Consider a low-speed open-circuit subsonic wind tunnel with an inlet-to-throat area ratio of 12. 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
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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.arrow_forwardConsider a wing mounted in the test-section of a subsonic wind tunnel. The velocity of the airflow is 160 ft/s. If the velocity at a point on the wing is 195 ft/s, what is the pressure coefficient at this point?arrow_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_forward
- An airplane is flying at a pressure altitude of 15 km with a velocity of 619 m/s. The outside air temperature is 220 K. What is the pressure measured by a Pitot tube mounted on the nose of the airplane?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_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
- Suppose an aircraft is flying at standard sea-level at M = 0.8 and using a pitot-static tube for airspeed measurement. Determine the actual difference in total and static pressures as would be measured by the pitot-static system. Compare the speed of the aircraft as determined using the actual pressure difference and incompressible flow versus that knowing the actual Mach number. Use k = 1.4.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_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_forward
- The 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_forwardHow to Calibration of Air Velocity from Pressure Using Sub-Sonic Wind Tunnel and Write the components of Sub-Sonic Wind Tunnel?arrow_forward85. Calculate the altitude of a supersonic airplane (v = 800 m-s1, T=-34 C°) when the time between seeing above observer and hearing is 9.9 sec.arrow_forward
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