Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 9, Problem 9.7P
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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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- When a pitot-static tube is placed in a supersonic flow, the probe reads P = 190 kPa and P = 150 kPa. If the stagnation temperature is 400 K, estimate the freestream Mach number and velocity.arrow_forwardA normal shock wave propagates into atmosphere where the atmospheric temperature is equal 336K and pressure is 0.8 atm. For a given pressure ratio of p2/p₁ 41 to T₁ = a) Assuming stationary atmosphere, calculate the shock wave velocity and the velocity induced behind the shock wave. Also calculate the temperature, total pressure and total temperature of the fluid particles behind the shock wave using equations for moving normal shock. =arrow_forward5. Airflow at Mach 2 passes through an oblique shock as shown and deflects 10°. A second oblique shock reflects from the solid wall. What is the pressure ratio (p3/p₁) across the two-shock system? Assume that there is no boundary layer near the wall, so the flow is uniform in each of the regions bounded by the shocks and that y = 1.4. Reflected oblique shock 10° M=2arrow_forward
- The converging-diverging flow domain is shown in Figure 1. The inlet diameter is 0.2 m, the throat diameter is 0.15 m, and the outlet diameter is 0.24 m. The axial distance from the inlet to the throat is 0.30 m—the same as the axial distance from the throat to the outlet. At the inlet section, the stagnation pressure P0 is set to 220 kPa (absolute), while the stagnation temperature T0, at the inlet is set to 300 K.arrow_forwardFollowing manufacture, it is planned to use the biodiesel to power a rocket which uses a de Laval nozzle. The nozzle throat has a cross-sectional area of 5.5 x 10-3 m2. The hot air enters the nozzle at 300 bar(atm) and 77 °C and exits nozzle at 1 bar(atm). If we have choked flow at the throat, what is the pressure at the throat? Considering we have choked flow at the throat, what is the flow rate at the exit of the nozzle?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
- Q4//B/ Heat is removed from airflow in a constant-area duct, if flow enters at Mach 2.20, T1= 310 K, and p1= 120 kPa. Determine exit Mach number and the amount of heat removing from the duct if the static temperature at exit section is 255 K. justify your solution for M2arrow_forwardQ6/ Air is heated as it flows subsonically through a duct. When the amount of heat transfer reaches 60 kJ/kg, the flow is observed to be choked, and the velocity and the static pressure are measured to be 620 m/s and 270 kPa. Disregarding frictional losses, determine the velocity, static temperature, and static pressure at the duct inlet.arrow_forwardConsider a low-speed open-circuit subsonic wind tunnel. The tunnel is turned on, and the pressure difference between the inlet (the settling chamber) and the test section is read as a height difference of 10 cm on a U-tube mercury manometer. (The density of liquid mercury is 1.36 × 104 kg/m3.) Assume that a Pitot tube is inserted into the test-section flow of the wind tunnel. The tunnel test section is completely sealed from the outside ambient pressure. Calculate the total pressure measured by the Pitot tube, assuming the static pressure at the tunnel inlet is atmospheric. Given that A2/A1 = 1/12. (Round the final answer to two decimal places.) The total pressure measured by the Pitot tube is × 105 N/m2.arrow_forward
- nozzle with an expansion ratio of 10. The inlet - 8. Air is flowing in a converging-diverging stagnation pressure P 7.0 10³ N/m², and the pressure in the exhaust region is 1 atm. A normal shock wave is standing in the nozzle divergence. Determine the area ratio at which the normal shock wave occurs. Assume y = 1.40.arrow_forwardA nozzle for a supersonic wind tunnel is designed to achieve a Mach number of 3.2, with a velocity of 2500 m/s, and a density of 1.0 kg/ m³ in the test section. Find the temperature and pressure in the test section and the upstream stagnation conditions. The fluid is helium. Te = i Pe= To= Po= Mi K kPa K kPaarrow_forward4. Determine the upstream Mach number, considering an oblique shock wave with ew = 32° and a pressure ratio, P2/P1 = 3.0.arrow_forward
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