6/6/2023 0 Comments Clark y airfoil![]() Rotate the turntable so that the angle of attack is 0°.The manometer panel should have 24 columns filled with colored oil and marked with water inch graduations. The ports on the Clark Y-14 model are located as follows: port 1: x/c = 0 (right on leading edge), ports 2 and 11: x/c = 5%, ports 3 and 12: x/c = 10%, ports 4 and 13: x/c = 20%, ports 5 and 14: x/c = 30% ports 6 and 15: x/c = 40%, ports 7 and 16: x/c = 50%, ports 8 and 17: x/c = 60%, ports 9 and 18: x/c = 70%, and ports 10 and 19: x/c = 80% (Figure 2). Connect the 19 pressure tubes labeled 1 - 19 to the corresponding ports of the manometer panel, respectively.Note that the model is touching both the floor and the ceiling of the wind tunnel test section so no 3D flow around the airfoil develops. Mount the aluminum Clark Y-14 model on the turntable inside the test section so that port #1 is facing upstream.The test section should be 1 ft x 1 ft and the wind tunnel should be able to sustain a maximum airspeed of 140 mph. Remove top cover of test section to install the Clark Y-14 model (chord length, c = 3.5 in).Where Δ x i is the increment between 2 adjacent ports. Once the pressure distribution is obtained, the non-dimensional lift coefficient, C l, can be numerically determined to evaluate Equation 3: Where Δ h is the height difference of the manometer with reference to free-stream pressure, ρ L is the density of the liquid in the manometer, and g is the acceleration due to gravity. The gage pressure reading is determined using the following equation: The gage pressures are measured using a manometer panel with 24 columns filled with liquid oil marked with water inch graduations. Airfoil profile of a Clark Y-14 wing with locations of gage pressure ports. It has a thickness of 14% and is flat on the lower surface from 30% of the chord length to the back.įigure 2. ![]() A Clark Y-14 airfoil is shown in Figure 2. Here, the overall pressure distribution along the airfoil is measured with 19 small tubes embedded in the wing and attached to a pressure transducer. Where the new parameter μ is the dynamics viscosity of the fluid. Where x is the horizontal coordinate position with origin starting from leading edge.Īirfoil performance takes into account the Reynolds number, Re, which is defined as: As such, at small angles of attack, the lift coefficient can be estimated by: Where L' is the lift per unit span, and c is the chord length of the airfoil.Įxcept for points along the leading-edge, the pressure forces uniformly point upward, in approximately the same direction as lift. ![]() The non-dimensional lift coefficient C l is similarly defined: Where P is the absolute pressure, P ∞ is the undisturbed free-stream pressure, P gage = P − P ∞ is the gage pressure, and is the dynamic pressure, which is based on the free-stream density, ρ ∞, and airspeed, V ∞. The non-dimensional pressure coefficient, C p, for an arbitrary point on the airfoil is defined as: ![]() Figure 1 shows a schematic of the pressure distribution over an airfoil.įigure 1. If the shear forces parallel to the surface of the airfoil are neglected (typically their contributions to lift are small), then the total pressure force is the reason for the lift generated by the airfoil. An airfoil develops lift at various angles of attack through lower gage pressures on the upper surface and higher gage pressures on the lower surface with respect to the pressure of the approaching air (free-stream pressure).
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