I'm trying to code a method to find the lift coefficient of a NACA airfoil using panel method.(adsbygoogle = window.adsbygoogle || []).push({});

(Coding mentioned below is MATLAB)

There are two things I am stuck at:

1) Finding the coordinates of NACA

Usually we use the given general formula for camber. However as we all know, if we compute the equation, the "actual" calculated coordinates do not meet at the trailing edge.

In other words, if the camber is 1, the trailing edge is located at slightly beyond 1 and the y-axis is smaller than 0.

Would I have to re-scale the output airfoil so that the trailing edge is located at x=1? If so, should I assume that the trailing edge is less than 0 for more accurate analysis?

So is there any method to accurately compute out the coordinates of NACA airfoil?

2) Finding the C_{l}from C_{p}

If I divide the airfoil as following:

12 panels. Starting point from trailing edge, then through the lower edge, then the leading edge, then the upper edge and the 13th point being back at point 1.

Then I obtained the values:

I X Y THETA S GAMA V CP

1.0000 0.9665 -0.0025 -3.0656 0.0672 0.0145 -0.8445 0.2868

2.0000 0.8415 -0.0115 -3.0713 0.1835 -0.1411 -0.9612 0.0761

3.0000 0.6250 -0.0257 -3.0800 0.2505 -0.1548 -1.0071 -0.0143

4.0000 0.3750 -0.0378 -3.1065 0.2502 -0.1635 -1.0525 -0.1077

5.0000 0.1585 -0.0380 3.0962 0.1832 -0.1720 -1.1071 -0.2256

6.0000 0.0335 -0.0169 2.6735 0.0751 -0.1850 -0.8112 0.3420

7.0000 0.0335 0.0231 0.6032 0.0813 -0.0042 0.8040 0.3536

8.0000 0.1585 0.0613 0.1647 0.1855 0.1947 1.2136 -0.4729

9.0000 0.3750 0.0744 -0.0170 0.2500 0.1989 1.2014 -0.4434

10.0000 0.6250 0.0583 -0.1112 0.2516 0.1873 1.1309 -0.2789

11.0000 0.8415 0.0290 -0.1670 0.1856 0.1722 1.0423 -0.0863

12.0000 0.9665 0.0068 -0.1993 0.0683 0.1505 0.8674 0.2477

13.0000 0 0 0 0 -0.0145 0 0

how would I find the lift coefficient?

For the Area, do I need the area of the airfoil or the area surrounded by panels?

This is my MATLAB code (gives highly strange values...)

The angle of attack was 0 degrees

x,y,theta,s... are all vectors

x= x coordinates of panels

y= y coordinates of panels

theta=angles in radians

s= panel lengths

v= dimensionless velocity at control panel

cp= pressure coefficient at control panel

ρ,μ,v were arbitrary values I gave.

The lift comes out as 3.2365, Although from what I know that it should be around 0.2

Where did go wrong? What should I have done? I'm completely lost...Code (Text):l=sum(cp.*s.*sin(theta)); %Lift

ro=1.225;

A=0;

for jjj=1:m

if jjj==1

A=A+abs((x(jjj)-x(end))*(y(jjj)+y(end)));

else

A=A+abs((x(jjj)-x(jjj-1))*(y(jjj)+y(jjj-1)));

end

end

A=A/2;

c=1; %camber length

mu=1.78*10^-5;

Re=25000;

v=Re*mu/(ro*c);

cl=l/(0.5*ro*v^2*A)

**Physics Forums | Science Articles, Homework Help, Discussion**

Dismiss Notice

Join Physics Forums Today!

The friendliest, high quality science and math community on the planet! Everyone who loves science is here!

The friendliest, high quality science and math community on the planet! Everyone who loves science is here!

# Code for finding the lift coefficient of NACA

Can you offer guidance or do you also need help?

Draft saved
Draft deleted

**Physics Forums | Science Articles, Homework Help, Discussion**