Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Textbook Question
Chapter 9, Problem 9.6P
Consider a flat plate at an angle of attack a to a Mach 2.4 airflow at 1 atm pressure. What is the maximum pressure that can occur on the plate surface and still have an attached shock wave at the leading edge? At what value of
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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?
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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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- An explosion occurs which creates a plane normal shock wave propagating into a region of air that is at rest (stagnation pressure po=1.0135×105Pa) and (stagnation temperature of To=290K). The speed of the shock is 1700 m/s. The air is modelled as an inviscid fluid, specific heat ratio γ=1.4 and gas constant R=287~J/kg⋅K. Calculate the air speed in m/s, relative to a stationary observer in the region behind the shock?arrow_forwardConsider 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_forwardASAParrow_forward
- Can the Mach number of a fluid be greater than 1 after a normal shock wave? Explain.arrow_forwardConsider a typical air flow around a cruising jetliner at 10km altitude. The speed is now 810 km/h, while the ambient conditions are 0.414 kg/m³ , 0.261 atm and -50°C. At the stagnation point, the temperature rises over by 25°C, while the density and pressure changes by more than 30% and 45 % respectively. Classify the following situations as compressible/incompressible flowarrow_forwardThe velocity and temperature behind a normal shock wave are 329 m/s and 1500 K, respectively. Calculate the velocity in front of the shock wave.arrow_forward
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- An infinitely thin flat plate is operating at a given freestream Mach number of 2.4 where ambient pressure is 0.53 atm. Calculate Pressure acting on the lower surface of the plate for an given angle of attack of 4 degrees using both shock-expansion theory and linearized theoryarrow_forwardvortex generators on the upper surface of a wing will a. decrease the spanwise flow at high Mach numbers b. increase the critical Mach number c. decrease the intensity of the shockwave effects d. increase intensity of the shockwave effectsarrow_forwardAn infinitely thin flat plate is operating at a given freestream Mach number of 2.4 where ambient pressure is 0.53 atm. Calculate Pressure acting on the upper surface of the plate for an given angle of attack of 4 degrees using both shock-expansion theory and linearized theoryarrow_forward
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