Consider an airfoil with chord length $c$ and running distance $x$ measured along the chord. The leading edge is located at $(\frac{x}{c})=0$ and the trailing edge at $(\frac{x}{c})=1.$ The pressure coefficient variations over the upper and lower surfaces are given, respectively, as

$C_{p,u}=1-300(\frac{x}{c}{)}^2$ for $0(\frac{x}{c})0\mathrm{.}1$

$C_{p,u}=-2\mathrm{.}2277+2\mathrm{.}2777(\frac{x}{c})$ for $0\mathrm{.}1(\frac{x}{c})1\mathrm{.}0$

$C_{p,l}=1-0\mathrm{.}95(\frac{x}{c})$ for $0(\frac{x}{c})1\mathrm{.}0$

Chula ISE

pg$.\frac{2}{2}$

a$.$ Using MATLAB, plot the pressure distribution

b$.$ Calculate the normal force coefficient

c$.$ Calculate the location of center of pressure

d$.$ Calculate the moment coefficient about the leading edge