GOE 117 (MVA MK.4) AIRFOIL (goe117-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 117 (MVA MK.4) AIRFOIL (goe117-il) Reynolds number: 200,000 Max Cl/Cd: 78.23 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe117-il-200000.txt Download as CSV file: xf-goe117-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 117 (MVA MK.4) 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3414 0.08855 0.08536 -0.0381 1.0000 0.0452
-7.750 -0.3541 0.08726 0.08415 -0.0352 1.0000 0.0458
-7.500 -0.3752 0.08669 0.08368 -0.0308 1.0000 0.0461
-7.250 -0.3847 0.08496 0.08201 -0.0298 0.9983 0.0471
-7.000 -0.3580 0.07901 0.07604 -0.0412 0.9906 0.0510
-6.750 -0.3250 0.06796 0.06489 -0.0642 0.9802 0.0544
-6.500 -0.3073 0.06634 0.06329 -0.0613 0.9758 0.0566
-6.250 -0.2597 0.05777 0.05432 -0.0819 0.9667 0.0668
-6.000 -0.2480 0.05575 0.05252 -0.0777 0.9623 0.0700
-5.750 -0.2067 0.04762 0.04404 -0.0912 0.9572 0.0818
-5.500 -0.1844 0.04641 0.04289 -0.0901 0.9498 0.0849
-5.250 -0.1462 0.02762 0.02248 -0.1026 0.9440 0.0502
-5.000 -0.1161 0.02423 0.01845 -0.1039 0.9365 0.0509
-4.750 -0.0792 0.02113 0.01474 -0.1059 0.9318 0.0507
-4.500 -0.0445 0.01922 0.01237 -0.1071 0.9256 0.0520
-4.250 -0.0105 0.01751 0.01032 -0.1082 0.9186 0.0551
-4.000 0.0231 0.01600 0.00870 -0.1092 0.9115 0.0575
-3.750 0.0562 0.01505 0.00765 -0.1101 0.9030 0.0612
-3.500 0.0861 0.01435 0.00684 -0.1101 0.8922 0.0651
-3.250 0.1157 0.01362 0.00601 -0.1101 0.8810 0.0677
-3.000 0.1439 0.01290 0.00527 -0.1099 0.8690 0.0726
-2.500 0.1979 0.01187 0.00433 -0.1090 0.8417 0.1202
-2.250 0.2238 0.01157 0.00423 -0.1084 0.8265 0.2112
-2.000 0.2503 0.01158 0.00417 -0.1079 0.8106 0.2420
-1.750 0.2767 0.01161 0.00411 -0.1074 0.7947 0.2657
-1.500 0.3030 0.01163 0.00403 -0.1069 0.7793 0.2885
-1.250 0.3290 0.01154 0.00392 -0.1065 0.7649 0.3109
-1.000 0.3551 0.01142 0.00375 -0.1060 0.7515 0.3328
-0.750 0.3808 0.01124 0.00361 -0.1056 0.7390 0.3601
-0.500 0.4063 0.01105 0.00351 -0.1051 0.7278 0.3954
-0.250 0.4314 0.01084 0.00345 -0.1046 0.7172 0.4444
0.000 0.4512 0.01013 0.00346 -0.1030 0.7074 0.6392
0.250 0.5108 0.00962 0.00335 -0.1092 0.6985 1.0000
0.500 0.5367 0.00976 0.00335 -0.1088 0.6899 1.0000
0.750 0.5625 0.00993 0.00341 -0.1084 0.6822 1.0000
1.000 0.5887 0.01009 0.00346 -0.1080 0.6750 1.0000
1.250 0.6144 0.01027 0.00357 -0.1076 0.6678 1.0000
1.500 0.6407 0.01046 0.00367 -0.1073 0.6612 1.0000
1.750 0.6665 0.01066 0.00384 -0.1069 0.6546 1.0000
2.000 0.6926 0.01086 0.00398 -0.1066 0.6480 1.0000
2.250 0.7187 0.01108 0.00417 -0.1063 0.6418 1.0000
2.500 0.7445 0.01128 0.00438 -0.1059 0.6354 1.0000
2.750 0.7715 0.01154 0.00460 -0.1057 0.6302 1.0000
3.000 0.7966 0.01175 0.00488 -0.1053 0.6240 1.0000
3.250 0.8231 0.01199 0.00512 -0.1050 0.6184 1.0000
3.500 0.8477 0.01219 0.00536 -0.1044 0.6095 1.0000
3.750 0.8720 0.01234 0.00551 -0.1035 0.5974 1.0000
4.000 0.8962 0.01247 0.00564 -0.1026 0.5850 1.0000
4.250 0.9189 0.01254 0.00572 -0.1014 0.5694 1.0000
4.500 0.9404 0.01255 0.00575 -0.0999 0.5502 1.0000
4.750 0.9613 0.01255 0.00576 -0.0982 0.5280 1.0000
5.000 0.9787 0.01251 0.00572 -0.0958 0.4898 1.0000
5.250 0.9766 0.01359 0.00589 -0.0901 0.3064 1.0000
5.500 0.9634 0.01679 0.00781 -0.0840 0.0775 1.0000
5.750 0.9772 0.01786 0.00889 -0.0816 0.0592 1.0000
6.000 0.9858 0.01914 0.01023 -0.0784 0.0506 1.0000
6.250 0.9985 0.02007 0.01128 -0.0759 0.0468 1.0000
6.500 1.0079 0.02118 0.01247 -0.0729 0.0435 1.0000
6.750 1.0123 0.02251 0.01381 -0.0693 0.0403 1.0000
7.000 1.0173 0.02407 0.01540 -0.0657 0.0378 1.0000
7.250 1.0310 0.02530 0.01670 -0.0634 0.0364 1.0000
7.500 1.0495 0.02677 0.01822 -0.0619 0.0350 1.0000
7.750 1.0776 0.02856 0.02007 -0.0618 0.0337 1.0000
8.000 1.1038 0.03013 0.02169 -0.0618 0.0312 1.0000
8.250 1.1409 0.03277 0.02436 -0.0635 0.0295 1.0000
8.500 1.1789 0.03572 0.02753 -0.0644 0.0301 1.0000
8.750 1.2560 0.04599 0.03909 -0.0645 0.0720 1.0000
9.000 1.2711 0.04871 0.04186 -0.0633 0.0659 1.0000
9.250 1.2693 0.05103 0.04459 -0.0597 0.0583 1.0000
9.500 1.2785 0.05365 0.04739 -0.0576 0.0549 1.0000
9.750 1.2939 0.05809 0.05186 -0.0572 0.0525 1.0000
10.000 1.2908 0.06796 0.06191 -0.0562 0.0504 1.0000
10.250 1.2742 0.06362 0.05816 -0.0489 0.0475 1.0000
10.500 1.2670 0.06582 0.06063 -0.0452 0.0455 1.0000
10.750 1.2579 0.06854 0.06356 -0.0419 0.0442 1.0000
11.000 1.2451 0.07120 0.06638 -0.0382 0.0434 1.0000
11.250 1.2308 0.07405 0.06939 -0.0352 0.0427 1.0000
11.500 1.2163 0.07725 0.07272 -0.0329 0.0422 1.0000
11.750 1.2007 0.08075 0.07637 -0.0312 0.0418 1.0000
12.000 1.1844 0.08459 0.08033 -0.0303 0.0414 1.0000
12.250 1.1669 0.08881 0.08469 -0.0300 0.0411 1.0000
12.500 1.1456 0.09346 0.08948 -0.0305 0.0410 1.0000
12.750 1.1195 0.09872 0.09491 -0.0321 0.0412 1.0000
13.000 1.0867 0.10536 0.10173 -0.0355 0.0417 1.0000
13.250 1.0523 0.11363 0.11016 -0.0407 0.0424 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 117 (MVA MK.4) AIRFOIL (goe117-il)