Connect with LearnSmart for Anderson: Fundamentals of Aerodynamics, 6e
6th Edition
ISBN: 9781259683268
Author: Anderson, John
Publisher: Mcgraw-hill Higher Education (us)
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Chapter 12, Problem 12.3P
Consider a diamond-wedge airfoil such as shown in Figure 9.26, with a half-angle
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Consider a diamond-wedge airfoil such as shown , with ahalf-angle ε = 10◦. The airfoil is at an angle of attack α = 15◦ to a Mach3 freestream. Using linear theory, calculate the lift and wave-dragcoefficients for the airfoil. Compare these approximate results with thosefrom the exact shock-expansion theory obtained.
4. For the 2D symmetric airfoil with a diamond profile A = 5° as shown in figure, compute
the lift and drag coefficients in the supersonic flow through air (y = 1.4) and free stream
mach number M = 2
%3D
with AOA a = 10°.
(2
(3
Consider the same Lockheed F-104 supersonic fighter shown, with the same flight conditions of Mach 2 at an altitude of 11 km. For these conditions the wing angle of attack is α = 0.035 rad = 1.98◦. Assume the chord length of the airfoil is 2.2 m, which is approximately the mean chord length for the wing. Also, assume fully turbulent flow over the airfoil. Calculate: (a) the airfoil skin friction drag coefficient, and (b) the airfoil wave-drag coefficient. Compare the two values of drag.
Chapter 12 Solutions
Connect with LearnSmart for Anderson: Fundamentals of Aerodynamics, 6e
Ch. 12 - Using the results of linearized theory, calculate...Ch. 12 - For the conditions of Problem 12.1, calculate the...Ch. 12 - Consider a diamond-wedge airfoil such as shown in...Ch. 12 - Equation (12.24). from linear supersonic theory....Ch. 12 - Consider a flat plate at an angle of attack in an...Ch. 12 - Consider a flat plate at an angle of attack in a...Ch. 12 - Using the same flight conditions and the same...Ch. 12 - The result from Problem 12.6 demonstrates that...
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- The lift curve for the 4 digit NACA 2421 airfoil is shown in the Figure. Consider a wing with AR = 8, sweep ? = 15 ???, and airplane flies at the Mach number equal to 0.3. The Oswald efficiency span number ?1 = 0.95 Calculate the lift slope for the finite wing. Clearly show the formula and explanations in the solution.arrow_forwardConsider a rectangular wing with a NACA 23015 airfoil that has a maximum thickness of 2.7 ftand a wing span of 20ft if the freestream density, temperature and mach number are 1.22x10^-3 slugs/ft^3, 444 R and 0.2 respectively. Calculate the lift if the coefficient of lift is 0.8. English unit not metricarrow_forwardCalculate the lift and drag coefficiens for a flat plate at a 5◦ angle of attack in a Mach 3 flow.arrow_forward
- Compute 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_forwardA tiny scratch in the side of a supersonic wind tunnelcreates a very weak wave of angle 17°, as shown in Fig., after which a normal shock occurs. The air temperaturein region (1) is 250 K. Estimate the temperaturein region (2).arrow_forwardConsider a flat plate at zero angle of attack in a hypersonic flow at Mach10 at standard sea level conditions. At a point 0.5 m downstream from theleading edge, the local shear stress at the wall is 282 N/m2. The gastemperature at the wall is equal to standard sea level temperature. At thispoint, calculate the velocity gradient at the wall normal to the wall.arrow_forward
- Air flows at Ma = 2.5 past a half-wedge airfoil whoseangles are 4°, as in Fig. Compute the lift and dragcoefficient at α equal to (a) 0° and (b) 6°.arrow_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_forwardIn the test section of a high-speed subsonic wind tunnel operating at Mach number of 0.75, a NACA 2415 airfoil is mounted. The compressible lift coefficient measured for that airfoil is 0.7. Determine the equivalent incompressible lift coefficient and the angle of attack at which the airfoil is flying.arrow_forward
- Consider a flat plate at α = 20◦ in a Mach 20 freestream. Using straightnewtonian theory, calculate the lift- and wave-drag coefficients. Comparethese results with exact shock-expansion theory.arrow_forwardA NACA 2412 airfoil with a chord of 0.64m is flying in an airstream of standard sea level conditions. The freestream velocity is 70 m/s. Given the lift per unit span is 1,254 N/m. By using the experimental data for NACA 2412 data plot in Figure Q1c, investigatethe angle of attack of the airfoil and the analyze the value of drag per unit spanof the airfoil. Given that at standard sea level, ?=1.789×10-5 kg/m.s.arrow_forward7. Using shock-expansion theory, calculate the lift and drag coefficients for a flat plate at a 5° angle of attack in a Mach 2.4 flow. Pa M, P3arrow_forward
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