lift capability which can be provided by using higher ASPECT ratio wings than military aircrafts.

Discussion

50.	For a cambered airfoil section lift is 30N. Airfoil is operating in such way that dynamic lift is 20N. Determine the lift co-efficient for the airfoil.(a) 1.5(b) 1.5N(c) 0.5(d) 0.5NThis question was posed to me during a job interview.Question is from Airfoil Selection-2 topic in portion Airfoil and Geometry Selection of Aircraft Design
Answer» Right CHOICE is (a) 1.5 Easiest EXPLANATION: GIVEN, SECTIONAL lift = 30N, dynamic lift=20N. Lift co-efficient = section lift/dynamic lift = 30/20 = 1.5.

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