Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 9, Problem 9.16P
Consider a circular cylinder (oriented with its axis perpendicular to the flow) and a symmetric diamond-wedge airfoil with a half-angle of
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For the shown continuously circular (i.e. not parabolic) arc airfoil with sharp corners of each
8 = 5°, at 0° angle of attack with Mo = 3 and assuming steady, two-dimensional flow,
calculate the:
a) pressure coefficient distribution;
b) lift coefficient;
c) drag coefficient.
Moo
8
*********
8
Consider an infinite wing with an NACA
1412 airfoil section and a chord length of
1 m. The wing is at an angle of attack 5°
in an airflow velocity of 44 m/s at
standard sea-level conditions. Calculate
the lift, drag, and moment about the
quarter chord per unit span.
Consider a finite wing with a NACA 1412 airfoil section with a wingspan of 12 ft and a wing area of 240 sq ft. The airflow velocity is 100 ft/s at 2000 ft. Calculate the lift and drag if the CL is 1.5 and Cdo is 0.035. span efficiency factor is 0.8.
Chapter 9 Solutions
Fundamentals of Aerodynamics
Ch. 9 - A slender missile is flying at Mach 1.5 at low...Ch. 9 - Consider an oblique shock wave with a wave angle...Ch. 9 - Equation (8.80) does not hold for an oblique shock...Ch. 9 - Consider an oblique shock wave with a wave angle...Ch. 9 - Consider the flow over a 22.2 half-angle wedge. If...Ch. 9 - Consider a flat plate at an angle of attack a to a...Ch. 9 - A 30.2 half-angle wedge is inserted into a...Ch. 9 - Consider a Mach 4 airflow at a pressure of 1 atm....Ch. 9 - Consider an oblique shock generated at a...Ch. 9 - Consider the supersonic flow over an expansion...
Ch. 9 - A supersonic flow at M1=1.58 and p1=1atm expands...Ch. 9 - A supersonic flow at M1=3,T1=285K, and p1=1atm is...Ch. 9 - Consider an infinitely thin flat plate at an angle...Ch. 9 - Consider a diamond-wedge airfoil such as shown in...Ch. 9 - Consider sonic flow. Calculate the maximum...Ch. 9 - Consider a circular cylinder (oriented with its...Ch. 9 - Consider the supersonic flow over a flat plate at...Ch. 9 - (The purpose of this problem is to calculate a...Ch. 9 - Repeat Problem 9.18, except with =30. Again, we...Ch. 9 - Consider a Mach 3 flow at 1 atm pressure initially...Ch. 9 - The purpose of this problem is to explain what...
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- A 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_forwardAn airplane with a NACA 23012 airfoil cruises at 150 m/s at an altitude of 6000 m. The airfoil has an aspect ratio of 10 with a span of 36 m. Using the airfoil data as in Fig. 3, determine the lift and drag forces. Then determine the power required to overcome drag. Consider the airplane flying at an angle of attack equal to 2. wWw wwarrow_forward5) A wedge airfoil is placed in a supersonic flow at M = 2. G. Calculate the lift and drag coefficients at 4° angle of attack. Hint: Consider the free-stream pressure at the base of the airfoil. 4° Pooarrow_forward
- Study how the lift and drag varies with a change in angle of attack for a flat plate, symmetrical airfoil and asymmetrical airfoil when Re is below 50’000 and compare it to higher Re.arrow_forwardAir flows parallel to the surface of a smooth flat plate 10 m long. The boundary thickness at the leading edge. The Reynolds number at the trailing edge of the plate is 107. Calculate the total drag force due to skin friction on one side of the plate per unit width. Assume that for laminar boundary layer, up to Re, = 5 x 105, the skin friction coefficient is Gr 1.328(Re)-1/2 and for turbulent boundary layer 0.078(Re)-1/5. Take the density of air as 1.2 kg/m³ and Its dynamic viscosity as 1.8 x 10-5 kg/msarrow_forwardAn aeroplane has a rectangular planform wing with a span of 12 m and a chord of 2 m. The aircraft is flying at cruising speed of 62 m/s at an altitude of 3,200 m. Assume that the skin-friction drag on the wing can be approximated by the drag on a flat plate of the same dimensions at 0° incidence. Calculate the skin-friction drag and the maximum boundary layer thickness assuming (i) a completely laminar flow, and (ii) a completely turbulent flow. Assume μ = 1.63.10-5 N s/m², p = 0.89 kg/m³. Then, calculate the skin-friction drag accounting for transition at Recr= 5.105. What would the drag benefit be if transition was delayed to a Reynolds number of 106? What reduction in a propeller engine power would this imply in the cruise condition, assuming a propeller efficiency of 0.88?arrow_forward
- The original Flettner rotor ship was approximately 100 ftlong, displaced 800 tons, and had a wetted area of 3500 ft2. Assketched in Fig. , it had two rotors 50 ft high and 9 ft indiameter rotating at 750 r/min, which is far outside the range. The measured lift and drag coefficients for eachrotor were about 10 and 4, respectively. If the ship is mooredand subjected to a crosswind of 25 ft/s, as in Fig. whatwill the wind force parallel and normal to the ship centerlinebe? Estimate the power required to drive the rotors.arrow_forward6. 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_forwardIs it possible to mix multiple airfoil profiles for an aircraft wing? If so, then why.arrow_forward
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