GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 562 AIRFOIL (goe562-il) Reynolds number: 200,000 Max Cl/Cd: 70.55 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe562-il-200000-n5.txt Download as CSV file: xf-goe562-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 562 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.1784 0.12353 0.12016 -0.0519 0.9880 0.0258
-11.000 -0.1704 0.11956 0.11620 -0.0553 0.9800 0.0259
-10.750 -0.1808 0.11758 0.11431 -0.0519 0.9341 0.0260
-10.500 -0.1625 0.11211 0.10879 -0.0592 0.9130 0.0261
-10.250 -0.1373 0.10577 0.10236 -0.0695 0.8970 0.0262
-10.000 -0.0832 0.10009 0.09652 -0.0766 0.8791 0.0277
-9.750 -0.0385 0.09385 0.09010 -0.0891 0.8578 0.0290
-9.500 -0.0008 0.08728 0.08333 -0.1048 0.8322 0.0319
-9.000 0.0473 0.07658 0.07225 -0.1240 0.7745 0.0323
-8.250 0.0746 0.06801 0.06331 -0.1266 0.7105 0.0304
-8.000 0.0716 0.06490 0.06013 -0.1278 0.6962 0.0299
-7.750 0.0643 0.06194 0.05713 -0.1279 0.6843 0.0291
-7.500 0.0566 0.05887 0.05401 -0.1278 0.6741 0.0288
-7.250 0.0543 0.05581 0.05084 -0.1279 0.6649 0.0292
-7.000 0.0538 0.05266 0.04759 -0.1278 0.6566 0.0301
-6.750 0.0536 0.04927 0.04404 -0.1271 0.6489 0.0304
-6.500 0.0542 0.04565 0.04024 -0.1259 0.6421 0.0309
-6.250 0.0567 0.04232 0.03669 -0.1241 0.6353 0.0309
-6.000 0.0601 0.03898 0.03306 -0.1217 0.6295 0.0311
-5.750 0.0577 0.03454 0.02809 -0.1176 0.6242 0.0327
-5.500 0.0733 0.03366 0.02717 -0.1163 0.6178 0.0335
-5.250 0.0863 0.03212 0.02544 -0.1143 0.6120 0.0341
-5.000 0.0989 0.03039 0.02348 -0.1119 0.6060 0.0346
-4.750 0.1120 0.02861 0.02138 -0.1094 0.6007 0.0349
-4.500 0.1279 0.02735 0.01990 -0.1074 0.5951 0.0362
-4.250 0.1437 0.02593 0.01818 -0.1051 0.5896 0.0373
-4.000 0.1605 0.02449 0.01640 -0.1028 0.5848 0.0377
-3.750 0.1790 0.02324 0.01485 -0.1009 0.5801 0.0381
-3.500 0.1987 0.02216 0.01350 -0.0991 0.5748 0.0385
-3.250 0.2200 0.02149 0.01252 -0.0976 0.5699 0.0397
-3.000 0.2419 0.02093 0.01169 -0.0962 0.5650 0.0404
-2.750 0.2643 0.01984 0.01045 -0.0951 0.5598 0.0408
-2.500 0.2879 0.01909 0.00955 -0.0942 0.5551 0.0413
-2.250 0.3125 0.01851 0.00883 -0.0936 0.5512 0.0419
-2.000 0.3367 0.01800 0.00827 -0.0928 0.5463 0.0423
-1.750 0.3609 0.01757 0.00777 -0.0920 0.5415 0.0431
-1.500 0.3851 0.01729 0.00741 -0.0913 0.5372 0.0448
-1.250 0.4094 0.01698 0.00705 -0.0905 0.5329 0.0461
-1.000 0.4331 0.01667 0.00671 -0.0896 0.5281 0.0470
-0.750 0.4562 0.01639 0.00637 -0.0886 0.5234 0.0479
-0.500 0.4795 0.01618 0.00608 -0.0877 0.5193 0.0488
-0.250 0.5016 0.01598 0.00588 -0.0865 0.5144 0.0497
0.000 0.5228 0.01577 0.00566 -0.0851 0.5090 0.0506
0.250 0.5442 0.01557 0.00540 -0.0838 0.5047 0.0525
0.500 0.5657 0.01546 0.00530 -0.0825 0.4994 0.0554
0.750 0.5876 0.01540 0.00520 -0.0813 0.4947 0.0593
1.000 0.6101 0.01536 0.00509 -0.0802 0.4907 0.0627
1.250 0.6319 0.01530 0.00500 -0.0789 0.4861 0.0692
1.500 0.6527 0.01522 0.00498 -0.0775 0.4806 0.0854
2.000 0.8994 0.01423 0.00586 -0.1208 0.4608 1.0000
2.500 0.9423 0.01452 0.00605 -0.1184 0.4529 1.0000
2.750 0.9634 0.01467 0.00616 -0.1171 0.4486 1.0000
3.000 0.9840 0.01483 0.00628 -0.1158 0.4445 1.0000
3.250 1.0048 0.01500 0.00639 -0.1145 0.4409 1.0000
3.500 1.0256 0.01516 0.00656 -0.1132 0.4363 1.0000
3.750 1.0459 0.01533 0.00672 -0.1118 0.4318 1.0000
4.000 1.0657 0.01552 0.00687 -0.1103 0.4279 1.0000
4.250 1.0857 0.01571 0.00705 -0.1089 0.4236 1.0000
4.500 1.1054 0.01589 0.00726 -0.1074 0.4189 1.0000
4.750 1.1242 0.01609 0.00745 -0.1058 0.4143 1.0000
5.000 1.1427 0.01631 0.00764 -0.1041 0.4100 1.0000
5.250 1.1613 0.01651 0.00789 -0.1024 0.4048 1.0000
5.500 1.1785 0.01672 0.00812 -0.1005 0.3992 1.0000
5.750 1.1952 0.01697 0.00834 -0.0985 0.3946 1.0000
6.000 1.2121 0.01718 0.00861 -0.0965 0.3889 1.0000
6.250 1.2262 0.01742 0.00885 -0.0940 0.3833 1.0000
6.500 1.2401 0.01766 0.00912 -0.0914 0.3784 1.0000
6.750 1.2542 0.01790 0.00941 -0.0890 0.3722 1.0000
7.000 1.2666 0.01818 0.00969 -0.0862 0.3666 1.0000
7.250 1.2806 0.01846 0.01002 -0.0838 0.3606 1.0000
7.500 1.2932 0.01878 0.01038 -0.0812 0.3542 1.0000
7.750 1.3054 0.01913 0.01075 -0.0785 0.3483 1.0000
8.000 1.3183 0.01949 0.01117 -0.0761 0.3417 1.0000
8.250 1.3285 0.01993 0.01161 -0.0732 0.3352 1.0000
8.500 1.3399 0.02037 0.01210 -0.0707 0.3272 1.0000
8.750 1.3481 0.02092 0.01266 -0.0677 0.3187 1.0000
9.000 1.3551 0.02154 0.01329 -0.0646 0.3092 1.0000
9.250 1.3626 0.02222 0.01398 -0.0618 0.3000 1.0000
9.500 1.3657 0.02310 0.01484 -0.0585 0.2892 1.0000
9.750 1.3713 0.02399 0.01574 -0.0557 0.2814 1.0000
10.000 1.3745 0.02505 0.01682 -0.0529 0.2720 1.0000
10.250 1.3773 0.02624 0.01801 -0.0502 0.2647 1.0000
10.500 1.3821 0.02741 0.01923 -0.0480 0.2581 1.0000
10.750 1.3849 0.02879 0.02064 -0.0457 0.2510 1.0000
11.000 1.3864 0.03034 0.02222 -0.0436 0.2444 1.0000
11.250 1.3905 0.03181 0.02377 -0.0419 0.2381 1.0000
11.500 1.3904 0.03367 0.02565 -0.0401 0.2324 1.0000
11.750 1.3922 0.03549 0.02756 -0.0386 0.2248 1.0000
12.000 1.3884 0.03785 0.02994 -0.0370 0.2170 1.0000
12.250 1.3893 0.03995 0.03213 -0.0358 0.2088 1.0000
12.500 1.3831 0.04276 0.03496 -0.0346 0.1987 1.0000
12.750 1.3787 0.04553 0.03778 -0.0336 0.1843 1.0000
13.000 1.3559 0.05026 0.04236 -0.0324 0.1378 1.0000
13.250 1.3323 0.05524 0.04720 -0.0314 0.1138 1.0000
13.500 1.2979 0.06162 0.05343 -0.0305 0.0783 1.0000
13.750 1.2801 0.06641 0.05822 -0.0301 0.0637 1.0000
14.000 1.2641 0.07117 0.06300 -0.0299 0.0524 1.0000
14.250 1.2493 0.07598 0.06787 -0.0300 0.0401 1.0000
14.500 1.2353 0.08082 0.07275 -0.0303 0.0287 1.0000
14.750 1.2220 0.08570 0.07770 -0.0307 0.0243 1.0000
15.000 1.2113 0.09033 0.08241 -0.0311 0.0227 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 562 AIRFOIL (goe562-il)