GOE 407 AIRFOIL (goe407-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 407 AIRFOIL (goe407-il) Reynolds number: 1,000,000 Max Cl/Cd: 121.64 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe407-il-1000000.txt Download as CSV file: xf-goe407-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 407 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 -0.2326 0.10871 0.10713 -0.0368 0.9968 0.0177
-10.500 -0.2319 0.10384 0.10226 -0.0405 0.9959 0.0179
-10.250 -0.2266 0.09922 0.09763 -0.0438 0.9950 0.0179
-10.000 -0.2174 0.09430 0.09271 -0.0457 0.9943 0.0181
-9.750 -0.2034 0.09117 0.08958 -0.0469 0.9933 0.0183
-9.500 -0.1933 0.08785 0.08626 -0.0486 0.9917 0.0184
-9.250 -0.1823 0.08457 0.08298 -0.0505 0.9901 0.0188
-9.000 -0.1721 0.08083 0.07925 -0.0530 0.9885 0.0190
-8.750 -0.1613 0.07701 0.07543 -0.0558 0.9870 0.0195
-8.500 -0.1497 0.07299 0.07141 -0.0590 0.9855 0.0201
-8.250 -0.1469 0.06701 0.06541 -0.0661 0.9838 0.0214
-8.000 -0.1451 0.06264 0.06105 -0.0694 0.9792 0.0215
-7.750 -0.1395 0.05786 0.05627 -0.0745 0.9748 0.0216
-7.500 -0.1280 0.05268 0.05107 -0.0807 0.9715 0.0216
-7.250 -0.1272 0.04761 0.04598 -0.0827 0.9638 0.0219
-7.000 -0.1071 0.04428 0.04263 -0.0859 0.9590 0.0221
-6.750 -0.0895 0.04083 0.03915 -0.0892 0.9501 0.0222
-6.500 -0.0659 0.03707 0.03533 -0.0939 0.9395 0.0225
-6.250 -0.0416 0.03335 0.03152 -0.0985 0.9155 0.0230
-6.000 -0.0150 0.02964 0.02752 -0.1032 0.8596 0.0237
-5.250 0.0055 0.01744 0.01444 -0.1017 0.7859 0.0261
-5.000 0.0179 0.01593 0.01280 -0.1002 0.7684 0.0264
-4.750 0.0326 0.01467 0.01143 -0.0988 0.7512 0.0266
-4.500 0.0484 0.01346 0.01008 -0.0975 0.7352 0.0269
-4.250 0.0648 0.01225 0.00873 -0.0962 0.7202 0.0273
-4.000 0.0818 0.01108 0.00740 -0.0948 0.7074 0.0279
-3.750 0.0994 0.00976 0.00590 -0.0933 0.6966 0.0287
-3.500 0.1081 0.00667 0.00217 -0.0895 0.6900 0.0316
-3.250 0.1289 0.00621 0.00166 -0.0885 0.6805 0.0320
-3.000 0.1497 0.00579 0.00113 -0.0874 0.6726 0.0324
-2.750 0.1710 0.00537 0.00060 -0.0863 0.6648 0.0332
-2.500 0.1926 0.00496 0.00004 -0.0851 0.6581 0.0344
-2.000 0.2401 0.01383 0.00824 -0.0848 0.6528 0.0383
-1.750 0.2636 0.01335 0.00771 -0.0839 0.6457 0.0390
-1.500 0.2884 0.01161 0.00555 -0.0821 0.6401 0.0307
-0.750 0.3637 0.00989 0.00358 -0.0798 0.6225 0.0306
-0.500 0.3893 0.00972 0.00336 -0.0792 0.6170 0.0309
0.000 0.4385 0.00903 0.00263 -0.0776 0.6061 0.0312
0.250 0.4623 0.00874 0.00232 -0.0766 0.6000 0.0315
0.500 0.4857 0.00855 0.00210 -0.0755 0.5940 0.0322
0.750 0.5107 0.00839 0.00195 -0.0748 0.5882 0.0331
1.000 0.5351 0.00830 0.00184 -0.0739 0.5815 0.0336
1.250 0.5598 0.00823 0.00176 -0.0731 0.5748 0.0346
1.500 0.5843 0.00817 0.00169 -0.0722 0.5656 0.0355
1.750 0.6088 0.00814 0.00163 -0.0714 0.5553 0.0367
2.000 0.6326 0.00812 0.00158 -0.0704 0.5435 0.0380
2.250 0.6562 0.00813 0.00155 -0.0694 0.5291 0.0413
2.500 0.6791 0.00814 0.00157 -0.0683 0.5089 0.0585
2.750 0.6987 0.00812 0.00166 -0.0665 0.4800 0.1609
3.000 0.9115 0.00758 0.00260 -0.1097 0.4084 1.0000
3.250 0.9339 0.00777 0.00270 -0.1085 0.3921 1.0000
3.500 0.9564 0.00796 0.00281 -0.1074 0.3773 1.0000
3.750 0.9789 0.00814 0.00293 -0.1063 0.3634 1.0000
4.000 1.0017 0.00831 0.00304 -0.1053 0.3526 1.0000
4.250 1.0240 0.00849 0.00317 -0.1042 0.3418 1.0000
4.500 1.0470 0.00864 0.00329 -0.1032 0.3324 1.0000
4.750 1.0694 0.00881 0.00343 -0.1021 0.3230 1.0000
5.000 1.0916 0.00900 0.00357 -0.1009 0.3133 1.0000
5.250 1.1142 0.00916 0.00371 -0.0999 0.3052 1.0000
5.500 1.1356 0.00937 0.00388 -0.0986 0.2952 1.0000
5.750 1.1580 0.00953 0.00403 -0.0975 0.2866 1.0000
6.000 1.1795 0.00974 0.00420 -0.0963 0.2776 1.0000
6.250 1.2007 0.00995 0.00438 -0.0950 0.2677 1.0000
6.500 1.2217 0.01016 0.00457 -0.0937 0.2569 1.0000
6.750 1.2419 0.01041 0.00477 -0.0922 0.2454 1.0000
7.000 1.2614 0.01069 0.00499 -0.0906 0.2329 1.0000
7.250 1.2811 0.01095 0.00522 -0.0891 0.2215 1.0000
7.500 1.2987 0.01130 0.00550 -0.0871 0.2038 1.0000
7.750 1.3136 0.01179 0.00584 -0.0848 0.1767 1.0000
8.000 1.3203 0.01268 0.00644 -0.0809 0.1301 1.0000
8.250 1.3334 0.01321 0.00690 -0.0782 0.1157 1.0000
8.500 1.3484 0.01362 0.00729 -0.0758 0.1076 1.0000
8.750 1.3649 0.01393 0.00762 -0.0737 0.1037 1.0000
9.000 1.3771 0.01432 0.00798 -0.0707 0.0981 1.0000
9.250 1.3897 0.01460 0.00829 -0.0678 0.0940 1.0000
9.500 1.4028 0.01487 0.00858 -0.0651 0.0901 1.0000
9.750 1.4134 0.01527 0.00897 -0.0619 0.0843 1.0000
10.000 1.4267 0.01561 0.00932 -0.0593 0.0785 1.0000
10.250 1.4370 0.01609 0.00975 -0.0563 0.0699 1.0000
10.500 1.4456 0.01669 0.01028 -0.0531 0.0578 1.0000
10.750 1.4480 0.01762 0.01107 -0.0489 0.0400 1.0000
11.000 1.4455 0.01882 0.01216 -0.0442 0.0231 1.0000
11.250 1.4515 0.01966 0.01301 -0.0410 0.0193 1.0000
11.500 1.4593 0.02044 0.01383 -0.0381 0.0178 1.0000
11.750 1.4646 0.02141 0.01483 -0.0350 0.0162 1.0000
12.000 1.4728 0.02226 0.01575 -0.0325 0.0155 1.0000
12.250 1.4810 0.02315 0.01669 -0.0302 0.0147 1.0000
12.500 1.4880 0.02417 0.01778 -0.0278 0.0144 1.0000
12.750 1.4920 0.02544 0.01912 -0.0254 0.0137 1.0000
13.000 1.4939 0.02697 0.02072 -0.0230 0.0131 1.0000
13.250 1.4970 0.02851 0.02235 -0.0209 0.0126 1.0000
13.500 1.5019 0.02999 0.02390 -0.0193 0.0124 1.0000
13.750 1.5074 0.03149 0.02548 -0.0178 0.0121 1.0000
14.000 1.5106 0.03326 0.02733 -0.0163 0.0118 1.0000
14.250 1.5127 0.03519 0.02934 -0.0150 0.0115 1.0000
14.500 1.5142 0.03726 0.03148 -0.0139 0.0113 1.0000
14.750 1.5140 0.03958 0.03389 -0.0129 0.0110 1.0000
15.000 1.5130 0.04208 0.03648 -0.0121 0.0109 1.0000
15.250 1.5074 0.04516 0.03965 -0.0113 0.0106 1.0000
15.500 1.4993 0.04864 0.04323 -0.0108 0.0104 1.0000
15.750 1.4903 0.05234 0.04704 -0.0105 0.0102 1.0000
16.000 1.4788 0.05645 0.05127 -0.0105 0.0101 1.0000
16.250 1.4681 0.06066 0.05559 -0.0107 0.0100 1.0000
16.500 1.4653 0.06397 0.05898 -0.0109 0.0099 1.0000
16.750 1.4587 0.06785 0.06297 -0.0113 0.0098 1.0000
17.000 1.4547 0.07147 0.06669 -0.0119 0.0097 1.0000
17.250 1.4449 0.07606 0.07138 -0.0128 0.0096 1.0000
17.500 1.4370 0.08042 0.07584 -0.0138 0.0095 1.0000
17.750 1.4278 0.08508 0.08059 -0.0149 0.0094 1.0000
18.000 1.4176 0.08997 0.08559 -0.0162 0.0093 1.0000
18.250 1.4123 0.09416 0.08986 -0.0174 0.0091 1.0000
18.500 1.4001 0.09943 0.09523 -0.0190 0.0090 1.0000
18.750 1.3910 0.10426 0.10015 -0.0205 0.0090 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 407 AIRFOIL (goe407-il)