Connect with LearnSmart for Anderson: Fundamentals of Aerodynamics, 6e
6th Edition
ISBN: 9781259683268
Author: Anderson, John
Publisher: Mcgraw-hill Higher Education (us)
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Chapter 9, Problem 9.11P
A supersonic flow at
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The flow just upstream of a normal shock wave is given by p1 = 1 atm,T1 = 288 K, and M1 = 2.6. Calculate the following properties justdownstream of the shock: p2, T2, ρ2, M2, p0,2, T0,2, and the change inentropy across the shock.
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An airplane is flying at a standard altitude of 10,000 ft. A Pitot tube mounted at the nose measures a pressure of 2220 lb/ft2. The airplane is flying at a high subsonic speed, faster than 300 mph. The flow should be considered compressible. Calculate the velocity of the airplane.
Chapter 9 Solutions
Connect with LearnSmart for Anderson: Fundamentals of Aerodynamics, 6e
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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- Consider 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_forwardA sample of steam is at a pressure of 300 kPa (x=0.75). Solve for its entropy.arrow_forwardA Pitot tube inserted at the exit of a supersonic nozzle reads8.92 × 104 N/m2. If the reservoir pressure is 2.02 × 105 N/m2, calculatethe area ratio Ae/A∗ of the nozzle.arrow_forward
- Good day. Here is my question (10) Consider a low-speed subsonic wind tunnel with a 12/1 contraction area ratio for the nozzle. If the flow in the test section is at a standard sea level conditions with a velocity of 50 m /s, calculate the height difference in a U-tube mercury manometer with one side connected to the nozzle inlet and the other to the test section. ρHg = 13.6 x 103 kg/ m^3.arrow_forwardEstimate the speed of sound of carbon monoxide at 200-kPa pressure and 300°C in m/s.arrow_forwardAn engineer is designing a subsonic wind tunnel. The test section is to have a cross-sectional area of 4 m2 and an airspeed of 60 m/s. The air density is 1.2 kg/m3. The area of the tunnel exit is 10 m2. The head loss through the tunnel is given by hL=0.025VT2/2g, where VT is the airspeed in the test section. Calculate the power needed to operate the wind tunnel. Hint: Assume negligible energy loss for the flow approaching the tunnel in region A, and assume atmospheric pressure at the outlet section of the tunnel. Assume α = 1.0 at all locations.arrow_forward
- The tank is filled with air at 20°C and 139 kPa in stationary condition. Air is leaving the tank with flowing in a nozzle under steady-state condition. The flow is under isentropic and subsonic condition. The nozzle exit area is 18,59 cm?. After leaving from the nozzle, air strikes a vertical plate. Define the force [N] required to hold the plate stationary. (Note: Assume Pe=1 atm, kair=1.4, Rair=287 J/kg.K) Pte F Yanıt:arrow_forward3. Assume a supersonic flow with M=2, P=1 atm, and T=288 K that is deflected via 15° at a compression corner. Determine M, P, T, as well as PO and TO behind the associated oblique shock wave.arrow_forwardGas flowing through a diverging nozzle has at inlet section a temperature of 20 °C, pressure 120 kN/m? and velocity 300 m/s. At the outlet of the nozzle the velocity has fallen to 100 m/s. Assuming an adiabatic flow, what is the values of outlet pressure, temperature, internal energy and specific enthalpy at outlet section. Take y = 1.333 and Cv= 0.86 kJ/kg K.arrow_forward
- An aircraft where the ambient parameters are density = 0.785 kg/m^3, p= 0.50 atm, and velocity = 350 m/s. At a point on the airfoil surface, the pressure is 0.40 atm. Assuming isentropic flow, calculate the velocity at that point.arrow_forwardConsider a circular cylinder in a hypersonic flow, with its axisperpendicular to the flow. Let φ be the angle measured between radiidrawn to the leading edge (the stagnation point) and to any arbitrary pointon the cylinder. The pressure coefficient distribution along the cylindricalsurface is given by Cp = 2 cos2 φ for 0 ≤ φ ≤ π/2 and 3π/2 ≤ φ ≤ 2πand Cp = 0 for π/2 ≤ φ ≤ 3π/2. Calculate the drag coefficient for thecylinder, based on projected frontal area of the cylinder.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
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