GOE 497 AIRFOIL (goe497-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 497 AIRFOIL (goe497-il) Reynolds number: 200,000 Max Cl/Cd: 84.99 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe497-il-200000.txt Download as CSV file: xf-goe497-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 497 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2576 0.09323 0.08973 -0.0595 0.9806 0.0926
-8.500 -0.4011 0.04270 0.03862 -0.1149 0.9500 0.0638
-8.250 -0.3852 0.03416 0.02916 -0.1248 0.9430 0.0648
-8.000 -0.3512 0.03036 0.02465 -0.1301 0.9404 0.0669
-7.750 -0.3341 0.02806 0.02203 -0.1301 0.9326 0.0684
-7.500 -0.3007 0.02673 0.02066 -0.1320 0.9290 0.0705
-7.250 -0.2632 0.02538 0.01911 -0.1346 0.9265 0.0729
-7.000 -0.2229 0.02410 0.01753 -0.1376 0.9248 0.0762
-6.750 -0.2018 0.02302 0.01620 -0.1369 0.9173 0.0784
-6.500 -0.1681 0.02196 0.01514 -0.1384 0.9138 0.0813
-6.250 -0.1295 0.02118 0.01425 -0.1407 0.9115 0.0854
-6.000 -0.0902 0.02021 0.01307 -0.1430 0.9095 0.0903
-5.750 -0.0502 0.01937 0.01227 -0.1454 0.9078 0.0957
-5.500 -0.0300 0.01902 0.01179 -0.1439 0.8985 0.1006
-5.250 0.0071 0.01813 0.01093 -0.1457 0.8951 0.1083
-5.000 0.0459 0.01735 0.01008 -0.1476 0.8925 0.1181
-4.750 0.0707 0.01702 0.00976 -0.1470 0.8851 0.1276
-4.500 0.1024 0.01647 0.00923 -0.1476 0.8799 0.1402
-4.250 0.1385 0.01593 0.00869 -0.1490 0.8765 0.1548
-4.000 0.1657 0.01564 0.00841 -0.1487 0.8698 0.1681
-3.750 0.1952 0.01534 0.00811 -0.1489 0.8637 0.1823
-3.500 0.2292 0.01497 0.00775 -0.1497 0.8598 0.1983
-3.250 0.2555 0.01480 0.00758 -0.1493 0.8524 0.2137
-3.000 0.2854 0.01454 0.00734 -0.1493 0.8460 0.2315
-2.750 0.3190 0.01427 0.00707 -0.1501 0.8419 0.2528
-2.500 0.3428 0.01426 0.00709 -0.1492 0.8336 0.2723
-2.250 0.3731 0.01411 0.00692 -0.1493 0.8282 0.2936
-2.000 0.4041 0.01395 0.00677 -0.1496 0.8235 0.3125
-1.750 0.4285 0.01395 0.00680 -0.1489 0.8155 0.3294
-1.500 0.4591 0.01380 0.00665 -0.1491 0.8104 0.3469
-1.250 0.4864 0.01377 0.00662 -0.1487 0.8038 0.3629
-1.000 0.5136 0.01368 0.00658 -0.1483 0.7967 0.3787
-0.750 0.5454 0.01353 0.00641 -0.1487 0.7918 0.3961
-0.500 0.5692 0.01355 0.00649 -0.1478 0.7832 0.4115
-0.250 0.5991 0.01343 0.00638 -0.1478 0.7772 0.4285
0.000 0.6257 0.01342 0.00640 -0.1474 0.7700 0.4461
0.250 0.6534 0.01335 0.00635 -0.1470 0.7625 0.4668
0.500 0.6812 0.01324 0.00630 -0.1467 0.7548 0.4874
0.750 0.7079 0.01311 0.00622 -0.1462 0.7458 0.5079
1.000 0.7349 0.01300 0.00616 -0.1457 0.7371 0.5299
1.250 0.7624 0.01284 0.00605 -0.1453 0.7280 0.5539
1.500 0.7878 0.01273 0.00605 -0.1445 0.7183 0.5811
1.750 0.8159 0.01255 0.00595 -0.1442 0.7099 0.6165
2.000 0.8388 0.01238 0.00602 -0.1430 0.6993 0.6689
2.250 0.8661 0.01171 0.00590 -0.1420 0.6897 1.0000
2.500 0.8938 0.01178 0.00583 -0.1417 0.6780 1.0000
2.750 0.9195 0.01189 0.00586 -0.1411 0.6654 1.0000
3.000 0.9459 0.01201 0.00589 -0.1406 0.6531 1.0000
3.250 0.9726 0.01214 0.00591 -0.1402 0.6408 1.0000
3.500 0.9985 0.01229 0.00598 -0.1396 0.6279 1.0000
3.750 1.0230 0.01248 0.00614 -0.1388 0.6148 1.0000
4.000 1.0482 0.01268 0.00629 -0.1381 0.6022 1.0000
4.250 1.0738 0.01290 0.00643 -0.1376 0.5904 1.0000
4.500 1.0988 0.01311 0.00660 -0.1369 0.5784 1.0000
4.750 1.1226 0.01334 0.00682 -0.1360 0.5659 1.0000
5.000 1.1469 0.01358 0.00702 -0.1352 0.5541 1.0000
5.250 1.1716 0.01383 0.00721 -0.1345 0.5429 1.0000
5.500 1.1948 0.01407 0.00745 -0.1336 0.5306 1.0000
5.750 1.2171 0.01432 0.00768 -0.1324 0.5171 1.0000
6.000 1.2386 0.01459 0.00791 -0.1311 0.5023 1.0000
6.250 1.2597 0.01489 0.00816 -0.1298 0.4877 1.0000
6.500 1.2808 0.01524 0.00845 -0.1285 0.4739 1.0000
6.750 1.3015 0.01563 0.00878 -0.1272 0.4603 1.0000
7.000 1.3221 0.01606 0.00913 -0.1259 0.4468 1.0000
7.250 1.3408 0.01646 0.00954 -0.1242 0.4324 1.0000
7.500 1.3591 0.01688 0.00996 -0.1225 0.4184 1.0000
7.750 1.3768 0.01731 0.01041 -0.1207 0.4046 1.0000
8.000 1.3935 0.01776 0.01086 -0.1188 0.3908 1.0000
8.250 1.4084 0.01821 0.01131 -0.1165 0.3766 1.0000
8.500 1.4203 0.01866 0.01175 -0.1137 0.3620 1.0000
8.750 1.4311 0.01914 0.01225 -0.1108 0.3470 1.0000
9.000 1.4414 0.01967 0.01280 -0.1079 0.3315 1.0000
9.250 1.4512 0.02028 0.01340 -0.1051 0.3156 1.0000
9.500 1.4600 0.02099 0.01410 -0.1023 0.2995 1.0000
9.750 1.4676 0.02183 0.01490 -0.0994 0.2836 1.0000
10.000 1.4740 0.02281 0.01583 -0.0966 0.2680 1.0000
10.250 1.4796 0.02392 0.01689 -0.0938 0.2526 1.0000
10.500 1.4843 0.02515 0.01808 -0.0910 0.2367 1.0000
10.750 1.4879 0.02651 0.01941 -0.0883 0.2192 1.0000
11.000 1.4902 0.02801 0.02089 -0.0857 0.1994 1.0000
11.250 1.4905 0.02973 0.02255 -0.0831 0.1784 1.0000
11.500 1.4918 0.03152 0.02426 -0.0808 0.1573 1.0000
11.750 1.4918 0.03350 0.02616 -0.0786 0.1423 1.0000
12.000 1.4913 0.03563 0.02823 -0.0766 0.1320 1.0000
12.250 1.4910 0.03785 0.03040 -0.0747 0.1239 1.0000
12.500 1.4934 0.03992 0.03251 -0.0732 0.1173 1.0000
12.750 1.4944 0.04219 0.03479 -0.0717 0.1119 1.0000
13.000 1.4951 0.04456 0.03717 -0.0703 0.1072 1.0000
13.250 1.5002 0.04659 0.03932 -0.0693 0.1026 1.0000
13.500 1.5014 0.04904 0.04178 -0.0683 0.0984 1.0000
13.750 1.5039 0.05144 0.04422 -0.0673 0.0945 1.0000
14.000 1.5084 0.05371 0.04662 -0.0667 0.0904 1.0000
14.250 1.5100 0.05632 0.04928 -0.0661 0.0865 1.0000
14.500 1.5115 0.05901 0.05201 -0.0655 0.0827 1.0000
14.750 1.5141 0.06170 0.05486 -0.0652 0.0786 1.0000
15.000 1.5135 0.06480 0.05799 -0.0651 0.0747 1.0000
15.250 1.5129 0.06798 0.06127 -0.0650 0.0705 1.0000
15.500 1.5113 0.07145 0.06486 -0.0653 0.0661 1.0000
15.750 1.5070 0.07527 0.06867 -0.0655 0.0619 1.0000
16.000 1.5037 0.07922 0.07279 -0.0662 0.0573 1.0000
16.250 1.4983 0.08332 0.07683 -0.0668 0.0536 1.0000
16.500 1.4948 0.08752 0.08124 -0.0677 0.0498 1.0000
16.750 1.4910 0.09169 0.08544 -0.0687 0.0472 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 497 AIRFOIL (goe497-il)