RG 12A-1.8/9.0 AIRFOIL (rg12a189-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RG 12A-1.8/9.0 AIRFOIL (rg12a189-il) Reynolds number: 200,000 Max Cl/Cd: 67.07 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg12a189-il-200000.txt Download as CSV file: xf-rg12a189-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RG 12A-1.8/9.0 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3795 0.10118 0.09771 -0.0250 1.0000 0.0439
-10.000 -0.3784 0.09761 0.09416 -0.0256 1.0000 0.0451
-9.750 -0.4871 0.10057 0.09697 -0.0258 1.0000 0.0416
-9.500 -0.4808 0.09770 0.09411 -0.0250 1.0000 0.0426
-9.250 -0.4775 0.09457 0.09099 -0.0252 1.0000 0.0438
-9.000 -0.4767 0.09118 0.08763 -0.0261 1.0000 0.0449
-8.750 -0.4779 0.08759 0.08408 -0.0273 1.0000 0.0460
-8.500 -0.4815 0.08380 0.08034 -0.0289 1.0000 0.0471
-8.250 -0.4884 0.07985 0.07645 -0.0309 1.0000 0.0480
-8.000 -0.5024 0.07569 0.07237 -0.0332 1.0000 0.0486
-7.750 -0.5133 0.07106 0.06773 -0.0362 1.0000 0.0493
-7.250 -0.5347 0.06340 0.05956 -0.0398 1.0000 0.0523
-7.000 -0.5411 0.05662 0.05271 -0.0394 1.0000 0.0534
-6.750 -0.5325 0.05316 0.04933 -0.0382 1.0000 0.0546
-6.500 -0.5241 0.05059 0.04676 -0.0368 1.0000 0.0560
-6.250 -0.5160 0.04784 0.04392 -0.0356 1.0000 0.0581
-6.000 -0.5068 0.04476 0.04059 -0.0345 1.0000 0.0621
-5.750 -0.4997 0.04076 0.03614 -0.0336 1.0000 0.0675
-5.500 -0.4859 0.03814 0.03351 -0.0324 1.0000 0.0695
-5.250 -0.4611 0.02816 0.02188 -0.0289 1.0000 0.0295
-5.000 -0.4419 0.02412 0.01750 -0.0279 1.0000 0.0281
-4.750 -0.4199 0.02169 0.01468 -0.0268 1.0000 0.0278
-4.500 -0.3975 0.02104 0.01373 -0.0255 1.0000 0.0293
-4.250 -0.3741 0.01829 0.01076 -0.0248 1.0000 0.0313
-4.000 -0.3507 0.01698 0.00936 -0.0239 1.0000 0.0328
-3.750 -0.3277 0.01601 0.00833 -0.0230 1.0000 0.0348
-3.500 -0.3050 0.01521 0.00747 -0.0220 1.0000 0.0386
-3.250 -0.2826 0.01436 0.00665 -0.0213 1.0000 0.0472
-3.000 -0.2594 0.01339 0.00577 -0.0206 1.0000 0.0766
-2.750 -0.2250 0.01221 0.00529 -0.0229 0.9969 0.2125
-2.500 -0.1919 0.01055 0.00527 -0.0251 0.9925 0.5684
-2.250 -0.1602 0.01044 0.00577 -0.0250 0.9853 0.7728
-2.000 -0.1262 0.01054 0.00587 -0.0256 0.9771 0.8213
-1.750 -0.0897 0.01067 0.00594 -0.0266 0.9705 0.8599
-1.500 -0.0620 0.01064 0.00590 -0.0256 0.9603 0.8907
-1.250 -0.0320 0.01060 0.00582 -0.0252 0.9514 0.9171
-1.000 0.0068 0.01055 0.00569 -0.0266 0.9451 0.9385
-0.750 0.0492 0.01050 0.00557 -0.0289 0.9380 0.9573
-0.500 0.1045 0.01047 0.00546 -0.0340 0.9342 0.9703
-0.250 0.1699 0.01038 0.00530 -0.0413 0.9324 0.9778
0.000 0.2328 0.01020 0.00507 -0.0482 0.9300 0.9845
0.250 0.2884 0.00995 0.00480 -0.0538 0.9211 0.9908
0.500 0.3404 0.00967 0.00450 -0.0586 0.9109 0.9951
0.750 0.3894 0.00938 0.00420 -0.0628 0.8965 0.9983
1.000 0.4311 0.00914 0.00393 -0.0655 0.8769 1.0000
1.250 0.4611 0.00899 0.00373 -0.0659 0.8526 1.0000
1.500 0.4873 0.00892 0.00359 -0.0654 0.8265 1.0000
1.750 0.5103 0.00890 0.00348 -0.0643 0.7993 1.0000
2.000 0.5315 0.00893 0.00344 -0.0629 0.7725 1.0000
2.250 0.5512 0.00901 0.00342 -0.0612 0.7454 1.0000
2.500 0.5696 0.00913 0.00344 -0.0593 0.7185 1.0000
2.750 0.5871 0.00926 0.00349 -0.0573 0.6919 1.0000
3.000 0.6043 0.00942 0.00360 -0.0552 0.6656 1.0000
3.250 0.6219 0.00960 0.00370 -0.0532 0.6392 1.0000
3.500 0.6408 0.00981 0.00382 -0.0514 0.6119 1.0000
3.750 0.6617 0.01004 0.00397 -0.0501 0.5835 1.0000
4.000 0.6837 0.01029 0.00413 -0.0489 0.5540 1.0000
4.250 0.7061 0.01057 0.00435 -0.0478 0.5232 1.0000
4.750 0.7512 0.01120 0.00480 -0.0458 0.4564 1.0000
5.000 0.7735 0.01158 0.00508 -0.0448 0.4214 1.0000
5.250 0.7960 0.01197 0.00541 -0.0439 0.3854 1.0000
5.500 0.8176 0.01243 0.00575 -0.0429 0.3447 1.0000
5.750 0.8382 0.01300 0.00612 -0.0418 0.2951 1.0000
6.000 0.8587 0.01363 0.00655 -0.0407 0.2438 1.0000
6.250 0.8792 0.01434 0.00706 -0.0397 0.1977 1.0000
6.500 0.8993 0.01515 0.00769 -0.0387 0.1507 1.0000
6.750 0.9176 0.01621 0.00847 -0.0375 0.1019 1.0000
7.000 0.9321 0.01783 0.00981 -0.0355 0.0588 1.0000
7.250 0.9482 0.01921 0.01120 -0.0335 0.0470 1.0000
7.500 0.9647 0.02054 0.01252 -0.0319 0.0410 1.0000
7.750 0.9827 0.02183 0.01392 -0.0303 0.0373 1.0000
8.000 1.0017 0.02313 0.01535 -0.0289 0.0346 1.0000
8.250 1.0210 0.02458 0.01687 -0.0277 0.0327 1.0000
8.500 1.0405 0.02648 0.01883 -0.0266 0.0311 1.0000
8.750 1.0602 0.02972 0.02226 -0.0257 0.0292 1.0000
9.000 1.0792 0.03103 0.02384 -0.0244 0.0279 1.0000
9.250 1.0968 0.03345 0.02658 -0.0230 0.0270 1.0000
9.500 1.1110 0.03631 0.02981 -0.0212 0.0266 1.0000
9.750 1.1202 0.03957 0.03348 -0.0191 0.0263 1.0000
10.000 1.1251 0.04274 0.03703 -0.0166 0.0258 1.0000
10.250 1.1271 0.04554 0.04016 -0.0141 0.0250 1.0000
10.500 1.1311 0.04755 0.04236 -0.0121 0.0237 1.0000
10.750 1.1173 0.05119 0.04638 -0.0083 0.0239 1.0000
11.000 1.0979 0.05495 0.05046 -0.0048 0.0242 1.0000
11.250 1.0747 0.05926 0.05506 -0.0025 0.0247 1.0000
11.500 1.0510 0.06399 0.06001 -0.0021 0.0253 1.0000
11.750 1.0274 0.06926 0.06549 -0.0034 0.0254 1.0000
12.000 1.0034 0.07549 0.07190 -0.0065 0.0258 1.0000
12.250 0.9775 0.08315 0.07973 -0.0115 0.0263 1.0000
12.500 0.9515 0.09220 0.08891 -0.0180 0.0271 1.0000
12.750 0.9261 0.10238 0.09914 -0.0246 0.0284 1.0000
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Polar data table (+)
Polar graphs
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