GOE 406 AIRFOIL (goe406-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 406 AIRFOIL (goe406-il) Reynolds number: 1,000,000 Max Cl/Cd: 114.24 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe406-il-1000000.txt Download as CSV file: xf-goe406-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 406 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.1786 0.10801 0.10626 -0.0636 0.9873 0.0192
-10.500 -0.1699 0.10337 0.10160 -0.0692 0.9856 0.0193
-10.250 -0.1579 0.09882 0.09705 -0.0745 0.9842 0.0194
-10.000 -0.1407 0.09472 0.09295 -0.0768 0.9833 0.0195
-9.750 -0.1326 0.09207 0.09031 -0.0773 0.9780 0.0196
-9.500 -0.1160 0.08929 0.08753 -0.0796 0.9749 0.0198
-9.250 -0.0970 0.08563 0.08387 -0.0841 0.9714 0.0199
-9.000 -0.0884 0.08336 0.08160 -0.0847 0.9613 0.0202
-8.750 -0.0726 0.07998 0.07821 -0.0885 0.9519 0.0204
-8.500 -0.0463 0.07602 0.07421 -0.0953 0.9352 0.0208
-8.250 0.0028 0.07036 0.06839 -0.1092 0.8957 0.0217
-8.000 0.0389 0.06176 0.05931 -0.1310 0.8235 0.0228
-7.750 0.0655 0.04490 0.04232 -0.1230 0.7721 0.0231
-7.500 0.0564 0.04256 0.03994 -0.1220 0.7627 0.0231
-7.250 0.0366 0.05328 0.05054 -0.1306 0.7752 0.0231
-7.000 0.0422 0.05125 0.04843 -0.1299 0.7658 0.0232
-6.750 0.0487 0.04915 0.04628 -0.1292 0.7570 0.0233
-6.500 0.0573 0.04749 0.04455 -0.1281 0.7488 0.0234
-6.250 0.0662 0.04552 0.04253 -0.1272 0.7416 0.0236
-6.000 0.0754 0.04350 0.04044 -0.1262 0.7347 0.0238
-5.750 0.0854 0.04152 0.03836 -0.1250 0.7279 0.0241
-5.500 0.0961 0.03937 0.03614 -0.1238 0.7218 0.0245
-5.250 0.1071 0.03721 0.03387 -0.1222 0.7152 0.0253
-5.000 0.1151 0.03205 0.02832 -0.1187 0.7101 0.0267
-4.750 0.1171 0.02896 0.02511 -0.1155 0.7043 0.0271
-4.500 0.1309 0.02765 0.02370 -0.1135 0.6979 0.0272
-4.250 0.1468 0.02665 0.02262 -0.1117 0.6913 0.0275
-4.000 0.1628 0.02551 0.02140 -0.1098 0.6848 0.0277
-3.750 0.1787 0.02433 0.02009 -0.1077 0.6794 0.0281
-3.500 0.1955 0.02307 0.01871 -0.1057 0.6744 0.0287
-3.250 0.2113 0.02096 0.01618 -0.1018 0.6695 0.0310
-3.000 0.2179 0.01813 0.01309 -0.0978 0.6647 0.0316
-2.750 0.2380 0.01751 0.01242 -0.0964 0.6596 0.0319
-2.500 0.2587 0.01687 0.01174 -0.0950 0.6544 0.0324
-2.250 0.2788 0.01623 0.01101 -0.0934 0.6483 0.0329
-2.000 0.2989 0.01551 0.01015 -0.0916 0.6428 0.0341
-1.750 0.3224 0.01564 0.00999 -0.0897 0.6370 0.0361
-1.500 0.3395 0.01369 0.00799 -0.0879 0.6314 0.0377
-1.250 0.3619 0.01327 0.00751 -0.0867 0.6256 0.0386
-1.000 0.3869 0.01141 0.00515 -0.0848 0.6198 0.0313
-0.750 0.4115 0.01080 0.00450 -0.0840 0.6122 0.0305
-0.500 0.4368 0.01029 0.00394 -0.0833 0.6058 0.0302
-0.250 0.4605 0.01012 0.00370 -0.0823 0.5971 0.0307
0.000 0.4845 0.00988 0.00342 -0.0813 0.5893 0.0308
0.250 0.5077 0.00962 0.00311 -0.0802 0.5802 0.0309
0.500 0.5309 0.00946 0.00292 -0.0791 0.5707 0.0310
0.750 0.5513 0.00914 0.00255 -0.0773 0.5598 0.0316
1.000 0.5725 0.00900 0.00238 -0.0758 0.5467 0.0322
1.250 0.5930 0.00896 0.00229 -0.0742 0.5307 0.0328
1.500 0.6123 0.00898 0.00222 -0.0722 0.5108 0.0333
1.750 0.6313 0.00903 0.00219 -0.0703 0.4882 0.0340
2.000 0.6473 0.00916 0.00220 -0.0677 0.4611 0.0348
2.250 0.6633 0.00932 0.00224 -0.0652 0.4338 0.0352
2.500 0.6796 0.00948 0.00229 -0.0627 0.4101 0.0364
2.750 0.6969 0.00961 0.00232 -0.0604 0.3902 0.0382
3.000 0.7151 0.00975 0.00239 -0.0584 0.3746 0.0401
3.250 0.7338 0.00989 0.00248 -0.0564 0.3623 0.0427
3.500 0.7536 0.00995 0.00261 -0.0547 0.3528 0.0821
3.750 1.0673 0.00956 0.00397 -0.1209 0.3249 1.0000
4.000 1.0891 0.00973 0.00411 -0.1197 0.3201 1.0000
4.250 1.1098 0.00994 0.00427 -0.1183 0.3148 1.0000
4.500 1.1326 0.01006 0.00439 -0.1173 0.3121 1.0000
4.750 1.1548 0.01019 0.00453 -0.1162 0.3088 1.0000
5.000 1.1763 0.01036 0.00468 -0.1149 0.3055 1.0000
5.250 1.1965 0.01057 0.00486 -0.1134 0.3009 1.0000
5.500 1.2170 0.01076 0.00504 -0.1120 0.2970 1.0000
5.750 1.2393 0.01087 0.00517 -0.1109 0.2932 1.0000
6.000 1.2601 0.01103 0.00532 -0.1096 0.2896 1.0000
6.250 1.2797 0.01123 0.00551 -0.1080 0.2859 1.0000
6.500 1.2979 0.01147 0.00573 -0.1062 0.2812 1.0000
6.750 1.3192 0.01159 0.00589 -0.1049 0.2792 1.0000
7.000 1.3386 0.01173 0.00605 -0.1033 0.2763 1.0000
7.250 1.3553 0.01190 0.00622 -0.1011 0.2724 1.0000
7.500 1.3699 0.01211 0.00643 -0.0986 0.2684 1.0000
7.750 1.3864 0.01229 0.00662 -0.0964 0.2643 1.0000
8.000 1.4045 0.01243 0.00679 -0.0945 0.2615 1.0000
8.250 1.4212 0.01262 0.00699 -0.0925 0.2578 1.0000
8.500 1.4363 0.01287 0.00724 -0.0901 0.2529 1.0000
8.750 1.4530 0.01309 0.00748 -0.0881 0.2486 1.0000
9.000 1.4705 0.01329 0.00770 -0.0863 0.2439 1.0000
9.250 1.4855 0.01358 0.00798 -0.0841 0.2383 1.0000
9.500 1.5016 0.01385 0.00827 -0.0820 0.2327 1.0000
9.750 1.5153 0.01421 0.00860 -0.0796 0.2228 1.0000
10.000 1.5284 0.01463 0.00898 -0.0772 0.2104 1.0000
10.250 1.5357 0.01529 0.00954 -0.0739 0.1904 1.0000
10.500 1.5288 0.01664 0.01062 -0.0684 0.1474 1.0000
10.750 1.5159 0.01841 0.01213 -0.0623 0.1090 1.0000
11.000 1.5085 0.02006 0.01360 -0.0574 0.0788 1.0000
11.250 1.4959 0.02214 0.01548 -0.0523 0.0471 1.0000
11.500 1.4996 0.02339 0.01673 -0.0495 0.0402 1.0000
11.750 1.5055 0.02457 0.01792 -0.0472 0.0365 1.0000
12.000 1.5127 0.02572 0.01910 -0.0452 0.0342 1.0000
12.250 1.5176 0.02710 0.02052 -0.0431 0.0322 1.0000
12.500 1.5245 0.02837 0.02185 -0.0414 0.0307 1.0000
12.750 1.5318 0.02966 0.02320 -0.0398 0.0295 1.0000
13.000 1.5372 0.03114 0.02471 -0.0381 0.0282 1.0000
13.250 1.5400 0.03289 0.02651 -0.0364 0.0271 1.0000
13.500 1.5428 0.03468 0.02837 -0.0349 0.0260 1.0000
13.750 1.5483 0.03628 0.03004 -0.0336 0.0255 1.0000
14.000 1.5530 0.03801 0.03183 -0.0323 0.0247 1.0000
14.250 1.5565 0.03989 0.03376 -0.0312 0.0237 1.0000
14.500 1.5575 0.04205 0.03599 -0.0300 0.0232 1.0000
14.750 1.5523 0.04489 0.03890 -0.0287 0.0222 1.0000
15.000 1.5495 0.04760 0.04169 -0.0278 0.0218 1.0000
15.250 1.5525 0.04977 0.04393 -0.0271 0.0213 1.0000
15.500 1.5533 0.05221 0.04645 -0.0264 0.0207 1.0000
15.750 1.5512 0.05505 0.04937 -0.0258 0.0202 1.0000
16.000 1.5499 0.05787 0.05226 -0.0254 0.0197 1.0000
16.250 1.5441 0.06134 0.05582 -0.0251 0.0195 1.0000
16.500 1.5378 0.06492 0.05949 -0.0250 0.0191 1.0000
16.750 1.5284 0.06902 0.06367 -0.0250 0.0188 1.0000
17.000 1.5133 0.07403 0.06879 -0.0254 0.0183 1.0000
17.250 1.4986 0.07912 0.07399 -0.0260 0.0181 1.0000
17.500 1.4930 0.08299 0.07795 -0.0264 0.0178 1.0000
17.750 1.4854 0.08720 0.08225 -0.0270 0.0176 1.0000
18.000 1.4754 0.09184 0.08700 -0.0278 0.0174 1.0000
18.250 1.4668 0.09632 0.09157 -0.0287 0.0171 1.0000
18.500 1.4555 0.10121 0.09655 -0.0297 0.0168 1.0000
18.750 1.4460 0.10588 0.10131 -0.0308 0.0167 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 406 AIRFOIL (goe406-il)