AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 100,000 Max Cl/Cd: 50.78 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg12a-il-100000.txt Download as CSV file: xf-rg12a-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4733 0.10861 0.10347 -0.0200 1.0000 0.0886
-9.500 -0.4866 0.10573 0.10069 -0.0246 1.0000 0.0914
-9.250 -0.5076 0.10267 0.09780 -0.0305 1.0000 0.0921
-9.000 -0.4790 0.09752 0.09257 -0.0246 1.0000 0.0952
-8.750 -0.4728 0.09438 0.08946 -0.0240 1.0000 0.0985
-8.500 -0.3863 0.08150 0.07694 -0.0271 1.0000 0.1144
-8.250 -0.4069 0.07775 0.07329 -0.0301 1.0000 0.1185
-8.000 -0.4293 0.07387 0.06951 -0.0328 1.0000 0.1191
-7.750 -0.4048 0.07011 0.06575 -0.0284 1.0000 0.1265
-7.500 -0.4221 0.06656 0.06230 -0.0292 1.0000 0.1303
-7.250 -0.4458 0.06258 0.05839 -0.0313 1.0000 0.1317
-7.000 -0.4797 0.05843 0.05410 -0.0349 1.0000 0.1335
-6.750 -0.5136 0.06598 0.06140 -0.0327 1.0000 0.1305
-6.500 -0.5137 0.06224 0.05764 -0.0326 1.0000 0.1379
-6.250 -0.5136 0.05895 0.05423 -0.0328 1.0000 0.1501
-5.750 -0.4968 0.05314 0.04840 -0.0302 1.0000 0.1703
-5.500 -0.4780 0.03983 0.03328 -0.0344 1.0000 0.0800
-5.250 -0.4578 0.03454 0.02748 -0.0329 1.0000 0.0626
-5.000 -0.4384 0.03066 0.02314 -0.0317 1.0000 0.0580
-4.750 -0.4153 0.02729 0.01894 -0.0301 1.0000 0.0543
-4.500 -0.3922 0.02499 0.01627 -0.0289 1.0000 0.0543
-4.250 -0.3693 0.02309 0.01419 -0.0279 1.0000 0.0567
-4.000 -0.3461 0.02186 0.01275 -0.0268 1.0000 0.0630
-3.750 -0.3224 0.02013 0.01091 -0.0256 1.0000 0.0679
-3.500 -0.2991 0.01896 0.00967 -0.0244 1.0000 0.0784
-3.250 -0.2768 0.01751 0.00846 -0.0235 1.0000 0.1038
-3.000 -0.2548 0.01583 0.00736 -0.0227 1.0000 0.1861
-2.750 -0.2423 0.01317 0.00706 -0.0202 1.0000 0.5903
-2.500 -0.2408 0.01320 0.00759 -0.0128 1.0000 0.8000
-2.250 -0.2360 0.01332 0.00773 -0.0067 1.0000 0.8703
-2.000 -0.2141 0.01345 0.00778 -0.0037 1.0000 0.9418
-1.750 -0.0928 0.01388 0.00758 -0.0204 1.0000 1.0000
-1.500 -0.0980 0.01373 0.00738 -0.0160 1.0000 1.0000
-1.250 -0.0996 0.01365 0.00721 -0.0123 1.0000 1.0000
-1.000 -0.0911 0.01370 0.00714 -0.0103 1.0000 1.0000
-0.750 -0.0763 0.01387 0.00718 -0.0094 1.0000 1.0000
-0.500 -0.0334 0.01433 0.00745 -0.0135 0.9923 1.0000
-0.250 0.0185 0.01479 0.00776 -0.0191 0.9811 1.0000
0.000 0.0706 0.01516 0.00802 -0.0245 0.9701 1.0000
0.250 0.1198 0.01536 0.00812 -0.0292 0.9576 1.0000
0.500 0.1676 0.01546 0.00817 -0.0335 0.9446 1.0000
0.750 0.2148 0.01548 0.00816 -0.0374 0.9315 1.0000
1.000 0.2623 0.01542 0.00809 -0.0413 0.9183 1.0000
1.250 0.3106 0.01526 0.00795 -0.0451 0.9052 1.0000
1.500 0.3607 0.01498 0.00773 -0.0490 0.8922 1.0000
1.750 0.4079 0.01466 0.00746 -0.0521 0.8780 1.0000
2.000 0.4529 0.01432 0.00717 -0.0547 0.8624 1.0000
2.250 0.4960 0.01396 0.00689 -0.0568 0.8454 1.0000
2.500 0.5262 0.01382 0.00678 -0.0566 0.8222 1.0000
2.750 0.5582 0.01368 0.00664 -0.0566 0.7993 1.0000
3.000 0.5881 0.01360 0.00655 -0.0562 0.7748 1.0000
3.250 0.6136 0.01366 0.00661 -0.0550 0.7477 1.0000
3.500 0.6386 0.01376 0.00668 -0.0538 0.7195 1.0000
3.750 0.6630 0.01393 0.00679 -0.0526 0.6906 1.0000
4.000 0.6866 0.01414 0.00695 -0.0512 0.6603 1.0000
4.250 0.7090 0.01439 0.00720 -0.0497 0.6282 1.0000
4.500 0.7315 0.01466 0.00742 -0.0483 0.5952 1.0000
4.750 0.7540 0.01497 0.00763 -0.0468 0.5619 1.0000
5.000 0.7752 0.01530 0.00794 -0.0452 0.5253 1.0000
5.250 0.7963 0.01568 0.00828 -0.0436 0.4879 1.0000
5.500 0.8167 0.01612 0.00861 -0.0420 0.4486 1.0000
5.750 0.8361 0.01662 0.00901 -0.0403 0.4061 1.0000
6.000 0.8548 0.01720 0.00951 -0.0386 0.3603 1.0000
6.250 0.8725 0.01792 0.01008 -0.0368 0.3114 1.0000
6.500 0.8881 0.01876 0.01074 -0.0348 0.2557 1.0000
6.750 0.9024 0.01981 0.01147 -0.0329 0.1970 1.0000
7.000 0.9156 0.02132 0.01263 -0.0309 0.1425 1.0000
7.250 0.9272 0.02340 0.01437 -0.0285 0.1019 1.0000
7.500 0.9434 0.02523 0.01613 -0.0266 0.0797 1.0000
7.750 0.9632 0.02725 0.01815 -0.0253 0.0694 1.0000
8.000 0.9871 0.03002 0.02074 -0.0249 0.0630 1.0000
8.250 1.0108 0.03207 0.02321 -0.0238 0.0592 1.0000
8.500 1.0315 0.03403 0.02537 -0.0228 0.0546 1.0000
8.750 1.0533 0.03742 0.02884 -0.0224 0.0516 1.0000
9.000 1.0693 0.04128 0.03312 -0.0209 0.0507 1.0000
9.250 1.0803 0.04472 0.03708 -0.0188 0.0503 1.0000
9.500 1.0868 0.04852 0.04138 -0.0166 0.0502 1.0000
9.750 1.0880 0.05214 0.04548 -0.0140 0.0498 1.0000
10.000 1.0837 0.05577 0.04957 -0.0113 0.0493 1.0000
10.250 1.0752 0.05961 0.05378 -0.0087 0.0491 1.0000
10.500 1.0610 0.06338 0.05787 -0.0061 0.0491 1.0000
10.750 1.0435 0.06714 0.06184 -0.0036 0.0494 1.0000
11.000 1.0251 0.07125 0.06613 -0.0023 0.0497 1.0000
11.250 1.0075 0.07605 0.07107 -0.0023 0.0502 1.0000
11.500 0.9311 0.08792 0.08344 -0.0106 0.0577 1.0000
11.750 0.9027 0.09771 0.09328 -0.0175 0.0600 1.0000
12.000 0.8905 0.10556 0.10111 -0.0216 0.0616 1.0000
12.250 0.8936 0.11071 0.10626 -0.0217 0.0628 1.0000
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Polar data table (+)
Polar graphs
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