Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 8, Problem 8.16P
In the test section of a supersonic wind tunnel, a Pilot tube in the flow reads a pressure of 1.13 atm. A static pressure measurement (from a pressure tap on the sidewall of the test section) yields 0.1 atm. Calculate the Mach number of the flow in the test section.
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Chapter 8 Solutions
Fundamentals of Aerodynamics
Ch. 8 - Consider air at a temperature of 230 K. Calculate...Ch. 8 - The temperature in the reservoir of a supersonic...Ch. 8 - At a given point in a flow, T=300K,p=1.2atm, and...Ch. 8 - At a given point in a flow, T=700R,p=1.6atm, and...Ch. 8 - Consider the isentropic flow through a supersonic...Ch. 8 - Consider the isentropic flow over an airfoil. The...Ch. 8 - The flow just upstream of a normal shock wave is...Ch. 8 - The pressure upstream of a normal shock wave is 1...Ch. 8 - The entropy increase across a normal shock wave is...Ch. 8 - The how just upstream of a normal shock wave is...
Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Repeat Problems 8.11 and 8.12 using (incorrectly)...Ch. 8 - Derive the Rayleigh Pitot tube formula, Equation...Ch. 8 - On March 16, 1990, an Air Force SR-71 set a new...Ch. 8 - In the test section of a supersonic wind tunnel, a...Ch. 8 - When the Apollo command module returned to earth...Ch. 8 - The stagnation temperature on the Apollo vehicle...Ch. 8 - Prove that the total pressure is constant...
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- Nonearrow_forwardThe nozzle of a supersonic wind tunnel has an exit-to-throat area ratioof 6.79. When the tunnel is running, a Pitot tube mounted in the testsection measures 1.448 atm. What is the reservoir pressure for thetunnel?arrow_forward1. A uniform supersonic air flow at Mach 2.0 passes over a wedge. An oblique shock, making an angle of 40° with the flow direction, is attached to the wedge. If the static pressure and temperature in the freestream are 0.5 x 10 N/m2 and 0°C, determine the static pressure and temperature behind the wave, the Mach number of the flow passing over the wedge, and the wedge angle.arrow_forward
- Q1/2. Air flow isentropically through a supersonic convergent divergent nozzle with 0.6 kg/s. At the inlet, the pressure is 700 kPa and the temperature is 300 K, and the area is 7 cm. If the exit area is 14 cm? calculate: a- The stagnation pressure and temperature. b- The exit Mach number. c- The exit pressure and temperature. d- The area and velocity at the throat. e- What will be the maximum rate of flow and the corresponding exit Mach number if the flow completely subsonic in the nozzle?arrow_forward1. The pressure in a reservoir of a supersonic wind tunnel is 178.42 KPa, and the density is 1.7 kg/m3. The pressure outside the nozzle is atmospheric, with a Mach number of 1.8 at the exit. The throat has a surface area of 0.15m2 with a Mach Number of 1.0. Solve for the value of the following: (a) Reservoir temperature and speed of sound. (b) Pressure, density, temperature and speed of sound at the throat. (c) Mass flow at the exit.arrow_forwardA high-speed subsonic private jet is flying at a pressure altitude of 11 km. A Pitot tube on the wing tip measures a pressure of 4.21 x 104 N/m². • Calculate the Mach number at which the airplane is flying. • If the ambient air temperature is 225 K, calculate; • the true airspeed and • the calibrated airspeedarrow_forward
- In a low-speed subsonic wind tunnel with a closed test section, a static pressure tap on the wall of the tunnel test section measures 0.98 atm. The temperature of the air in the test section is 80 ˚F. A Pitot tube is inserted in the middle of the flow in the test section in order to measure the flow velocity. The pressure measured by the Pitot tube is 2,200 psf. Calculate the flow velocity (in KPH) in the test section.arrow_forwardStagnation pressure and temperature probes are located on the nose of a supersonic aircraft. At 35,000 ft altitude a normal shock stands in front of the probes. The temperature probe indicates T0 = 420F behind the shock. Calculate the Mach number and air speed of the plane. Find the static and stagnation pressures behind the shock. Show the process and the static and stagnation state points on a Ts diagram.arrow_forwardPlease provide correct answerarrow_forward
- If the Mach number behind a normal shock is equal to 0.4, calculate the ratio of the readings of pitot tubes located before and after the shock.arrow_forward1. An air flow with a Mach number of 2.0 passes through an oblique shock wave inclined at an angle of 45° . Find the flow deflection angle α. 2. An air flow with a Mach number of 8 is deflected by a wedge through an angle α. What is the maximum value of α for an attached oblique shock?arrow_forwardQuestion Two Consider one-dimensional air isentropic flow through a duct. At a certain section of this duct, the velocity is 360 m/s, the temperature is 45 °C, and the pressure is 120 kPa. Find the Mach number and the stagnation temperature and pressure at this point in the flow. If the temperature at some other point in the flow is 90 °C, find the Mach number and pressure at this point in the flow.arrow_forward
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