I thought I'd try and make a simple example to give an easy understanding on this subject, and I wanted to use some images as well to illustrate and make it just a little more inspirational than just a lot of words.
This example is made for everyone to understand, so please don't feel offended if the information seems too obvious for you, it might not be that obvious to someone else.
I have made two examples using the basic NACA airfoils, the first is a lifting wing (sport/aerobatic) and the second is a symmetrical wing (aerobatic).
I thought the wing example should be a relatively small model airplane compared to today's standards, because it will always work if you make it bigger - but not always if you make it smaller.
I have made a simple plan form view of a wing half that have a root chord at 10inches, a tip chord of 6inches and a full span at 60inches. The span is just a reference and could be about anything, it would not make a difference in this example. The relation between the size of the root and tip chord is what matters here.
In the first example I have used 1412 as a root airfoil and 2312 as a tip airfoil, the 1412 have a chamber at 1% @ 40% of the chord and have a 12% thickness, and the 2312 have a chamber at 2% @ 30% of the chord and have a 12% thickness.
In the second example I have used 0012 as a root airfoil and 0014 as a tip airfoil, the 0012 have a chamber at 0% @ 0% of the chord and have a 12% thickness, and the 0014 have a chamber at 0% @ 0% of the chord and have a 14% thickness.
I have created the NACA airfoils at this site using 80 points:
http://airfoiltools.com/airfoil/naca4digit
Here is an information page on the NACA airfoils:
http://www.aerospaceweb.org/question...ls/q0041.shtml
To calculate how the size of the wing chord affects the airfoil characteristics we use Reynolds Numbers, I have calculated this with an anticipated air speed at about 26mph.
10 inches at this airspeed is about Re 200.000 and 6 inches at this air speed is about Re 120.000, you can find simple to use Reynolds Numbers calculators online if you search for it - there are several different alternatives.
If the wing is scaled up lets say by 25% then the Re numbers at the root would go up from Re200k to Re250k, and so at Re200k this scaled up wing would be at a slower airspeed at about 19,5mph with the same results.
The left of the polar data images shows Cl (lift or Lift Coefficient) in relation to alpha (AoA or Angle of Attack), it shows that when the airfoil reaches a certain AoA the lift starts to drop dramatically - this indicates stall.
You can see that in both my examples these two curves is somewhat matched, the two airfoils drops at about the same AoA even though they have different characteristics and is analyzed at different Re numbers.
If you do this with the same airfoil at the tip and root on a tapered wing the tip will show a drop at an earlier AoA than the root, this indicates that the wing will be prone to tip stall when it reaches it's limits.
The polar data is created by a program called XFoil by professor Mark Drela, this is free to download and use but it might be a little complicated to understand at first glance. I have used something called "Profili" which is basically an easier and simpler interface for the XFoil program, Profili is free to download and use with some restrictions.