GOE 285 AIRFOIL (goe285-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 285 AIRFOIL (goe285-il) Reynolds number: 1,000,000 Max Cl/Cd: 103.52 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe285-il-1000000-n5.txt Download as CSV file: xf-goe285-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 285 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
------ -------- --------- --------- -------- -------- --------
-14.000 -0.9706 0.04075 0.03811 -0.0815 1.0000 0.0197
-13.750 -0.9965 0.03631 0.03345 -0.0813 1.0000 0.0198
-13.500 -1.0026 0.03368 0.03065 -0.0797 1.0000 0.0199
-13.250 -1.0026 0.03165 0.02846 -0.0777 1.0000 0.0200
-13.000 -0.9986 0.02994 0.02661 -0.0756 1.0000 0.0201
-12.750 -0.9918 0.02847 0.02499 -0.0735 1.0000 0.0202
-12.500 -0.9828 0.02717 0.02356 -0.0713 1.0000 0.0203
-12.250 -0.9621 0.02582 0.02207 -0.0714 0.9989 0.0204
-12.000 -0.9360 0.02464 0.02075 -0.0723 0.9971 0.0205
-11.750 -0.9081 0.02363 0.01962 -0.0733 0.9953 0.0207
-11.500 -0.8791 0.02270 0.01857 -0.0745 0.9938 0.0208
-11.250 -0.8546 0.02145 0.01718 -0.0749 0.9909 0.0209
-11.000 -0.8290 0.02013 0.01573 -0.0755 0.9878 0.0211
-10.750 -0.8008 0.01918 0.01470 -0.0763 0.9856 0.0213
-10.500 -0.7715 0.01840 0.01385 -0.0771 0.9836 0.0215
-10.250 -0.7460 0.01776 0.01314 -0.0770 0.9786 0.0216
-10.000 -0.7175 0.01713 0.01245 -0.0774 0.9743 0.0218
-9.750 -0.6892 0.01655 0.01180 -0.0778 0.9695 0.0219
-9.500 -0.6617 0.01599 0.01119 -0.0779 0.9629 0.0221
-9.250 -0.6334 0.01546 0.01058 -0.0782 0.9562 0.0223
-9.000 -0.6055 0.01496 0.01002 -0.0783 0.9468 0.0225
-8.750 -0.5781 0.01451 0.00949 -0.0782 0.9338 0.0227
-8.500 -0.5515 0.01410 0.00898 -0.0779 0.9158 0.0230
-8.250 -0.5261 0.01372 0.00847 -0.0774 0.8941 0.0232
-8.000 -0.5008 0.01335 0.00798 -0.0768 0.8738 0.0233
-7.750 -0.4751 0.01301 0.00752 -0.0763 0.8582 0.0235
-7.500 -0.4491 0.01268 0.00708 -0.0759 0.8444 0.0237
-7.250 -0.4228 0.01237 0.00667 -0.0755 0.8314 0.0239
-7.000 -0.3962 0.01207 0.00629 -0.0752 0.8200 0.0240
-6.750 -0.3695 0.01180 0.00593 -0.0748 0.8090 0.0242
-6.500 -0.3426 0.01154 0.00560 -0.0746 0.7978 0.0243
-6.250 -0.3155 0.01129 0.00528 -0.0743 0.7872 0.0244
-6.000 -0.2886 0.01105 0.00496 -0.0740 0.7749 0.0246
-5.750 -0.2623 0.01070 0.00452 -0.0736 0.7586 0.0250
-5.500 -0.2356 0.01045 0.00419 -0.0733 0.7396 0.0253
-5.250 -0.2087 0.01025 0.00391 -0.0730 0.7176 0.0256
-5.000 -0.1817 0.01010 0.00366 -0.0727 0.6973 0.0259
-4.750 -0.1544 0.00995 0.00345 -0.0724 0.6805 0.0263
-4.500 -0.1270 0.00983 0.00325 -0.0722 0.6651 0.0267
-4.250 -0.0995 0.00971 0.00307 -0.0720 0.6504 0.0271
-4.000 -0.0720 0.00960 0.00289 -0.0718 0.6368 0.0275
-3.500 -0.0164 0.00939 0.00257 -0.0715 0.6130 0.0282
-3.250 0.0114 0.00931 0.00243 -0.0713 0.6014 0.0286
-3.000 0.0390 0.00923 0.00229 -0.0712 0.5871 0.0290
-2.750 0.0668 0.00913 0.00214 -0.0710 0.5726 0.0298
-2.500 0.0946 0.00906 0.00203 -0.0709 0.5581 0.0306
-2.250 0.1222 0.00903 0.00193 -0.0707 0.5379 0.0315
-2.000 0.1496 0.00903 0.00184 -0.0705 0.5132 0.0326
-1.500 0.2042 0.00907 0.00171 -0.0701 0.4657 0.0374
-1.250 0.2313 0.00900 0.00167 -0.0699 0.4481 0.0578
-1.000 0.2589 0.00900 0.00166 -0.0698 0.4339 0.0691
-0.750 0.2867 0.00903 0.00165 -0.0697 0.4218 0.0746
-0.500 0.3147 0.00904 0.00164 -0.0696 0.4122 0.0789
-0.250 0.3425 0.00907 0.00164 -0.0695 0.4026 0.0834
0.000 0.3707 0.00908 0.00164 -0.0694 0.3944 0.0872
0.250 0.3984 0.00910 0.00166 -0.0693 0.3853 0.0952
0.500 0.4264 0.00909 0.00167 -0.0693 0.3772 0.1072
0.750 0.4541 0.00910 0.00170 -0.0692 0.3679 0.1255
1.000 0.4819 0.00911 0.00173 -0.0691 0.3584 0.1425
1.250 0.5095 0.00914 0.00177 -0.0690 0.3480 0.1604
1.500 0.5368 0.00915 0.00182 -0.0689 0.3350 0.1921
2.000 0.5891 0.00883 0.00198 -0.0685 0.3043 0.4262
2.250 0.6138 0.00852 0.00209 -0.0681 0.2890 0.6058
2.500 0.6295 0.00778 0.00229 -0.0652 0.2783 0.9429
2.750 0.6742 0.00800 0.00244 -0.0688 0.2637 0.9907
3.000 0.7178 0.00824 0.00257 -0.0723 0.2491 0.9984
3.250 0.7481 0.00843 0.00269 -0.0729 0.2401 1.0000
3.500 0.7730 0.00855 0.00278 -0.0722 0.2357 1.0000
3.750 0.7977 0.00869 0.00288 -0.0716 0.2308 1.0000
4.000 0.8222 0.00886 0.00300 -0.0708 0.2250 1.0000
4.250 0.8473 0.00898 0.00311 -0.0702 0.2213 1.0000
4.500 0.8724 0.00912 0.00322 -0.0696 0.2177 1.0000
4.750 0.8976 0.00926 0.00334 -0.0691 0.2146 1.0000
5.000 0.9228 0.00943 0.00348 -0.0685 0.2109 1.0000
5.250 0.9483 0.00958 0.00362 -0.0680 0.2076 1.0000
5.500 0.9743 0.00970 0.00375 -0.0677 0.2052 1.0000
5.750 1.0001 0.00985 0.00389 -0.0673 0.2017 1.0000
6.000 1.0253 0.01005 0.00405 -0.0668 0.1958 1.0000
6.250 1.0510 0.01023 0.00422 -0.0664 0.1913 1.0000
6.500 1.0766 0.01040 0.00438 -0.0660 0.1850 1.0000
6.750 1.1015 0.01064 0.00457 -0.0655 0.1758 1.0000
7.000 1.1177 0.01161 0.00517 -0.0638 0.1146 1.0000
7.250 1.1398 0.01209 0.00556 -0.0629 0.0995 1.0000
7.500 1.1624 0.01251 0.00593 -0.0620 0.0883 1.0000
7.750 1.1842 0.01299 0.00633 -0.0611 0.0734 1.0000
8.000 1.2006 0.01391 0.00702 -0.0594 0.0389 1.0000
8.250 1.2228 0.01433 0.00742 -0.0586 0.0347 1.0000
8.500 1.2457 0.01468 0.00777 -0.0578 0.0329 1.0000
8.750 1.2683 0.01503 0.00813 -0.0571 0.0318 1.0000
9.000 1.2903 0.01541 0.00852 -0.0562 0.0308 1.0000
9.250 1.3117 0.01582 0.00894 -0.0553 0.0299 1.0000
9.500 1.3324 0.01626 0.00939 -0.0543 0.0288 1.0000
9.750 1.3534 0.01664 0.00981 -0.0533 0.0283 1.0000
10.000 1.3739 0.01704 0.01024 -0.0523 0.0279 1.0000
10.250 1.3935 0.01746 0.01069 -0.0511 0.0274 1.0000
10.500 1.4111 0.01790 0.01115 -0.0496 0.0268 1.0000
10.750 1.4275 0.01838 0.01167 -0.0479 0.0263 1.0000
11.000 1.4433 0.01890 0.01222 -0.0461 0.0258 1.0000
11.250 1.4583 0.01949 0.01285 -0.0444 0.0253 1.0000
11.500 1.4720 0.02017 0.01356 -0.0426 0.0246 1.0000
11.750 1.4844 0.02095 0.01439 -0.0407 0.0241 1.0000
12.000 1.4994 0.02159 0.01507 -0.0392 0.0238 1.0000
12.250 1.5137 0.02229 0.01582 -0.0377 0.0235 1.0000
12.500 1.5274 0.02305 0.01663 -0.0363 0.0231 1.0000
12.750 1.5403 0.02390 0.01754 -0.0349 0.0227 1.0000
13.000 1.5524 0.02485 0.01853 -0.0336 0.0223 1.0000
13.250 1.5636 0.02589 0.01963 -0.0323 0.0219 1.0000
13.500 1.5740 0.02705 0.02084 -0.0311 0.0215 1.0000
13.750 1.5833 0.02835 0.02219 -0.0299 0.0211 1.0000
14.000 1.5908 0.02985 0.02374 -0.0288 0.0206 1.0000
14.250 1.5959 0.03163 0.02559 -0.0278 0.0202 1.0000
14.500 1.6034 0.03328 0.02730 -0.0270 0.0200 1.0000
14.750 1.6103 0.03502 0.02912 -0.0263 0.0197 1.0000
15.000 1.6159 0.03695 0.03113 -0.0257 0.0195 1.0000
15.250 1.6206 0.03906 0.03331 -0.0253 0.0191 1.0000
15.500 1.6244 0.04131 0.03564 -0.0250 0.0188 1.0000
15.750 1.6265 0.04382 0.03823 -0.0248 0.0184 1.0000
16.000 1.6275 0.04653 0.04101 -0.0248 0.0181 1.0000
16.250 1.6270 0.04949 0.04405 -0.0249 0.0177 1.0000
16.500 1.6241 0.05281 0.04745 -0.0252 0.0175 1.0000
16.750 1.6194 0.05644 0.05116 -0.0256 0.0172 1.0000
17.000 1.6125 0.06047 0.05529 -0.0263 0.0170 1.0000
17.250 1.6027 0.06499 0.05991 -0.0272 0.0167 1.0000
17.500 1.5943 0.06942 0.06445 -0.0282 0.0166 1.0000
17.750 1.5846 0.07415 0.06928 -0.0294 0.0164 1.0000
18.000 1.5736 0.07910 0.07435 -0.0307 0.0163 1.0000
18.250 1.5610 0.08434 0.07969 -0.0322 0.0161 1.0000
18.500 1.5485 0.08962 0.08508 -0.0337 0.0159 1.0000
18.750 1.5358 0.09497 0.09054 -0.0354 0.0157 1.0000
19.000 1.5226 0.10041 0.09607 -0.0371 0.0156 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 285 AIRFOIL (goe285-il)