Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 14, Problem 14.2P
Consider a fiat plate at
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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.
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- Air at 20C flows at 1 m/s between two parallel flat plates spaced 5 cm apart. Estimate the distance from the entrance to the point at which the hydrodynamic boundary layers meet.arrow_forwardAir at 1000C flows at an inlet velocity of 2 m/s between two parallel flat plates spaced 1 cm apart. Estimate the distance from the entrance to the point where the boundary layers meet.arrow_forwardConsider a flat-plate airfoil at an angle of attack of 6°. TheMach number is Ma∞ = 3.2 and the stream pressure p∞ isunspecified. Calculate the predicted lift and drag coefficientsby (a) shock-expansion theory and (b) Ackeret theory.arrow_forward
- Consider a flat plate at an angle of attack α to a Mach 2.4 airflow at 1 atmpressure. What is the maximum pressure that can occur on the platesurface and still have an attached shock wave at the leading edge? At whatvalue of α does this occur?arrow_forwardConsider a cone at zero angle of attack in a hypersonic flow. (Hypersonic flow is very high-speed flow, generally defined as any flow above a Mach number of 5.) The half-angle of the cone is θc, as shown inthe figure. An approximate expression for the pressure coefficient on the surface of ahypersonic body is given by the newtonian sine-squared law : Cp = 2 sin2 θcNote that Cp, hence, p, is constant along the inclined surface of the cone. Along the base of the body, we assume that p = p∞. Neglecting the effect of friction, obtain an expression for the drag coefficient of the cone, where CD is based on the area of the base Sb.arrow_forwarda balloon is 4 m in diameter and contains helium at 125 kpa and 15°c. balloon material and payload weigh 200 n, not including the helium. estimate (a) the terminal ascent velocity in sea-level standard air, (b) the final standard altitude (neglecting winds) at which the balloon will come to rest, and (c) the minimum diameter (< 4 m) for which the balloon will just barely begin to rise in sea-level standard air.arrow_forward
- Pu Wings at supersonic flows can be modeled as a flat plate that have constant pressure distributions across their chord. Ignoring any viscous effects determine the drag force of the wing section. Round to the full N/m and make sure you use the standard sign convention (i.e. right hand rule). Use the following data chord = 1.5 m epsilon = 5 degree angle of attack = 20.0 alpha upper surface pressure = 60,000 Pa lower surface pressure = 150,000 Paarrow_forwardConsider 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.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_forward
- In low-speed incompressible flow, the peak pressure coefficient (at theminimum pressure point) on an airfoil is −0.41. Estimate the criticalMach number for this airfoil, using the Prandtl-Glauert rule.arrow_forwardSR-71 Blackbird cruised at Mach 3.2 at 80,000 ft altitude. However, it was designed to leak a finite amount of fuel during takeoff because the surface temperature of the aircraft increases (which leads to expansion of the surfaces) due to aerodynamic heating during flight. Estimate the flow temperature around the aircraft during cruise condition. Assume specific heat ratio is 1.4.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
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