Talking Airfoils

[bomber]

The Beagle Pup uses a NACA 63-615 airfoil from the root to tip... how do we know, go here

http://m-selig.ae.illinois.edu/ads/aircraft.html

google the airfoil and you'll get to here...

http://airfoiltools.com/airfoil/details ... a632615-il

and the data file gives us the polars..

NACA 63(2)-615
1.00000 0.00000
0.95042 0.01245
0.90089 0.02398
0.85127 0.03555
0.80153 0.04693
0.75163 0.05800
0.70159 0.06847
0.65139 0.07809
0.60105 0.08665
0.55058 0.09393
0.50000 0.09974
0.44932 0.10384
0.39857 0.10598
0.34778 0.10587
0.29700 0.10331
0.24625 0.09830
0.19558 0.09066
0.14504 0.08010
0.09473 0.06578
0.06973 0.05667
0.04492 0.04560
0.02050 0.03129
0.00866 0.02159
0.00418 0.01634
0.00205 0.01317
0.00000 0.00000
0.00795 -0.01017
0.01082 -0.01214
0.01634 -0.01517
0.02950 -0.02013
0.05508 -0.02664
0.08027 -0.03123
0.10527 -0.03476
0.15496 -0.03972
0.20442 -0.04290
0.25375 -0.04460
0.30300 -0.04499
0.35222 -0.04407
0.40143 -0.04172
0.45068 -0.03814
0.50000 -0.03356
0.54942 -0.02823
0.59895 -0.02239
0.64861 -0.01629
0.69841 -0.01015
0.74837 -0.00430
0.79847 0.00083
0.84873 0.00483
0.89911 0.00704
0.94958 0.00651
1.00000 0.00000

How easy was that....

[bomber]

Thing is you can't use the polars directly in JSBsim you have to use the coeeficients.... so it's off to Javafoil

http://www.mh-aerotools.de/airfoils/javafoil.htm

=================================================

On the geometry tab paste the polars into the big table... and press update view (you're not interested in anything else on that tab)

[bomber]

next challenge is a bit of spreadsheet jockeying to determine a realistic reynolds number

m/s = 26.4 = cruise speed
l = 1.1965 = cord at 50% wing span
p = 70000 = rho (a genuine magic number)

reynold no. = m/s * l * P = 2211132

[bomber]

So in the polar tab paste in the reynolds number into each of the 3 reynolds number boxes..

first angle of attack = -40
last angle of attack = 40

press analyse it...

within the polar tab open the 2211132 tab and there's ya co-efficients of lift n drag.

Code: Select all

Name = NACA 63-615
Mach = 0; Re = 2211132; T.U. = 1.0; T.L. = 1.0
Surface Finish = 0; Stall model = 0; Transition model = 1; Aspect Ratio = 0; ground effect = 0
α   Cl   Cd   Cm 0.25   T.U.   T.L.   S.U.   S.L.   L/D   A.C.   C.P.
[°]   [-]   [-]   [-]   [-]   [-]   [-]   [-]   [-]   [-]   [-]
-40.0   -0.201   0.61336   -0.033   1.000   0.003   1.000   0.030   -0.328   0.255   0.088
-39.0   -0.212   0.58364   -0.032   1.000   0.003   1.000   0.029   -0.363   0.253   0.097
-38.0   -0.224   0.57054   -0.032   1.000   0.002   1.000   0.029   -0.393   0.247   0.105
-37.0   -0.237   0.55342   -0.033   1.000   0.003   1.000   0.029   -0.428   0.247   0.113
-36.0   -0.251   0.51252   -0.033   1.000   0.003   1.000   0.029   -0.490   0.247   0.120
-35.0   -0.266   0.49762   -0.033   1.000   0.003   1.000   0.028   -0.535   0.245   0.127
-34.0   -0.283   0.47441   -0.033   1.000   0.002   1.000   0.028   -0.596   0.243   0.134
-33.0   -0.301   0.44815   -0.033   1.000   0.002   1.000   0.027   -0.671   0.239   0.141
-32.0   -0.320   0.42268   -0.033   1.000   0.002   1.000   0.027   -0.757   0.238   0.147
-31.0   -0.341   0.40120   -0.033   1.000   0.002   1.000   0.027   -0.850   0.228   0.152
-30.0   -0.364   0.36660   -0.034   1.000   0.002   1.000   0.028   -0.993   0.227   0.156
-29.0   -0.388   0.35032   -0.034   1.000   0.002   1.000   0.028   -1.109   0.230   0.161
-28.0   -0.415   0.31948   -0.035   1.000   0.003   1.000   0.029   -1.298   0.230   0.165
-27.0   -0.443   0.30301   -0.036   1.000   0.003   1.000   0.029   -1.461   0.234   0.170
-26.0   -0.472   0.27922   -0.036   1.000   0.002   1.000   0.028   -1.692   0.227   0.174
-25.0   -0.504   0.25859   -0.037   1.000   0.003   1.000   0.030   -1.948   0.228   0.177
-24.0   -0.536   0.24018   -0.037   0.941   0.003   1.000   0.029   -2.232   0.232   0.180
-23.0   -0.569   0.21900   -0.038   0.931   0.003   1.000   0.029   -2.598   0.213   0.183
-22.0   -0.602   0.20163   -0.040   0.931   0.003   1.000   0.034   -2.988   0.205   0.184
-21.0   -0.635   0.18301   -0.041   0.921   0.004   1.000   0.036   -3.467   0.208   0.185
-20.0   -0.664   0.16716   -0.042   0.911   0.004   1.000   0.038   -3.974   0.194   0.186
-19.0   -0.690   0.15024   -0.044   0.891   0.005   1.000   0.043   -4.595   0.176   0.186
-18.0   -0.711   0.13479   -0.046   0.868   0.005   1.000   0.047   -5.272   0.125   0.185
-17.0   -0.724   0.12021   -0.048   0.840   0.006   1.000   0.055   -6.020   -0.101   0.183
-16.0   -0.729   0.10391   -0.052   0.813   0.008   1.000   0.074   -7.014   -1.255   0.178
-15.0   -0.732   0.08176   -0.062   0.779   0.008   1.000   0.151   -8.959   -1.021   0.166
-14.0   -0.747   0.04825   -0.075   0.763   0.008   1.000   0.373   -15.482   -1.108   0.149
-13.0   -0.750   0.02379   -0.085   0.735   0.009   1.000   0.636   -31.536   0.572   0.136
-12.0   -0.704   0.01778   -0.089   0.713   0.009   1.000   0.709   -39.574   0.309   0.123
-11.0   -0.637   0.01514   -0.092   0.690   0.009   1.000   0.755   -42.063   0.289   0.105
-10.0   -0.553   0.01276   -0.095   0.668   0.010   1.000   0.797   -43.317   0.282   0.078
-9.0   -0.456   0.01111   -0.098   0.646   0.010   1.000   0.831   -41.055   0.278   0.035
-8.0   -0.352   0.00993   -0.101   0.623   0.010   1.000   0.861   -35.421   0.275   -0.036
-7.0   -0.241   0.00904   -0.103   0.598   0.011   1.000   0.888   -26.681   0.274   -0.179
-6.0   -0.126   0.00828   -0.106   0.573   0.020   1.000   0.908   -15.208   0.272   -0.592
-5.0   -0.007   0.00784   -0.108   0.552   0.025   1.000   0.919   -0.930   0.273   -14.628
-4.0   0.114   0.00552   -0.111   0.526   0.332   1.000   0.952   20.620   0.272   1.230
-3.0   0.235   0.00542   -0.114   0.507   0.370   1.000   0.953   43.461   0.269   0.733
-2.0   0.357   0.00551   -0.116   0.485   0.394   1.000   0.953   64.793   0.269   0.575
-1.0   0.478   0.00562   -0.118   0.465   0.416   1.000   0.954   85.079   0.269   0.497
0.0   0.604   0.00570   -0.121   0.448   0.437   1.000   0.955   106.041   0.269   0.450
1.0   0.726   0.00595   -0.123   0.426   0.459   1.000   0.955   121.961   0.269   0.419
2.0   0.847   0.00623   -0.125   0.409   0.478   1.000   0.955   136.004   0.269   0.398
3.0   0.967   0.00678   -0.128   0.389   0.503   1.000   0.956   142.717   0.270   0.382
4.0   1.085   0.00698   -0.130   0.366   0.524   1.000   0.956   155.579   0.271   0.370
5.0   1.197   0.00774   -0.132   0.340   0.551   1.000   0.957   154.703   0.272   0.361
6.0   1.302   0.01394   -0.135   0.004   0.573   1.000   0.957   93.390   0.273   0.353
7.0   1.398   0.01506   -0.137   0.004   0.602   1.000   0.957   92.835   0.268   0.348
8.0   1.464   0.01631   -0.138   0.004   0.627   0.934   0.962   89.748   0.266   0.344
9.0   1.525   0.01797   -0.139   0.003   0.654   0.904   0.962   84.848   0.277   0.341
10.0   1.573   0.01986   -0.141   0.003   0.676   0.881   0.962   79.197   0.289   0.339
11.0   1.604   0.02219   -0.142   0.003   0.698   0.857   0.962   72.269   0.316   0.339
12.0   1.616   0.02470   -0.143   0.003   0.721   0.832   0.963   65.452   0.627   0.339
13.0   1.610   0.02803   -0.145   0.003   0.743   0.801   0.963   57.434   0.175   0.340
14.0   1.588   0.03175   -0.146   0.003   0.771   0.771   0.963   50.015   0.214   0.342
15.0   1.552   0.03571   -0.147   0.003   0.798   0.743   0.963   43.475   0.231   0.344
16.0   1.502   0.04099   -0.147   0.003   0.821   0.708   0.962   36.655   0.242   0.348
17.0   1.441   0.04790   -0.148   0.003   0.859   0.666   0.965   30.080   0.250   0.352
18.0   1.372   0.05585   -0.147   0.003   0.879   0.624   0.955   24.570   0.257   0.357
19.0   1.296   0.06654   -0.147   0.003   0.909   0.571   0.952   19.482   0.268   0.363
20.0   1.213   0.08267   -0.144   0.003   0.912   0.499   0.955   14.675   0.279   0.369
21.0   1.131   0.10160   -0.142   0.003   0.914   0.428   0.954   11.135   0.299   0.375
22.0   1.045   0.13006   -0.136   0.003   0.917   0.330   0.954   8.033   0.330   0.380
23.0   0.961   0.16527   -0.128   0.003   0.917   0.230   0.955   5.817   0.356   0.383
24.0   0.888   0.20136   -0.120   0.003   0.919   0.154   0.954   4.409   0.334   0.385
25.0   0.829   0.22374   -0.117   0.003   0.919   0.132   0.955   3.707   0.297   0.391
26.0   0.768   0.24932   -0.114   0.003   0.912   0.109   0.955   3.080   0.298   0.398
27.0   0.704   0.27300   -0.111   0.003   0.912   0.091   0.955   2.580   0.282   0.408
28.0   0.648   0.29783   -0.110   0.003   0.914   0.084   0.954   2.175   0.284   0.420
29.0   0.595   0.32562   -0.107   0.003   0.914   0.070   0.955   1.828   0.285   0.430
30.0   0.549   0.34941   -0.107   0.003   0.914   0.065   0.955   1.571   0.268   0.444
31.0   0.507   0.37307   -0.106   0.003   0.917   0.060   0.955   1.358   0.275   0.459
32.0   0.468   0.39779   -0.105   0.003   0.927   0.055   0.952   1.177   0.268   0.473
33.0   0.434   0.42892   -0.104   0.003   0.927   0.053   0.952   1.011   0.270   0.491
34.0   0.402   0.45335   -0.103   0.003   0.927   0.048   0.955   0.887   0.289   0.507
35.0   0.374   0.48127   -0.102   0.003   0.917   0.044   0.956   0.777   0.263   0.523
36.0   0.348   0.50975   -0.103   0.003   0.919   0.044   0.955   0.683   0.244   0.545
37.0   0.325   0.53866   -0.102   0.003   0.919   0.042   0.956   0.603   0.238   0.565
38.0   0.304   0.56446   -0.103   0.002   0.929   0.042   0.955   0.538   0.240   0.589
39.0   0.285   0.61042   -0.103   0.002   0.929   0.041   0.955   0.466   0.224   0.611
40.0   0.267   0.62994   -0.104   0.002   0.919   0.042   0.957   0.424   0.170   0.639


so all we're after is the first 3 collumns

[bomber]

so that's the co-efficients for the airfoils as they stand..... but parts of our wing has control surfaces

if we go back to the type test for the beagle we find this

Control Surface Movements
Wing

flap
Up 0
Down 39.5

Aileron
Up 28
Down 12

Elevator
Up 30
Down 20

Elevator trim tab
Up 13
Down 32

Rudder
Left 25
Right 25

now the percentage of aileron to cord is 27%

================================

on the modify tab enter 27 in the flap cord xf/c box and -28 in the flap deflection press enter

---------------------------------------------

on the polar tab tick the add to plots box and analyise it.

---------------------------------------------
back to the modify tab press undo, enter 12 into the flap deflection box and press enter...
---------------------------------------------
on the polar tab analyse it

----------------------------------------------

We can now using these values of coefficients create a table within JSBsim to give accurate values for the airfoils with or without control deflection.

[jwocky]

Wow, you go the extra mile and them some more. That's beautiful from an engineering point of view.

[bomber]

It's one of the joys of flight modelling in real time... you don't have to follow convention.

[jwocky]

For me, it is trying to get those birds in the air reasonable in a limited amount of time. But one day, I have really time to go all the way with one. Just for the heck of it!

[bomber]

Time ?

Well to create the co-efficients for the 8 airfoil sections took me an evening, about 5 hours... 4 of the hours was used to update the design of the spreadsheet I use to collate the information ready for cutting and pasting into the xml file. Because I wanted to make the spreadsheet more easily understood with data for lift and drag but also the 2 axial forces and graphs.

How long does it take to create a planes 3d and 2 textures... should an FDM be created in the same time period rather than being presented with a wtg 3d model and then thinking lets bang an fdm on it in the next half hour and we can go fly it ?

[jwocky]

Again ... you got me wrong. I LACK the time. Right now, I fight for every hour due to out of FG activities. Some nasties are on the road, my early prediction model for detecting such nasties needs a lot of work and aside of all, I actually try to write also now and then a book. Over the last few weeks, I worked from about 8am till 4am which gives me on average four hours sleep per night ... and yes, under such circumstances another hour later hurts me. Especially, since you needed 1hour effective work for ONE plane and I have, oled ones and new ones, I would like to work on, about 40 to 45 planes in my list. Roughly a complete work week for the average guy. Additionally to the two average work weeks I do already per week. So, that is what I meant with TIIIIMMMMEEEEE!