Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Textbook Question
Chapter 5, Problem 5.6P
Consider a finite wing with an aspect ratio of 6. Assume an elliptical lift distribution. The lift slope for the airfoil section is 0.1/degree. Calculate and compare the lift slopes for (a) a straight wing, and (b) a swept wing, with a half-chord line sweep of 45 degrees.
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Consider a finite wing with an aspect ratio of 5. Assume an elliptical lift distribution. The lift slope for the airfoil section is
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Chapter 5 Solutions
Fundamentals of Aerodynamics
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- Consider a finite wing with an aspect ratio of 5. Assume an elliptical lift distribution. The lift slope for the airfoil section is 0.1/degree. Calculate the lift slope (in 1/rad) for a straight wing. (Round the final answer to three decimal places.) The lift slope for a straight wing is per radian.arrow_forwardConsider a thin, symmetric airfoil at 1.5◦ angle of attack. From the resultsof thin airfoil theory, calculate the lift coefficient and the momentcoefficient about the leading edge.arrow_forwardConsider a finite wing with an aspect ratio of 5. Assume an elliptical lift distribution. The lift slope for the airfoil section is 0.1/degree. Identify the correct statement about the lift slopes for the straight wing and the swept wing with a half-chord line sweep of 45 degrees. Multiple Choice A swept wing with a half-chord line sweep of 45 degrees has a slightly more lift slope than a straight wing. A swept wing with a half-chord line sweep of 45 degrees has a substantially more lift slope than a straight wing. A swept wing with a half-chord line sweep of 45 degrees has a slightly less lift slope than a straight wing. A swept wing with a half-chord line sweep of 45 degrees has a substantially less lift slope than a straight wing.arrow_forward
- Consider a swept wing with a taper ratio of 0.2, a leading edge sweep angle of 70 degrees, and and aspect ratio of 1.5. Calculate the lift coefficient at M = 2 at an angle of attack of 2 degrees. (one answer only) * 0.059 0.061 0.063 0.065arrow_forwardExplain how the three wing planforms, have the same local lift distribution local lift coefficient. All wings twisted to have elliptic distribution of lift 1(y) root tip shown below, could while having different cl tip cl(y) cl root root root tip tiparrow_forwardA given symmetrical airfoil has a lift curve slope of 6 per radian. What is the angle of attack (in degrees) corresponding to a lift coefficient of 0.4? Round off answer to the nearest tenths. (for symmetrical airfoils zero-lift angle of attack is 0°)arrow_forward
- 6. For a wing of aspect ratio AR, having an elliptical lift distribution, the induced drag coefficient is (where CL is the lift coefficient) (b) TAR CL 2TAR (d) TAR2 (a) TAR 7. Laminar flow airfoil is used to reduce the (a) pressure drag (b) induced drag (c) skin friction drag (d) wave drag 8. Set the aerofoil your chosen angle of incidence is gives a good installed pressure distribution. (a) twelve degree (b) ten degree (c) eight degree 9. The critical Mach number for a thick airfoil will be (a) lesser than a thin airfoil. (b) greater than a thin airfoil. (c) equal to a thin airfoil. (d) cannot be related to thin airfoil 10. What is the dimension for Lift coefficient? (a) N/s (b) kg/N (c) Dimensionlessarrow_forwardConsider a swept wing with a taper ratio of 0.2, a leading edge sweep angle of 70 degrees, and and aspect ratio of 1.5. Calculate the lift coefficient at M = 2 at an angle of attack of 2 degrees. (one answer only) * 0.059 0.061 0.063 0.065 Wrong answer means thumbs downarrow_forwardA wing has a planform area S of 200 ft? and a total span b of 40 feet. The airfoils are symmetric all along the span. The airfoil has a 2-D lift curve slope of 27 per radian. The wing has a rectangular planform, and thus has zero taper. The wing is untwisted. a. Compute the lift coefficient C and the drag coefficient Coi at an angle of attack of 4 degrees. Use two terms in the series expansion for circulation. T= 2bV,[4, sin ø + A, sin 3ø] b. Repeat the above calculation, now with just one term T=2bVA1sino. Compare the lift drag coefficient C and Cp values to problem #2 above. c. Compare the results for drag coefficient from part (b) above with that for an elliptically loaded wing at this lift coefficient.arrow_forward
- Consider the NACA 2412 airfoil discussed . The airfoil is flying at a velocity of 60 m/s at a standard altitude of 3 km . The chord length of the airfoil is 2 m. Calculate the lift per unit span when the angle of attack is 4◦.arrow_forwardFor a 10 deg included angle wedge at 0 deg AOA, calculate the lift and drag coefficient (Cl = 0, Cd = 0.082). Then calculate the lift and drag at AOA=4 deg (Cl = 0.16, Cd = 0.093). Let M=2.arrow_forwardAn untwisted wing uses the same airfoil along the span and the lift distribution can be considered elliptical. Data: mo = 6.1/rad b = 30 ft; S = 150ft²; a = 4°; aoL = -2.5° Find: CL(a); Cpi; CL; and Induced angle of attack at y = 5 ftarrow_forward
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