Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 4, Problem 4.5P
Consider a thin, symmetric airfoil at
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Consider 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.
Consider an airfoil at 12◦ angle of attack. The normal and axial forcecoefficients are 1.2 and 0.03, respectively. Calculate the lift and dragcoefficients.
Consider the North American P-51D Mustang. Its wingspan is 37 ft; wing area is 233.6 ft2; and gross weight is 10,100 lb. Assume the Oswald efficiency factor as 0.8. The airplane is flying in steady; level flight at a velocity of 750 mi/h at a standard altitude of 5000 ft.
Calculate the drag due to lift using the concept of span loading.
Chapter 4 Solutions
Fundamentals of Aerodynamics
Ch. 4 - Consider the data for the NACA 2412 airfoil given...Ch. 4 - Consider an NACA 2412 airfoil with a 2-m chord in...Ch. 4 - Starting with the definition of circulation,...Ch. 4 - Starting with Equation (4.35), derive Equation...Ch. 4 - Consider a thin, symmetric airfoil at 1.5 angle of...Ch. 4 - The NACA 4412 airfoil has a mean camber line given...Ch. 4 - For the airfoil given in Problem 4.6, calculate...Ch. 4 - Compare the results of Problems 4.6 and 4.7 with...Ch. 4 - Starting with Equations (4.35) and (4.43), derive...Ch. 4 - For the NACA 2412 airfoil, the lift coefficient...
Ch. 4 - Consider again the NACA 2412 airfoil discussed in...Ch. 4 - For the airfoil in Problem 4.11, calculate the...Ch. 4 - In Section 3.15 we studied the case of the lifting...Ch. 4 - The question is often asked: Can an airfoil fly...Ch. 4 - The airfoil section of the wing of the British...Ch. 4 - For the conditions given in Problem 4.15, a more...
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- For 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_forwardConsider a finite wing with an aspect ratio of 3. Assume an elliptical liftdistribution. The lift slope for the airfoil section is 0.1/degree. Calculateand compare the lift slopes for (a) a straight wing, and (b) a swept wing,with a half-chord line sweep of 45 degrees.arrow_forwardCompute the lift and drag coefficients for a symmetric, diamond-shaped airfoil with a thickness-to-chord ratio t/c equal to 0.10, flying at Mach 3.5 in air (y=1.4) at zero angle of attack. M₂ - 3.5arrow_forward
- A 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_forward4. Write down the lift curve equation (lift coefficient versus angle of attack) according to the thin airfoil theory. What is the amount of the lift curve slope for a thin airfoil ? NACA 2412 airfoil Eg. (4,57) C 2.0 - 1.e 1.2 Lift coefficient 0.8 0.4 -0.4 Eg. (4,64) -0.1 Moment coefficient -0.8 -0.2 -1.2 -0.3 Re - 3.1 X 10 O Re- 8.9 x 10 -0.4 16 a, degrees de; = ? daarrow_forwardCalculate the lift and moment coefficients using thin airfoil theory approximations for a range of angle of attacks for the following airfoils below. 1. NACA 0008 2. NACA 0018 3. AG04 4. Clark-Y 5. NACA 2415 Angle of attacks taken to be between 0 degree to 15 degree. Also calculate value of Cm at quarter chord. (Cm at 0.25c) for each. Reynolds number 1000000arrow_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_forwardA flagpole 16 m high has the shape of a cylinder 100 mm in diameter. The air temperature is 30°C and the atmospheric pressure is 101 KPaa. With what speed is the air blowing against the pole if the moment developed at the base is 2.7 KN.m? The drag coefficient is 1.3.arrow_forward4arrow_forward
- Answer this olearrow_forward6. Write down the lift curve equation (lift coefficient versus angle of attack) according to the thin airfoil theory. What is the amount of the lift curve slope for a thin airfoil ? NACA 2412 airfoil LEg. (4.57) * 2.0 1.6 1.2 Lift coefficient 0.8 0.4 -0.4 Eg. (4,64) --0.1 Moment coefficient -0.8- -1.2 -0.3 • Re -3.1 X 10 O Re -8.9 X 10 L-0.4 -8 16 24 a, degrees dc =? daarrow_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
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