pressure distribution on wing

Published on: **Mar 4, 2016**

Published in:
Engineering

Source: www.slideshare.net

- 1. Airfoil Pressure Distributions (From http://adg.stanford.edu/aa241/AircraftDesign.html) (also from Peery:Aircraft Structures) The aerodynamic performance of airfoil sections can be studied most easily by reference to the distribution of pressure over the airfoil. This distribution is usually expressed in terms of the pressure coefficient: Cp is the difference between local static pressure and freestream static pressure, nondimensionalized by the freestream dynamic pressure. What does an airfoil pressure distribution look like? We generally plot Cp vs. x/c. x/c varies from 0 at the leading edge to 1.0 at the trailing edge. Cp is plotted "upside-down" with negative values (suction), higher on the plot. (This is done so that the upper surface of a conventional lifting airfoil corresponds to the upper curve.) The Cp starts from about 1.0 at the stagnation point near the leading edge... It rises rapidly (pressure decreases) on both the upper and lower surfaces... ...and finally recovers to a small positive value of Cp near the trailing edge. Various parts of the pressure distribution are described in the following sections. Airfoil Pressure Distributions file:///home/kamle/Desktop/Old_HDD/courses_combin... 1 of 3 Saturday 08 August 2015 10:48 AM
- 2. Upper Surface The upper surface pressure is lower (plotted higher on the usual scale) than the lower surface Cp in this case. But it doesn't have to be. Lower Surface The lower surface sometimes carries a positive pressure, but at many design conditions is actually pulling the wing downward. In this case, some suction (negative Cp -> downward force on lower surface) is present near the midchord. Pressure Recovery This region of the pressure distribution is called the pressure recovery region. The pressure increases from its minimum value to the value at the trailing edge. This area is also known as the region of adverse pressure gradient. As discussed in other sections, the adverse pressure gradient is associated with boundary layer transition and possibly separation, if the gradient is too severe. Trailing Edge Pressure The pressure at the trailing edge is related to the airfoil thickness and shape near the trailing edge. For thick airfoils the pressure here is slightly positive (the velocity is a bit less than the freestream velocity). For infinitely thin sections Cp = 0 at the trailing edge. Large positive values of Cp at the trailing edge imply more severe adverse pressure gradients. CL and Cp The section lift coefficient is related to the Cp by: Cl = ∫ (Cpl - Cpu) dx/c (It is the area between the curves.) with Cpu = upper surface Cp Also Cl = section lift / (q c) Stagnation Point The stagnation point occurs near the leading edge. It is the place at which V = 0. Note that in incompressible flow Cp = 1.0 at this point. In compressible flow it may be somewhat larger. Airfoil Pressure Distributions file:///home/kamle/Desktop/Old_HDD/courses_combin... 2 of 3 Saturday 08 August 2015 10:48 AM
- 3. Distribution of pressure distribution over airfoils Airfoil Pressure Distributions file:///home/kamle/Desktop/Old_HDD/courses_combin... 3 of 3 Saturday 08 August 2015 10:48 AM