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• ## A circular cylinder and a sphere are in a free stream flow with their axis perpendicular to the flow. At the top of the sphere there is a pressure tap which is connected by a tube to one side of a manometer. There is a pressure tap on the surface of the cylinder which is connected to the other side of the manometer. Find the location of this tap on cylindrical surface such that there is no deflection of fluid in the manometer.

on 1st September 2021 in Aerodynamics.
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• ## A pitot tube of an airplane which is flying at an standard sea level measures a pressure of $$1.09 \times 10^{5} N/m^{2}$$. Find the velocity of the airplane.

on 16th June 2021 in Aerodynamics.
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• ## Find the velocity of air in the test section of a low speed open circuit subsonic wind tunnel which has an inlet -to-throat area ratio of $$10$$. The U-tube mercury manometer read as a height difference of $$10.5\,cm$$ for the pressure difference between the inlet and the test section of the wind tunnel.

on 15th June 2021 in Aeronautics.
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• ## Find the velocity of an airplane flying at a standard sea level, if the static pressure at the throat of venturi is $$100 000 N/m^{2}$$.

on 15th June 2021 in Aerodynamics.
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• ## Bernoulli’s Principle and Equation:

on 8th December 2020 in Fluid dynamics.
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• ## Consider an incompressible two-dimensional viscous flow over a curved surface. Let the pressure distribution on the surface be $$p(s)=2+sin\left ( \frac{\pi}{2}+s \right )\;N/m^2$$,where $$s$$ is the distance along the curved surface from the leading edge. Find the distance from where the flow separates.

on 30th December 2019 in Aerodynamics.
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• ## The measured lift slope for the NACA $$23012$$ airfoil is $$0.1080\; degree^{−1}$$, and $$α_{L}=0 = −1.3^{\circ}$$. Consider a finite wing using this airfoil, with AR = $$8$$ and taper ratio = $$0.8$$. Assume that $$δ = τ$$ . Calculate the lift and induced drag coefficients for this wing at a geometric angle of attack = $$7^{\circ}$$.

on 12th November 2019 in Flight mechanics.
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• ## Using thin airfoil theory, calculate $$(a) α_{ L=0}$$ (b) $$cl$$ when $$α = 3^{\circ}$$.

on 11th November 2019 in Aerodynamics.
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• ## For the NACA $$2412$$ airfoil, the lift coefficient and moment coefficient about the quarter-chord at $$−6 ^{\circ}$$ angle of attack are $$−0.39$$ and $$−0.045$$,respectively. At $$4^{\circ}$$ angle of attack, these coefficients are $$0.65$$ and $$−0.037$$,respectively. Calculate the location of the aerodynamic center.

on 10th November 2019 in Aerodynamics.
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• ## Consider a thin, symmetric airfoil at $$1.5^{\circ}$$ angle of attack. From the results of thin airfoil theory, calculate the lift coefficient and the moment coefficient about the leading edge.

on 9th November 2019 in Aerodynamics.
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