Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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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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- I need the answer quicklyarrow_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_forwardAn 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_forward
- Consider a normal shock wave in air where the upstream flow properties are u1 = 660 m/s, T1 = 288 K, and p1=1 atm. Calculate the velocity, temperature, and pressure downstream of the shock.arrow_forwardPlease do correctly.arrow_forwardCan the Mach number of a fluid be greater than 1 after a normal shock wave? Explain.arrow_forward
- I need the answer as soon as possiblearrow_forwardI need the answer as soon as possiblearrow_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_forward
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