GOE 614 AIRFOIL (goe614-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 614 AIRFOIL (goe614-il) Reynolds number: 500,000 Max Cl/Cd: 77.06 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe614-il-500000-n5.txt Download as CSV file: xf-goe614-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 614 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.500 -0.3819 0.09586 0.09271 -0.0963 0.9848 0.0240
-15.250 -0.4287 0.08201 0.07860 -0.1069 0.9830 0.0239
-15.000 -0.4531 0.07301 0.06941 -0.1147 0.9801 0.0239
-14.750 -0.4714 0.06605 0.06229 -0.1206 0.9751 0.0239
-14.500 -0.4799 0.05999 0.05608 -0.1269 0.9713 0.0239
-14.250 -0.4880 0.05445 0.05038 -0.1326 0.9661 0.0240
-14.000 -0.4920 0.04976 0.04554 -0.1375 0.9591 0.0240
-13.750 -0.4894 0.04517 0.04079 -0.1435 0.9534 0.0241
-13.500 -0.4832 0.04166 0.03716 -0.1478 0.9448 0.0241
-13.250 -0.4712 0.03763 0.03293 -0.1550 0.9380 0.0242
-13.000 -0.4535 0.03463 0.02982 -0.1599 0.9291 0.0243
-12.750 -0.4245 0.03194 0.02701 -0.1660 0.9223 0.0245
-12.500 -0.3992 0.02969 0.02463 -0.1710 0.9107 0.0245
-12.250 -0.3762 0.02778 0.02260 -0.1748 0.8970 0.0247
-12.000 -0.3581 0.02628 0.02096 -0.1767 0.8814 0.0248
-11.750 -0.3459 0.02514 0.01968 -0.1766 0.8659 0.0249
-11.500 -0.3377 0.02420 0.01861 -0.1751 0.8509 0.0251
-11.250 -0.3315 0.02348 0.01777 -0.1726 0.8372 0.0251
-11.000 -0.3243 0.02282 0.01702 -0.1701 0.8250 0.0253
-10.750 -0.3125 0.02219 0.01628 -0.1683 0.8145 0.0255
-10.500 -0.3002 0.02158 0.01557 -0.1664 0.8032 0.0257
-10.250 -0.2863 0.02098 0.01488 -0.1646 0.7937 0.0259
-10.000 -0.2716 0.02042 0.01422 -0.1629 0.7837 0.0260
-9.750 -0.2559 0.01985 0.01357 -0.1612 0.7743 0.0261
-9.500 -0.2395 0.01934 0.01296 -0.1596 0.7647 0.0263
-9.250 -0.2224 0.01882 0.01237 -0.1580 0.7557 0.0264
-9.000 -0.2048 0.01836 0.01182 -0.1565 0.7459 0.0266
-8.750 -0.1867 0.01788 0.01127 -0.1550 0.7364 0.0267
-8.500 -0.1682 0.01746 0.01076 -0.1535 0.7265 0.0269
-8.250 -0.1489 0.01704 0.01027 -0.1522 0.7173 0.0270
-8.000 -0.1298 0.01667 0.00980 -0.1507 0.7072 0.0272
-7.750 -0.1100 0.01629 0.00935 -0.1493 0.6966 0.0273
-7.500 -0.0905 0.01596 0.00894 -0.1479 0.6859 0.0275
-7.250 -0.0701 0.01563 0.00853 -0.1466 0.6751 0.0276
-7.000 -0.0496 0.01534 0.00816 -0.1452 0.6654 0.0278
-6.750 -0.0289 0.01505 0.00779 -0.1439 0.6550 0.0279
-6.500 -0.0078 0.01479 0.00746 -0.1427 0.6450 0.0280
-6.250 0.0114 0.01443 0.00703 -0.1411 0.6342 0.0283
-6.000 0.0322 0.01412 0.00665 -0.1398 0.6241 0.0285
-5.500 0.0745 0.01360 0.00601 -0.1373 0.6050 0.0292
-5.000 0.1187 0.01322 0.00551 -0.1350 0.5879 0.0300
-4.750 0.1405 0.01307 0.00528 -0.1338 0.5791 0.0304
-4.500 0.1633 0.01291 0.00507 -0.1328 0.5711 0.0307
-4.250 0.1856 0.01278 0.00487 -0.1316 0.5627 0.0310
-4.000 0.2082 0.01268 0.00470 -0.1305 0.5551 0.0314
-3.750 0.2315 0.01257 0.00454 -0.1295 0.5478 0.0318
-3.500 0.2537 0.01250 0.00440 -0.1283 0.5411 0.0322
-3.250 0.2772 0.01239 0.00424 -0.1274 0.5351 0.0327
-3.000 0.3002 0.01229 0.00410 -0.1263 0.5287 0.0335
-2.750 0.3220 0.01224 0.00399 -0.1251 0.5224 0.0344
-2.500 0.3452 0.01216 0.00388 -0.1241 0.5169 0.0355
-2.250 0.3673 0.01211 0.00379 -0.1228 0.5106 0.0368
-2.000 0.3878 0.01204 0.00369 -0.1213 0.5051 0.0405
-1.750 0.4079 0.01174 0.00357 -0.1198 0.5002 0.0909
-1.500 0.4295 0.01169 0.00353 -0.1185 0.4933 0.1043
-1.250 0.4508 0.01166 0.00350 -0.1172 0.4881 0.1145
-0.750 0.4934 0.01145 0.00348 -0.1147 0.4771 0.1735
-0.500 0.5150 0.01145 0.00353 -0.1134 0.4720 0.1973
-0.250 0.5358 0.01153 0.00357 -0.1120 0.4667 0.2089
0.000 0.5594 0.01155 0.00362 -0.1111 0.4620 0.2205
0.250 0.5817 0.01160 0.00368 -0.1100 0.4564 0.2299
0.500 0.6023 0.01171 0.00376 -0.1086 0.4505 0.2393
0.750 0.6240 0.01177 0.00384 -0.1074 0.4444 0.2543
1.000 0.6457 0.01184 0.00393 -0.1062 0.4382 0.2671
1.250 0.6664 0.01194 0.00404 -0.1049 0.4330 0.2824
1.500 0.6879 0.01203 0.00414 -0.1037 0.4286 0.2952
1.750 0.7100 0.01208 0.00425 -0.1026 0.4237 0.3159
2.000 0.7307 0.01217 0.00438 -0.1013 0.4184 0.3376
2.250 0.7501 0.01231 0.00453 -0.0998 0.4132 0.3568
2.500 0.7722 0.01239 0.00466 -0.0988 0.4086 0.3783
2.750 0.7931 0.01248 0.00481 -0.0975 0.4035 0.4022
3.000 0.8130 0.01259 0.00498 -0.0961 0.3993 0.4285
3.250 0.8324 0.01274 0.00516 -0.0947 0.3956 0.4522
3.500 0.8540 0.01284 0.00532 -0.0936 0.3921 0.4754
3.750 0.8750 0.01295 0.00549 -0.0924 0.3880 0.4985
4.000 0.8941 0.01311 0.00570 -0.0909 0.3831 0.5236
4.250 0.9109 0.01331 0.00594 -0.0891 0.3777 0.5539
4.500 0.9305 0.01337 0.00615 -0.0877 0.3732 0.5997
4.750 0.9465 0.01343 0.00641 -0.0856 0.3679 0.6664
5.000 0.9576 0.01333 0.00668 -0.0825 0.3640 0.7999
5.500 1.0727 0.01392 0.00744 -0.0962 0.3479 1.0000
5.750 1.0912 0.01425 0.00773 -0.0948 0.3432 1.0000
6.000 1.1108 0.01456 0.00802 -0.0937 0.3370 1.0000
6.250 1.1281 0.01495 0.00838 -0.0922 0.3306 1.0000
6.500 1.1463 0.01533 0.00873 -0.0909 0.3250 1.0000
6.750 1.1636 0.01576 0.00913 -0.0896 0.3171 1.0000
7.000 1.1795 0.01626 0.00959 -0.0880 0.3098 1.0000
7.250 1.1961 0.01676 0.01006 -0.0866 0.3016 1.0000
7.500 1.2109 0.01734 0.01060 -0.0851 0.2937 1.0000
7.750 1.2255 0.01797 0.01119 -0.0835 0.2846 1.0000
8.000 1.2396 0.01865 0.01184 -0.0820 0.2767 1.0000
8.250 1.2531 0.01938 0.01253 -0.0805 0.2682 1.0000
8.500 1.2663 0.02015 0.01328 -0.0790 0.2607 1.0000
8.750 1.2792 0.02097 0.01407 -0.0775 0.2532 1.0000
9.000 1.2921 0.02181 0.01490 -0.0761 0.2476 1.0000
9.250 1.3045 0.02271 0.01579 -0.0747 0.2402 1.0000
9.500 1.3152 0.02374 0.01680 -0.0733 0.2340 1.0000
9.750 1.3277 0.02469 0.01775 -0.0720 0.2287 1.0000
10.000 1.3400 0.02568 0.01875 -0.0708 0.2237 1.0000
10.250 1.3498 0.02687 0.01993 -0.0695 0.2187 1.0000
10.500 1.3614 0.02797 0.02104 -0.0684 0.2143 1.0000
10.750 1.3723 0.02913 0.02223 -0.0673 0.2096 1.0000
11.000 1.3821 0.03041 0.02351 -0.0662 0.2053 1.0000
11.250 1.3901 0.03187 0.02498 -0.0650 0.2007 1.0000
11.500 1.4010 0.03312 0.02627 -0.0641 0.1969 1.0000
11.750 1.4050 0.03497 0.02810 -0.0628 0.1886 1.0000
12.000 1.4116 0.03664 0.02980 -0.0618 0.1820 1.0000
12.250 1.4126 0.03883 0.03197 -0.0606 0.1711 1.0000
12.750 1.3750 0.04713 0.04006 -0.0569 0.1227 1.0000
13.000 1.3717 0.05004 0.04299 -0.0560 0.1177 1.0000
13.250 1.3745 0.05239 0.04539 -0.0554 0.1157 1.0000
13.500 1.3766 0.05483 0.04789 -0.0549 0.1140 1.0000
13.750 1.3794 0.05725 0.05037 -0.0544 0.1132 1.0000
14.000 1.3792 0.06001 0.05318 -0.0540 0.1115 1.0000
14.250 1.3810 0.06259 0.05582 -0.0536 0.1106 1.0000
14.500 1.3791 0.06561 0.05891 -0.0533 0.1092 1.0000
14.750 1.3789 0.06845 0.06181 -0.0530 0.1083 1.0000
15.000 1.3775 0.07145 0.06487 -0.0527 0.1070 1.0000
15.250 1.3745 0.07470 0.06819 -0.0526 0.1058 1.0000
15.500 1.3742 0.07764 0.07118 -0.0525 0.1050 1.0000
15.750 1.3765 0.08029 0.07391 -0.0524 0.1046 1.0000
16.000 1.3785 0.08297 0.07666 -0.0524 0.1041 1.0000
16.250 1.3794 0.08580 0.07957 -0.0524 0.1033 1.0000
16.500 1.3814 0.08850 0.08234 -0.0525 0.1029 1.0000
16.750 1.3811 0.09152 0.08544 -0.0526 0.1022 1.0000
17.000 1.3818 0.09440 0.08838 -0.0527 0.1010 1.0000
17.250 1.3809 0.09750 0.09155 -0.0530 0.1002 1.0000
17.500 1.3807 0.10049 0.09460 -0.0532 0.0992 1.0000
17.750 1.3788 0.10374 0.09790 -0.0535 0.0982 1.0000
18.000 1.3781 0.10683 0.10105 -0.0539 0.0973 1.0000
18.250 1.3761 0.11012 0.10440 -0.0543 0.0963 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 614 AIRFOIL (goe614-il)