How do I choose the correct airfoil?

I've written an add-in for Fusion360 which will probably help you: it collects your proposed operating conditions, then inserts for you the best airfoil shape that it knows, out of a library of 768 base foils each of which has been optimised through thousands of hours of CPU time using particle swarms.

You can read a bit more about the history in my autodesk forum post here: https://forums.autodesk.com/t5/fusion-360-api-and-scripts/add-in-announcement-hydrofoil-and-airfoil-tools-seeking-your/m-p/9453985/highlight/false

Here's the add-in page: https://apps.autodesk.com/FUSION/en/Detail/Index?id=5447707798035545266&appLang=en&os=Win64

p.s. NACA shapes are never the right answer - that idea pre-dates computers, and if you read their latest application notes, they tell you not to use their shapes because modern computational methods are always superior. By way of example - I analyzed all 9999 series 4digit NACA foils, and several thousand 5 digit ones (200CPU cores took 3 months), then selected the best-performing CL/CD for each of the following Reynolds numbers. The margin by which CFD-optimised shapes outperform the BEST possible NACA shapes is the second column.

Re          CFD outperformance
300.00              146%
400.00              155%
500.00              148%
750.00              146%
1,000.00            150%
1,250.00            178%
1,500.00            182%
2,000.00            175%
2,500.00            172%
3,000.00            172%
3,500.00            175%
4,500.00            180%
5,500.00            111%
7,000.00            150%
8,192.00            136%
10,000.00           139%
12,500.00           148%
15,000.00           145%
20,000.00           145%
25,000.00           147%
30,000.00           142%
35,000.00           139%
45,000.00           136%
55,000.00           133%
65,536.00           129%
80,000.00           125%
100,000.00          131%
125,000.00          139%
150,000.00          143%
200,000.00          150%
275,000.00          155%
350,000.00          153%
400,000.00          149%
500,000.00          152%
620,000.00          158%
750,000.00          169%
1,000,000.00        184%
1,400,000.00        208%
1,700,000.00        225%
2,000,000.00        224%
2,500,000.00        203%
3,000,000.00        179%
4,000,000.00        184%
5,000,000.00        178%
6,000,000.00        174%
7,000,000.00        160%
8,192,000.00        142%
10,000,000.00       142%
13,000,000.00       132%
16,000,000.00       114%
20,000,000.00       134%
25,000,000.00       123%
30,000,000.00       133%
36,000,000.00       201%
45,000,000.00       127%
65,536,000.00       249%
81,920,000.00       154%
100,000,000.00      131%

200% means that the CFD-optimised shape Cl/Cd is double the NACA Cl/Cd

I am of the idea that @Peter Kämpf posted, and would have liked to have a modification of my own to the http://airfoiltools.com/airfoil/details?airfoil=nlf0215f-il NASA NFL-(1)-0215 airfoil that he recommended. I saw it to be a good idea to Put a larger airfoil section on the base and taper it to a smaller verion of itself (and in fact the only problem i have with this is the tail at the end of the airfoil could get too small if i tended to go for a delta win formation, so that is out of question. A delta wing would be too hard to material-engineer the end of the wing. ) so the best i thought of was it to taper it not so much, and increase the wing area at the base a bit. Would have been interesting as it would have made for a stronger structure to hold my weight, id like to keep stuff light.

The problem i have with the airfoild with an xfoil calculation is that, i pretty much made a Aoa-Cl-Cd table for angles, but xofil predicted the sepparation zone being a bit lowere in the airfoil at higher Aoa's * but not much * , 10 DEGREES.

The Simulation is made in CFD at a decent mesh size at quite a few iterations.

Is the xfoil table, for engineering purposes, accurate to a degree?

Also i thought of making a small modification to the airfoil, and add a bitsy more camber at the 60-65% of the camber line length. And see if i can stall the sepparation with 3 decently sized vortex generators. Also ill do some flaps simulations aswell.

A few of the obvious factors to consider in choosing an airfoil include:

1. intended speed range
2. presumed pilot's skill level
3. intended maximum weight
4. size of aircraft
5. Your skill in fabrication

From the intended maximum weight and intended speed range, you can get some idea of the wing loading you want. Given your intended (relatively low) minimum takeoff speed, you need fairly low wing loading.

With some idea of the wing area, you can start to think about how to achieve that: greater wing span vs. greater chord. As a simple guideline, something that has a really "clean", low-drag design will tend to benefit from a high-aspect ratio planform. Conversely, a higher-drag design will work better with a lower aspect ratio of wing. To give a couple of extreme examples, competition sail planes often have an aspect ratio around 50:1. A piper cub had an aspect ratio of 7:1, and hang gliders are often around 1:1.

With the chord and operating speed chosen, we can compute the range of Reynolds numbers we care about.

With that, we can also start to think about pilot skill level. For example, let's consider the NACA airfoil you suggested. With a minimum of 20 m/s and a 0.1 meter chord, we'd get a Reynolds number around 50K. Looking at the Cl vs. alpha curve for this airfoil at 50K:

For the moment, the part we care most about is the right side of the curve. Note how it curves upward, then has a very sharp peak and drops again very quickly.

That's telling us that (at this Reynolds number) the airfoil has a really nasty stall--as we move to higher angles of attack, it produces more lift--then with just slightly greater angle of attack, it stalls and the lift drops like a rock. In this range of Reynolds numbers, this could be pretty tricky to fly, even for quite an experienced pilot.

By contrast, if you can be reasonably certain your Reynolds number will be at least a million, the picture changes quite a bit:

In this case, the stall is much more gentle--in fact, the lift isn't dropping at all. We have to look at a separate plot to see that while the lift isn't dropping, the drag is rising sharply, so our L/D ratio (and with it, our glide path) drops very quickly:

Based on these, we can guess that this airfoil was probably really designed for higher Reynolds numbers. At 50K it's basically evil. At 1M it's more manageable, but still kind of a handful--especially since, with such high drag, when you stall your air speed (and with it the Reynolds number) will start to drop very quickly, so unless you react to recover from the stall quickly, you're going to drop into the range where you're going to lose lift as well.

Let's give a little thought to that final factor: fabrication skill. Here what we're most interested in is the accuracy with which you're likely to build the airfoil. Some airfoils are really picky about following their profile precisely to get good results. For example, quite a few of Selig's designs have a subtle "bubble ramp" along the upper surface. If you get this even a little wrong, you can end up with quite a bit higher drag than intended.

Other airfoils have quite a bit higher tolerance, so minor deviations from the intended profile won't affect their performance nearly as much. If you aren't very experienced, one of these may be a better choice.