AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 200,000 Max Cl/Cd: 67.06 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg12a-il-200000.txt Download as CSV file: xf-rg12a-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL PROFILE12A 9.00%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4777 0.08764 0.08413 -0.0273 1.0000 0.0460
-8.500 -0.4811 0.08385 0.08038 -0.0289 1.0000 0.0470
-8.250 -0.4888 0.07976 0.07637 -0.0310 1.0000 0.0482
-8.000 -0.5024 0.07569 0.07237 -0.0333 1.0000 0.0486
-7.750 -0.5129 0.07114 0.06782 -0.0361 1.0000 0.0493
-7.250 -0.5346 0.06339 0.05955 -0.0398 1.0000 0.0523
-7.000 -0.5413 0.05668 0.05276 -0.0394 1.0000 0.0534
-6.750 -0.5325 0.05320 0.04937 -0.0382 1.0000 0.0545
-6.500 -0.5242 0.05059 0.04675 -0.0368 1.0000 0.0560
-6.250 -0.5161 0.04783 0.04390 -0.0356 1.0000 0.0582
-6.000 -0.5069 0.04478 0.04062 -0.0345 1.0000 0.0620
-5.250 -0.4614 0.02824 0.02197 -0.0288 1.0000 0.0295
-5.000 -0.4423 0.02415 0.01753 -0.0278 1.0000 0.0281
-4.750 -0.4203 0.02170 0.01470 -0.0267 1.0000 0.0278
-4.500 -0.3979 0.02089 0.01358 -0.0255 1.0000 0.0295
-4.250 -0.3745 0.01830 0.01078 -0.0248 1.0000 0.0313
-4.000 -0.3511 0.01699 0.00937 -0.0238 1.0000 0.0326
-3.750 -0.3282 0.01602 0.00835 -0.0229 1.0000 0.0347
-3.500 -0.3055 0.01522 0.00747 -0.0219 1.0000 0.0386
-3.250 -0.2832 0.01438 0.00667 -0.0212 1.0000 0.0481
-3.000 -0.2599 0.01339 0.00577 -0.0205 1.0000 0.0770
-2.750 -0.2247 0.01222 0.00530 -0.0230 0.9967 0.2131
-2.500 -0.1918 0.01056 0.00528 -0.0252 0.9923 0.5691
-2.250 -0.1599 0.01045 0.00577 -0.0250 0.9851 0.7722
-2.000 -0.1262 0.01054 0.00587 -0.0256 0.9768 0.8215
-1.750 -0.0893 0.01067 0.00595 -0.0267 0.9703 0.8596
-1.500 -0.0621 0.01064 0.00591 -0.0256 0.9601 0.8910
-1.250 -0.0321 0.01060 0.00582 -0.0251 0.9511 0.9173
-1.000 0.0069 0.01055 0.00570 -0.0266 0.9449 0.9386
-0.750 0.0488 0.01050 0.00558 -0.0288 0.9377 0.9571
-0.500 0.1047 0.01047 0.00546 -0.0341 0.9340 0.9701
-0.250 0.1701 0.01038 0.00530 -0.0413 0.9323 0.9776
0.000 0.2331 0.01021 0.00508 -0.0483 0.9299 0.9844
0.250 0.2883 0.00996 0.00481 -0.0538 0.9210 0.9907
0.500 0.3404 0.00968 0.00451 -0.0586 0.9108 0.9950
0.750 0.3892 0.00938 0.00420 -0.0628 0.8961 0.9982
1.000 0.4314 0.00914 0.00393 -0.0656 0.8768 1.0000
1.250 0.4614 0.00899 0.00373 -0.0659 0.8523 1.0000
1.500 0.4876 0.00892 0.00359 -0.0654 0.8264 1.0000
1.750 0.5106 0.00890 0.00349 -0.0644 0.7995 1.0000
2.000 0.5317 0.00894 0.00345 -0.0629 0.7724 1.0000
2.250 0.5513 0.00901 0.00343 -0.0612 0.7454 1.0000
2.500 0.5698 0.00913 0.00344 -0.0594 0.7187 1.0000
2.750 0.5873 0.00926 0.00350 -0.0573 0.6921 1.0000
3.000 0.6045 0.00942 0.00360 -0.0552 0.6658 1.0000
3.250 0.6220 0.00960 0.00370 -0.0532 0.6394 1.0000
3.500 0.6408 0.00981 0.00382 -0.0514 0.6121 1.0000
3.750 0.6617 0.01004 0.00397 -0.0500 0.5838 1.0000
4.000 0.6836 0.01029 0.00413 -0.0489 0.5542 1.0000
4.250 0.7060 0.01057 0.00435 -0.0478 0.5233 1.0000
4.500 0.7285 0.01088 0.00456 -0.0468 0.4913 1.0000
4.750 0.7511 0.01120 0.00480 -0.0458 0.4566 1.0000
5.000 0.7733 0.01158 0.00507 -0.0448 0.4213 1.0000
5.250 0.7958 0.01197 0.00541 -0.0439 0.3855 1.0000
5.500 0.8176 0.01243 0.00575 -0.0429 0.3451 1.0000
5.750 0.8381 0.01299 0.00612 -0.0417 0.2956 1.0000
6.000 0.8586 0.01363 0.00655 -0.0407 0.2439 1.0000
6.250 0.8791 0.01434 0.00706 -0.0397 0.1977 1.0000
6.500 0.8991 0.01515 0.00769 -0.0387 0.1509 1.0000
6.750 0.9175 0.01620 0.00847 -0.0375 0.1022 1.0000
7.000 0.9320 0.01782 0.00981 -0.0354 0.0590 1.0000
7.250 0.9481 0.01921 0.01120 -0.0335 0.0470 1.0000
7.500 0.9646 0.02053 0.01252 -0.0318 0.0410 1.0000
7.750 0.9826 0.02183 0.01392 -0.0303 0.0373 1.0000
8.000 1.0017 0.02313 0.01535 -0.0289 0.0347 1.0000
8.250 1.0210 0.02457 0.01686 -0.0277 0.0327 1.0000
8.500 1.0405 0.02646 0.01882 -0.0266 0.0311 1.0000
8.750 1.0603 0.02980 0.02234 -0.0258 0.0293 1.0000
9.000 1.0792 0.03102 0.02384 -0.0244 0.0279 1.0000
9.250 1.0968 0.03346 0.02660 -0.0230 0.0270 1.0000
9.500 1.1109 0.03632 0.02982 -0.0212 0.0265 1.0000
9.750 1.1203 0.03958 0.03349 -0.0191 0.0263 1.0000
10.000 1.1252 0.04270 0.03700 -0.0166 0.0257 1.0000
10.250 1.1277 0.04541 0.04002 -0.0142 0.0248 1.0000
10.500 1.1303 0.04766 0.04249 -0.0120 0.0237 1.0000
10.750 1.1199 0.05091 0.04606 -0.0086 0.0237 1.0000
11.000 1.0974 0.05502 0.05054 -0.0048 0.0243 1.0000
11.250 1.0746 0.05928 0.05509 -0.0026 0.0248 1.0000
11.500 1.0524 0.06375 0.05976 -0.0021 0.0249 1.0000
11.750 1.0278 0.06919 0.06542 -0.0034 0.0254 1.0000
12.000 1.0026 0.07562 0.07204 -0.0066 0.0258 1.0000
12.250 0.9781 0.08300 0.07958 -0.0114 0.0261 1.0000
12.500 0.9511 0.09230 0.08900 -0.0179 0.0273 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AIRFOIL PROFILE12A 9.00% (rg12a-il)