Does someone know the C4 airfoil data or the eqautions to calculate its coordinates? thanks !

dragon

I've found RC(4) airfoil on this page: http://www.nasg.com/afdb/show-airfoil-e.phtml?id=996 Click on contour link and y'll find cordinates.

thanks,dear Bounecer.

But I guess RC(4) is not the one I want. I read in the paper that the airfoil's name is C4. I dont know if it is NACA,but I guess not. This is the paper I read: "Non-Newtonian Flow over the trailing edge of an airfoil" Journal name:Experimental thermal and fluid science 1996;12:244-249 there is a picture of C4, but it does not look like RC(4).

thanks again.

dragon

I'm studing a 2D, incompressible, laminar, viscous flow about airfoil. I've used "structured C grid generation", with finite volumes method and second order upwind numerical scheem and SIMPLER algorithm. After great time and efforts, my code program diverge after i reach a precision of 10e-3.

In my work i'am using only NACA airfoil, but if i found what you look for, i'll contact you.

Good chance friend

please send all information about airfoil HN273. thank you.

[arungeth]

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[Ammofreak]

Here is the code to generate C4 thickness distribution and cambered airfoil

clc
clear all

t = 0.1;%thickness, t and x are in % of chord length
theta = 18.5*pi/180; %camber angle/2
mu = pi/6;%it is actually dy_c/dx, for simplicity i used it equal to 30 degrees.

x = 0:0.02:1;

y = (t/0.2)*(0.3048*x.^0.5-0.0914*x-0.8614*x.^2+2.1236*x.^3-2.9163*x.^4+1.9744*x.^5-0.5231*x.^6);%equation for symmetric airfoil C4

y_c = sqrt(((0.5/sin(theta/2)).^2-(x-0.5).^2))-0.5/tan(theta/2);%equation of camber line

%for upper s/f wrapped co-ordinates around camber line y_c

x_u = transpose(115*(x-y*sin(mu)));
y_u = transpose(115*(y_c+y*cos(mu)));

%for lower s/f wrapped co-ordinates around camber line y_c

x_L = transpose(115*(x-y*cos(mu)));
y_L = transpose(115*(y_c-y*sin(mu)));

plot(x_u,y_u,x_L,y_L);
ylim([-20 50]);

regards,
Ammofreak