AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-il) Reynolds number: 500,000 Max Cl/Cd: 86.69 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg12a-il-500000.txt Download as CSV file: xf-rg12a-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4909 0.08429 0.08205 -0.0276 1.0000 0.0172
-8.750 -0.4958 0.07997 0.07777 -0.0295 1.0000 0.0172
-8.500 -0.5032 0.07558 0.07341 -0.0317 1.0000 0.0176
-8.250 -0.5209 0.07071 0.06861 -0.0347 1.0000 0.0175
-8.000 -0.5356 0.06574 0.06361 -0.0371 1.0000 0.0174
-7.750 -0.5444 0.06101 0.05882 -0.0381 1.0000 0.0176
-7.500 -0.5506 0.05634 0.05406 -0.0382 1.0000 0.0180
-7.250 -0.5530 0.05177 0.04934 -0.0376 1.0000 0.0188
-7.000 -0.5428 0.04857 0.04582 -0.0362 1.0000 0.0205
-6.750 -0.5407 0.04546 0.04247 -0.0343 1.0000 0.0206
-6.500 -0.5551 0.03697 0.03362 -0.0328 1.0000 0.0219
-6.250 -0.5427 0.03553 0.03220 -0.0316 1.0000 0.0227
-5.500 -0.4565 0.02096 0.01581 -0.0342 0.9941 0.0145
-5.250 -0.4275 0.01743 0.01193 -0.0347 0.9919 0.0132
-5.000 -0.3951 0.01549 0.00971 -0.0356 0.9894 0.0131
-4.750 -0.3609 0.01415 0.00821 -0.0369 0.9870 0.0135
-4.500 -0.3242 0.01340 0.00735 -0.0387 0.9849 0.0143
-4.250 -0.2893 0.01184 0.00565 -0.0404 0.9831 0.0158
-4.000 -0.2581 0.01134 0.00515 -0.0412 0.9781 0.0185
-3.750 -0.2223 0.01083 0.00456 -0.0429 0.9746 0.0207
-3.500 -0.1852 0.00999 0.00373 -0.0448 0.9720 0.0351
-3.250 -0.1504 0.00931 0.00333 -0.0465 0.9687 0.0989
-3.000 -0.1188 0.00879 0.00306 -0.0476 0.9630 0.1633
-2.750 -0.0831 0.00807 0.00280 -0.0497 0.9597 0.2853
-2.500 -0.0472 0.00702 0.00254 -0.0521 0.9572 0.4969
-2.250 -0.0212 0.00630 0.00246 -0.0517 0.9495 0.6789
-2.000 0.0148 0.00612 0.00238 -0.0531 0.9452 0.7380
-1.750 0.0464 0.00601 0.00227 -0.0536 0.9373 0.7633
-1.500 0.0813 0.00591 0.00216 -0.0548 0.9305 0.7888
-1.250 0.1117 0.00584 0.00211 -0.0549 0.9192 0.8078
-1.000 0.1435 0.00578 0.00202 -0.0554 0.9069 0.8207
-0.750 0.1741 0.00576 0.00195 -0.0557 0.8922 0.8328
-0.500 0.2036 0.00576 0.00189 -0.0557 0.8751 0.8440
-0.250 0.2307 0.00577 0.00186 -0.0552 0.8554 0.8536
0.000 0.2574 0.00582 0.00184 -0.0546 0.8341 0.8628
0.250 0.2833 0.00589 0.00182 -0.0539 0.8116 0.8724
0.500 0.3080 0.00596 0.00182 -0.0529 0.7887 0.8811
0.750 0.3324 0.00603 0.00183 -0.0519 0.7646 0.8905
1.250 0.3791 0.00623 0.00187 -0.0494 0.7178 0.9102
1.500 0.4029 0.00631 0.00189 -0.0483 0.6939 0.9184
1.750 0.4268 0.00641 0.00192 -0.0474 0.6708 0.9260
2.000 0.4508 0.00651 0.00194 -0.0464 0.6475 0.9329
2.250 0.4749 0.00661 0.00199 -0.0455 0.6236 0.9408
2.500 0.4994 0.00671 0.00203 -0.0447 0.5999 0.9485
2.750 0.5245 0.00684 0.00208 -0.0441 0.5751 0.9578
3.500 0.6169 0.00736 0.00237 -0.0463 0.4888 0.9834
3.750 0.6520 0.00760 0.00250 -0.0482 0.4554 0.9914
4.000 0.6831 0.00788 0.00265 -0.0492 0.4192 1.0000
4.250 0.7056 0.00814 0.00282 -0.0484 0.3907 1.0000
4.500 0.7296 0.00847 0.00301 -0.0480 0.3525 1.0000
4.750 0.7537 0.00886 0.00323 -0.0475 0.3117 1.0000
5.000 0.7779 0.00927 0.00348 -0.0471 0.2731 1.0000
5.250 0.8015 0.00975 0.00378 -0.0466 0.2290 1.0000
5.500 0.8251 0.01025 0.00411 -0.0460 0.1896 1.0000
5.750 0.8486 0.01077 0.00446 -0.0455 0.1542 1.0000
6.000 0.8720 0.01129 0.00485 -0.0450 0.1213 1.0000
6.250 0.8949 0.01189 0.00530 -0.0443 0.0901 1.0000
6.500 0.9162 0.01266 0.00586 -0.0434 0.0531 1.0000
6.750 0.9364 0.01360 0.00666 -0.0422 0.0283 1.0000
7.000 0.9592 0.01422 0.00731 -0.0414 0.0241 1.0000
7.250 0.9785 0.01524 0.00844 -0.0400 0.0210 1.0000
7.500 1.0006 0.01588 0.00916 -0.0391 0.0201 1.0000
7.750 1.0223 0.01655 0.00994 -0.0381 0.0189 1.0000
8.000 1.0432 0.01727 0.01073 -0.0371 0.0175 1.0000
8.250 1.0627 0.01814 0.01168 -0.0359 0.0165 1.0000
8.500 1.0799 0.01927 0.01289 -0.0343 0.0156 1.0000
8.750 1.0923 0.02112 0.01488 -0.0321 0.0147 1.0000
9.000 1.1078 0.02276 0.01668 -0.0304 0.0140 1.0000
9.250 1.1279 0.02347 0.01750 -0.0293 0.0132 1.0000
9.500 1.1452 0.02463 0.01880 -0.0280 0.0125 1.0000
9.750 1.1609 0.02606 0.02039 -0.0264 0.0120 1.0000
10.000 1.1756 0.02747 0.02195 -0.0248 0.0114 1.0000
10.250 1.1886 0.02895 0.02356 -0.0230 0.0109 1.0000
10.500 1.1990 0.03060 0.02533 -0.0211 0.0104 1.0000
10.750 1.2027 0.03319 0.02810 -0.0185 0.0100 1.0000
11.000 1.1963 0.03707 0.03230 -0.0150 0.0097 1.0000
11.250 1.1829 0.04046 0.03601 -0.0109 0.0094 1.0000
11.500 1.1836 0.04143 0.03715 -0.0084 0.0091 1.0000
11.750 1.1801 0.04342 0.03933 -0.0063 0.0089 1.0000
12.000 1.1763 0.04558 0.04167 -0.0047 0.0086 1.0000
12.250 1.1615 0.04948 0.04580 -0.0037 0.0086 1.0000
12.500 1.1471 0.05364 0.05017 -0.0037 0.0086 1.0000
12.750 1.1315 0.05839 0.05512 -0.0049 0.0085 1.0000
13.000 1.1146 0.06392 0.06085 -0.0072 0.0085 1.0000
13.250 1.1000 0.06965 0.06674 -0.0103 0.0084 1.0000
13.500 1.0747 0.07800 0.07529 -0.0153 0.0086 1.0000
13.750 1.0537 0.08640 0.08385 -0.0209 0.0087 1.0000
14.000 1.0332 0.09547 0.09307 -0.0270 0.0088 1.0000
14.250 1.0045 0.10721 0.10494 -0.0346 0.0091 1.0000
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Polar data table (+)
Polar graphs
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