GOE 626 AIRFOIL (goe626-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 626 AIRFOIL (goe626-il) Reynolds number: 1,000,000 Max Cl/Cd: 94.86 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe626-il-1000000-n5.txt Download as CSV file: xf-goe626-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 626 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.750 -0.7658 0.09063 0.08729 -0.0690 1.0000 0.0234
-16.500 -0.7661 0.08773 0.08435 -0.0699 1.0000 0.0235
-16.250 -0.7667 0.08488 0.08147 -0.0708 1.0000 0.0237
-16.000 -0.7659 0.08235 0.07891 -0.0713 1.0000 0.0237
-15.750 -0.7657 0.07980 0.07634 -0.0718 1.0000 0.0239
-15.500 -0.7651 0.07683 0.07334 -0.0731 0.9999 0.0240
-15.250 -0.7554 0.07371 0.07017 -0.0762 0.9992 0.0242
-15.000 -0.7449 0.07081 0.06724 -0.0791 0.9984 0.0243
-14.750 -0.7343 0.06802 0.06442 -0.0818 0.9974 0.0243
-14.500 -0.7253 0.06484 0.06119 -0.0850 0.9960 0.0245
-14.250 -0.7138 0.06202 0.05833 -0.0881 0.9942 0.0246
-14.000 -0.7037 0.05934 0.05562 -0.0906 0.9919 0.0247
-13.750 -0.6962 0.05654 0.05279 -0.0930 0.9892 0.0247
-13.500 -0.6880 0.05361 0.04982 -0.0959 0.9862 0.0249
-13.250 -0.6770 0.05087 0.04705 -0.0991 0.9833 0.0249
-12.750 -0.6712 0.04420 0.04030 -0.1060 0.9722 0.0253
-12.500 -0.6674 0.04033 0.03637 -0.1112 0.9660 0.0254
-12.000 -0.6486 0.02865 0.02444 -0.1334 0.9434 0.0256
-11.750 -0.6288 0.02256 0.01813 -0.1446 0.9314 0.0258
-11.500 -0.6041 0.01936 0.01476 -0.1494 0.9220 0.0260
-11.250 -0.5815 0.01805 0.01334 -0.1507 0.9131 0.0262
-11.000 -0.5572 0.01709 0.01228 -0.1515 0.9047 0.0264
-10.750 -0.5339 0.01632 0.01144 -0.1517 0.8962 0.0266
-10.500 -0.5097 0.01567 0.01070 -0.1518 0.8879 0.0268
-10.250 -0.4853 0.01508 0.01004 -0.1518 0.8812 0.0271
-10.000 -0.4609 0.01456 0.00945 -0.1517 0.8732 0.0273
-9.750 -0.4358 0.01410 0.00890 -0.1516 0.8650 0.0276
-9.500 -0.4107 0.01366 0.00840 -0.1513 0.8568 0.0278
-9.000 -0.3598 0.01292 0.00751 -0.1508 0.8376 0.0283
-8.750 -0.3343 0.01261 0.00711 -0.1505 0.8269 0.0285
-8.500 -0.3083 0.01231 0.00675 -0.1502 0.8169 0.0288
-8.250 -0.2824 0.01199 0.00636 -0.1499 0.8079 0.0292
-8.000 -0.2563 0.01164 0.00595 -0.1496 0.7981 0.0296
-7.750 -0.2303 0.01136 0.00560 -0.1492 0.7873 0.0301
-7.500 -0.2042 0.01111 0.00529 -0.1489 0.7744 0.0305
-7.250 -0.1782 0.01089 0.00500 -0.1484 0.7587 0.0310
-7.000 -0.1528 0.01073 0.00473 -0.1478 0.7379 0.0315
-6.750 -0.1284 0.01063 0.00449 -0.1470 0.7079 0.0320
-6.500 -0.1049 0.01060 0.00429 -0.1460 0.6718 0.0326
-6.250 -0.0803 0.01056 0.00411 -0.1452 0.6461 0.0331
-6.000 -0.0542 0.01044 0.00391 -0.1448 0.6303 0.0344
-5.750 -0.0277 0.01031 0.00372 -0.1444 0.6186 0.0361
-5.500 -0.0010 0.01016 0.00352 -0.1440 0.6077 0.0396
-5.250 0.0250 0.00969 0.00316 -0.1438 0.5996 0.0726
-5.000 0.0524 0.00954 0.00299 -0.1436 0.5919 0.0809
-4.750 0.0794 0.00940 0.00284 -0.1433 0.5841 0.0894
-4.500 0.1069 0.00915 0.00263 -0.1432 0.5773 0.1066
-4.250 0.1337 0.00885 0.00241 -0.1429 0.5699 0.1341
-4.000 0.1607 0.00862 0.00224 -0.1427 0.5635 0.1609
-3.750 0.1885 0.00845 0.00213 -0.1426 0.5573 0.1821
-3.500 0.2159 0.00838 0.00209 -0.1423 0.5492 0.1980
-3.250 0.2437 0.00836 0.00206 -0.1420 0.5420 0.2099
-3.000 0.2716 0.00835 0.00206 -0.1418 0.5341 0.2200
-2.750 0.2992 0.00839 0.00205 -0.1415 0.5262 0.2249
-2.500 0.3273 0.00839 0.00205 -0.1413 0.5190 0.2294
-2.250 0.3545 0.00844 0.00207 -0.1409 0.5082 0.2345
-2.000 0.3819 0.00851 0.00208 -0.1405 0.4953 0.2385
-1.750 0.4088 0.00858 0.00209 -0.1401 0.4823 0.2412
-1.500 0.4357 0.00865 0.00211 -0.1396 0.4711 0.2440
-1.250 0.4633 0.00870 0.00214 -0.1393 0.4625 0.2471
-1.000 0.4902 0.00879 0.00219 -0.1389 0.4531 0.2504
-0.750 0.5175 0.00887 0.00222 -0.1385 0.4436 0.2528
-0.500 0.5439 0.00899 0.00227 -0.1380 0.4324 0.2544
-0.250 0.5707 0.00907 0.00230 -0.1375 0.4211 0.2566
0.000 0.5973 0.00914 0.00235 -0.1371 0.4126 0.2589
0.250 0.6244 0.00920 0.00239 -0.1367 0.4056 0.2607
0.500 0.6511 0.00929 0.00245 -0.1362 0.3988 0.2622
0.750 0.6782 0.00935 0.00250 -0.1358 0.3937 0.2640
1.000 0.7052 0.00942 0.00255 -0.1354 0.3881 0.2654
1.250 0.7317 0.00951 0.00262 -0.1350 0.3820 0.2667
1.500 0.7584 0.00959 0.00268 -0.1345 0.3769 0.2679
1.750 0.7849 0.00968 0.00274 -0.1340 0.3690 0.2690
2.000 0.8105 0.00981 0.00283 -0.1334 0.3604 0.2700
2.250 0.8371 0.00989 0.00290 -0.1329 0.3547 0.2708
2.500 0.8627 0.01001 0.00298 -0.1323 0.3465 0.2723
2.750 0.8882 0.01013 0.00307 -0.1317 0.3373 0.2743
3.000 0.9129 0.01029 0.00319 -0.1309 0.3267 0.2760
3.250 0.9384 0.01040 0.00329 -0.1303 0.3193 0.2776
3.500 0.9634 0.01054 0.00340 -0.1296 0.3115 0.2792
3.750 0.9884 0.01067 0.00353 -0.1289 0.3053 0.2808
4.000 1.0132 0.01081 0.00364 -0.1281 0.2976 0.2824
4.250 1.0370 0.01099 0.00379 -0.1272 0.2891 0.2839
4.500 1.0605 0.01118 0.00394 -0.1263 0.2775 0.2853
4.750 1.0824 0.01145 0.00413 -0.1250 0.2619 0.2866
5.000 1.1033 0.01171 0.00432 -0.1236 0.2470 0.2877
5.250 1.1199 0.01213 0.00460 -0.1215 0.2223 0.2889
5.500 1.1310 0.01278 0.00504 -0.1184 0.1833 0.2910
5.750 1.1472 0.01321 0.00538 -0.1162 0.1668 0.2934
6.000 1.1673 0.01347 0.00564 -0.1147 0.1608 0.2957
6.250 1.1872 0.01374 0.00590 -0.1132 0.1558 0.2978
6.500 1.2068 0.01403 0.00618 -0.1116 0.1507 0.2998
6.750 1.2277 0.01424 0.00641 -0.1103 0.1483 0.3019
7.000 1.2479 0.01450 0.00668 -0.1089 0.1452 0.3038
7.250 1.2674 0.01479 0.00696 -0.1073 0.1416 0.3055
7.500 1.2860 0.01511 0.00728 -0.1057 0.1377 0.3071
7.750 1.3056 0.01537 0.00757 -0.1042 0.1351 0.3104
8.000 1.3249 0.01565 0.00788 -0.1028 0.1312 0.3138
8.250 1.3417 0.01605 0.00826 -0.1009 0.1253 0.3168
8.500 1.3591 0.01642 0.00863 -0.0992 0.1201 0.3201
8.750 1.3736 0.01694 0.00910 -0.0972 0.1104 0.3230
9.000 1.3688 0.01838 0.01029 -0.0925 0.0707 0.3249
9.250 1.3787 0.01914 0.01104 -0.0899 0.0623 0.3299
9.500 1.3915 0.01977 0.01169 -0.0878 0.0590 0.3354
9.750 1.4043 0.02043 0.01236 -0.0858 0.0562 0.3411
10.000 1.4181 0.02103 0.01301 -0.0840 0.0549 0.3476
10.250 1.4317 0.02167 0.01370 -0.0822 0.0537 0.3571
10.500 1.4446 0.02236 0.01444 -0.0804 0.0529 0.3672
11.000 1.4690 0.02387 0.01609 -0.0769 0.0513 0.4070
11.250 1.4800 0.02472 0.01705 -0.0751 0.0504 0.4433
11.500 1.4903 0.02563 0.01808 -0.0734 0.0494 0.4909
11.750 1.5010 0.02636 0.01914 -0.0717 0.0489 0.6366
12.250 1.5348 0.02789 0.02135 -0.0710 0.0479 1.0000
12.500 1.5452 0.02903 0.02252 -0.0697 0.0473 1.0000
12.750 1.5547 0.03027 0.02380 -0.0683 0.0468 1.0000
13.000 1.5633 0.03164 0.02521 -0.0670 0.0463 1.0000
13.250 1.5712 0.03311 0.02672 -0.0658 0.0458 1.0000
13.500 1.5786 0.03467 0.02832 -0.0647 0.0454 1.0000
13.750 1.5851 0.03637 0.03007 -0.0636 0.0449 1.0000
14.000 1.5904 0.03827 0.03201 -0.0627 0.0443 1.0000
14.250 1.5949 0.04031 0.03410 -0.0619 0.0438 1.0000
14.500 1.5995 0.04243 0.03628 -0.0612 0.0435 1.0000
14.750 1.6020 0.04482 0.03874 -0.0606 0.0431 1.0000
15.000 1.6070 0.04703 0.04101 -0.0602 0.0428 1.0000
15.250 1.6113 0.04937 0.04341 -0.0599 0.0427 1.0000
15.500 1.6161 0.05169 0.04579 -0.0596 0.0423 1.0000
15.750 1.6193 0.05421 0.04838 -0.0594 0.0422 1.0000
16.000 1.6225 0.05677 0.05100 -0.0593 0.0417 1.0000
16.250 1.6241 0.05957 0.05387 -0.0593 0.0413 1.0000
16.500 1.6254 0.06245 0.05681 -0.0594 0.0409 1.0000
16.750 1.6259 0.06546 0.05990 -0.0596 0.0405 1.0000
17.000 1.6253 0.06868 0.06318 -0.0599 0.0401 1.0000
17.250 1.6239 0.07206 0.06663 -0.0603 0.0398 1.0000
17.500 1.6225 0.07551 0.07014 -0.0608 0.0393 1.0000
17.750 1.6210 0.07901 0.07371 -0.0615 0.0389 1.0000
18.000 1.6177 0.08282 0.07759 -0.0623 0.0385 1.0000
18.250 1.6138 0.08680 0.08163 -0.0632 0.0380 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 626 AIRFOIL (goe626-il)