Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 12, Problem 12.5P
Consider a flat plate at an angle of attack in an inviscid supersonic flow. From linear theory, what is the value of the maximum lift-to-drag ratio, and at what angle of attack does it occur
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Consider a flat plate at an angle of attack in an inviscid supersonic flow.From linear theory, what is the value of the maximum lift-to-drag ratio,and at what angle of attack does it occur?
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Chapter 12 Solutions
Fundamentals of Aerodynamics
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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- Using linearized theory, calculate the lift and drag coefficients for a flat plate at a 5◦ angle of attack in a Mach 3 flow.arrow_forwardConsider a NACA 2412 airfoil at an angle of attack of 4 degrees. If the freestream Mach number is 0.72 determine the lift coefficient.arrow_forwardAt subsonic speeds, how do the lift coefficient CL and drag coefficientCD for a wing vary with angle of attack α?arrow_forward
- An infinitely thin flat plate is operating at a given freestream Mach number of 2.4 where ambient pressure is 0.53 atm. Calculate Lift to drag ratio for an given angle of attack of 4 degrees using both shock-expansion theory and linearized theoryarrow_forwardAn infinitely thin flat plate is operating at a given freestream Mach number of 2.4 where ambient pressure is 0.53 atm. Calculate Drag per unit span for an given angle of attack of 4 degrees using both shock-expansion theory and linearized theoryarrow_forwardAn infinitely thin flat plate is operating at a given freestream Mach number of 2.4 where ambient pressure is 0.53 atm. Calculate Lift per unit span for an given angle of attack of 4 degrees using both shock-expansion theory and linearized theoryarrow_forward
- How does the construction of a low speed and high-speed aircraft differ? What effects can be expected from supersonic flow on fuselage structure, discuss steps to minimise these effects.arrow_forwardStarting from the basic principles, drive an expression to compute the lift coefficient when the free stream Mach number equals (0.8)arrow_forwardConsider a flat plate at an angle of attack in a viscous supersonic flow; i.e.,there is both skin friction drag and wave drag on the plate. Use lineartheory for the lift and wave-drag coefficients. Denote the total skin frictiondrag coefficient by C f , and assume that it does not change with angle ofattack. (a) Derive the expression for the angle of attack at which maximumlift-to-drag ratio occurs as a function of C f and freestream Mach number.(b) Derive the expression for the maximum lift-to-drag ratio as a functionof C f and freestream Mach number M.arrow_forward
- In 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_forwardConsider an infinitely thin flat plate with a 1 m chord at an angle of attackof 10◦ in a supersonic flow. The pressure and shear stress distributions onthe upper and lower surfaces are given by pu = 4 × 104(x − 1)2 +5.4 × 104, pl = 2 × 104(x − 1)2 + 1.73 × 105, τu = 288x−0.2, andτl = 731x−0.2, respectively, where x is the distance from the leading edgein meters and p and τ are in newtons per square meter. Calculate thenormal and axial forces, the lift and drag, moments about the leadingedge, and moments about the quarter chord, all per unit span. Also,calculate the location of the center of pressure.arrow_forwardA 20◦ half-angle wedge is mounted at 0◦ angle of attack in the testsection of a supersonic wind tunnel. When the tunnel is operating, thewave angle from the wedge leading edge is measured to be 41.8◦. Whatis the exit-to-throat area ratio of the tunnel nozzle?arrow_forward
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