• Question Tag: incompressible flow
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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.

A circular cylinder and a sphere are in a free stream flow with their axis perpendicular to the flow. At …

Kisan Kumar Asked on 1st September 2021 in
• 921 views

A pitot tube of an airplane which is flying at an standard sea level measures a pressure of $$1.09 \times … techAir Asked on 16th June 2021 in • 914 views • 1 answers • 0 votes • ## 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.

Find the velocity of air in the test section of a low speed open circuit subsonic wind tunnel which has …

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

Find the velocity of an airplane flying at a standard sea level, if the static pressure at the throat of …

Kumar59 Asked on 15th June 2021 in
• 951 views
• ## Bernoulli’s Principle and Equation:

What is Bernoulli’s theorem?

techAir Asked on 8th December 2020 in
• 1K views

Consider an incompressible two-dimensional viscous flow over a curved surface. Let the pressure distribution on the surface be $$p(s)=2+sin\left ( … Kisan Kumar Asked on 30th December 2019 in • 1K views • 1 answers • 0 votes • ## 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}$$.

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 …

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

The NACA 4412 airfoil has a mean camber line given by\[\frac{z}{c} = \left\{ {\begin{array}{*{20}{c}} {0.25\left[ {0.8\frac{x}{c} – {{\left( {\frac{x}{c}} \right)}^2}} …

Worldtech Asked on 11th November 2019 in
• 4K views
• ## 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.

For the NACA $$2412$$ airfoil, the lift coefficient and moment coefficient about the quarter-chord at $$−6 ^{\circ}$$ angle of attack …

Worldtech Asked on 10th November 2019 in
• 2K views
• ## 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.
Consider a thin, symmetric airfoil at $$1.5^{\circ}$$ angle of attack. From the results of thin airfoil theory, calculate the lift …