GOE 515 AIRFOIL (goe515-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 515 AIRFOIL (goe515-il) Reynolds number: 500,000 Max Cl/Cd: 87.99 at α=2.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe515-il-500000-n5.txt Download as CSV file: xf-goe515-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 515 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.7149 0.07786 0.07532 -0.0415 1.0000 0.0097
-12.250 -0.7795 0.06457 0.06201 -0.0505 1.0000 0.0090
-12.000 -0.9193 0.03870 0.03567 -0.0656 1.0000 0.0077
-11.750 -0.9301 0.03208 0.02867 -0.0656 1.0000 0.0077
-11.500 -0.9288 0.02903 0.02538 -0.0640 1.0000 0.0079
-11.250 -0.9219 0.02697 0.02314 -0.0621 1.0000 0.0083
-11.000 -0.9126 0.02534 0.02133 -0.0601 1.0000 0.0085
-10.750 -0.9015 0.02400 0.01984 -0.0581 1.0000 0.0090
-10.500 -0.8751 0.02248 0.01808 -0.0592 0.9976 0.0095
-10.250 -0.8478 0.02116 0.01655 -0.0602 0.9948 0.0100
-10.000 -0.8214 0.01994 0.01513 -0.0609 0.9917 0.0106
-9.750 -0.7941 0.01873 0.01376 -0.0617 0.9885 0.0115
-9.500 -0.7645 0.01782 0.01275 -0.0627 0.9861 0.0125
-9.250 -0.7361 0.01707 0.01186 -0.0633 0.9831 0.0136
-9.000 -0.7078 0.01633 0.01100 -0.0639 0.9795 0.0151
-8.750 -0.6769 0.01576 0.01038 -0.0649 0.9769 0.0170
-8.500 -0.6446 0.01530 0.00983 -0.0661 0.9749 0.0191
-8.250 -0.6162 0.01479 0.00921 -0.0664 0.9703 0.0206
-8.000 -0.5856 0.01433 0.00871 -0.0673 0.9663 0.0226
-7.750 -0.5527 0.01398 0.00830 -0.0685 0.9631 0.0247
-7.500 -0.5238 0.01367 0.00789 -0.0688 0.9573 0.0264
-7.250 -0.4929 0.01338 0.00749 -0.0695 0.9524 0.0273
-7.000 -0.4627 0.01271 0.00675 -0.0703 0.9482 0.0295
-6.750 -0.4359 0.01234 0.00634 -0.0701 0.9411 0.0313
-6.500 -0.4061 0.01201 0.00594 -0.0706 0.9355 0.0331
-6.250 -0.3784 0.01171 0.00556 -0.0706 0.9287 0.0346
-6.000 -0.3502 0.01144 0.00519 -0.0706 0.9218 0.0357
-5.750 -0.3234 0.01105 0.00473 -0.0704 0.9146 0.0373
-5.500 -0.2965 0.01069 0.00432 -0.0702 0.9070 0.0395
-5.250 -0.2697 0.01041 0.00399 -0.0699 0.8995 0.0411
-5.000 -0.2427 0.01016 0.00367 -0.0696 0.8921 0.0431
-4.750 -0.2160 0.00994 0.00338 -0.0693 0.8847 0.0452
-4.500 -0.1892 0.00971 0.00311 -0.0690 0.8773 0.0493
-4.250 -0.1628 0.00950 0.00289 -0.0686 0.8702 0.0560
-4.000 -0.1361 0.00931 0.00272 -0.0683 0.8632 0.0665
-3.750 -0.1093 0.00918 0.00261 -0.0680 0.8564 0.0774
-3.500 -0.0823 0.00909 0.00248 -0.0677 0.8495 0.0844
-3.250 -0.0553 0.00899 0.00237 -0.0675 0.8433 0.0904
-3.000 -0.0283 0.00892 0.00225 -0.0672 0.8365 0.0947
-2.750 -0.0012 0.00885 0.00213 -0.0669 0.8304 0.0979
-2.500 0.0255 0.00873 0.00201 -0.0666 0.8236 0.1021
-2.250 0.0524 0.00866 0.00191 -0.0662 0.8171 0.1064
-2.000 0.0788 0.00861 0.00181 -0.0658 0.8064 0.1103
-1.750 0.1048 0.00854 0.00170 -0.0653 0.7930 0.1144
-1.500 0.1308 0.00846 0.00161 -0.0647 0.7799 0.1203
-1.250 0.1574 0.00841 0.00153 -0.0644 0.7711 0.1259
-1.000 0.1839 0.00832 0.00146 -0.0640 0.7623 0.1326
-0.750 0.2106 0.00826 0.00140 -0.0636 0.7541 0.1394
-0.500 0.2370 0.00819 0.00135 -0.0632 0.7448 0.1504
-0.250 0.2634 0.00811 0.00133 -0.0629 0.7358 0.1700
0.000 0.2896 0.00802 0.00131 -0.0624 0.7263 0.1981
0.250 0.3157 0.00794 0.00130 -0.0620 0.7154 0.2254
0.500 0.3417 0.00787 0.00130 -0.0616 0.7043 0.2553
0.750 0.3673 0.00779 0.00130 -0.0610 0.6913 0.2922
1.000 0.3916 0.00767 0.00131 -0.0603 0.6734 0.3528
1.250 0.4139 0.00742 0.00133 -0.0592 0.6536 0.4781
1.750 0.5102 0.00651 0.00163 -0.0679 0.6078 0.9775
2.000 0.5511 0.00668 0.00172 -0.0708 0.5856 0.9919
2.250 0.5936 0.00686 0.00179 -0.0740 0.5562 0.9994
2.500 0.6189 0.00704 0.00186 -0.0735 0.5298 1.0000
2.750 0.6406 0.00728 0.00196 -0.0722 0.4950 1.0000
3.000 0.6614 0.00759 0.00209 -0.0708 0.4540 1.0000
3.250 0.6819 0.00795 0.00226 -0.0693 0.4126 1.0000
3.500 0.7030 0.00828 0.00244 -0.0680 0.3783 1.0000
3.750 0.7245 0.00859 0.00262 -0.0667 0.3485 1.0000
4.000 0.7463 0.00888 0.00281 -0.0655 0.3215 1.0000
4.250 0.7679 0.00919 0.00301 -0.0643 0.2941 1.0000
4.500 0.7887 0.00956 0.00324 -0.0630 0.2569 1.0000
4.750 0.8071 0.01013 0.00354 -0.0613 0.1982 1.0000
5.000 0.8263 0.01065 0.00387 -0.0597 0.1601 1.0000
5.250 0.8469 0.01106 0.00418 -0.0584 0.1400 1.0000
5.500 0.8680 0.01142 0.00447 -0.0571 0.1224 1.0000
5.750 0.8894 0.01177 0.00477 -0.0559 0.1075 1.0000
6.000 0.9106 0.01212 0.00507 -0.0547 0.0909 1.0000
6.250 0.9305 0.01258 0.00540 -0.0533 0.0689 1.0000
6.500 0.9505 0.01302 0.00577 -0.0519 0.0545 1.0000
6.750 0.9709 0.01344 0.00618 -0.0505 0.0451 1.0000
7.000 0.9908 0.01389 0.00659 -0.0491 0.0362 1.0000
7.250 1.0103 0.01436 0.00703 -0.0477 0.0270 1.0000
7.500 1.0295 0.01486 0.00749 -0.0462 0.0202 1.0000
7.750 1.0489 0.01534 0.00798 -0.0447 0.0172 1.0000
8.000 1.0672 0.01590 0.00856 -0.0430 0.0145 1.0000
8.250 1.0867 0.01635 0.00909 -0.0416 0.0136 1.0000
8.500 1.1054 0.01685 0.00965 -0.0400 0.0124 1.0000
8.750 1.1229 0.01744 0.01027 -0.0384 0.0114 1.0000
9.000 1.1382 0.01818 0.01107 -0.0363 0.0104 1.0000
9.250 1.1542 0.01883 0.01180 -0.0344 0.0100 1.0000
9.500 1.1704 0.01944 0.01248 -0.0325 0.0096 1.0000
9.750 1.1843 0.02011 0.01323 -0.0303 0.0091 1.0000
10.000 1.1974 0.02074 0.01394 -0.0280 0.0086 1.0000
10.250 1.2086 0.02148 0.01476 -0.0254 0.0084 1.0000
10.500 1.2213 0.02215 0.01548 -0.0232 0.0079 1.0000
10.750 1.2311 0.02305 0.01644 -0.0207 0.0076 1.0000
11.000 1.2369 0.02425 0.01773 -0.0178 0.0073 1.0000
11.250 1.2452 0.02532 0.01889 -0.0154 0.0071 1.0000
11.500 1.2556 0.02626 0.01994 -0.0134 0.0069 1.0000
11.750 1.2635 0.02742 0.02121 -0.0113 0.0067 1.0000
12.000 1.2699 0.02874 0.02265 -0.0092 0.0066 1.0000
12.250 1.2762 0.03013 0.02415 -0.0072 0.0063 1.0000
12.500 1.2825 0.03154 0.02569 -0.0055 0.0061 1.0000
12.750 1.2862 0.03324 0.02751 -0.0038 0.0060 1.0000
13.000 1.2905 0.03493 0.02931 -0.0023 0.0059 1.0000
13.250 1.2934 0.03684 0.03135 -0.0010 0.0058 1.0000
13.500 1.2974 0.03867 0.03327 0.0001 0.0056 1.0000
13.750 1.2969 0.04107 0.03580 0.0011 0.0056 1.0000
14.000 1.2964 0.04356 0.03840 0.0017 0.0054 1.0000
14.250 1.2953 0.04623 0.04119 0.0021 0.0053 1.0000
14.500 1.2925 0.04924 0.04431 0.0022 0.0053 1.0000
14.750 1.2861 0.05284 0.04806 0.0020 0.0052 1.0000
15.000 1.2799 0.05660 0.05194 0.0013 0.0051 1.0000
15.250 1.2660 0.06164 0.05713 0.0000 0.0050 1.0000
15.500 1.2590 0.06600 0.06164 -0.0014 0.0050 1.0000
15.750 1.2435 0.07194 0.06775 -0.0036 0.0050 1.0000
16.000 1.2346 0.07722 0.07318 -0.0060 0.0049 1.0000
16.250 1.2210 0.08352 0.07965 -0.0090 0.0049 1.0000
16.500 1.2083 0.08991 0.08619 -0.0122 0.0049 1.0000
16.750 1.1911 0.09733 0.09377 -0.0160 0.0049 1.0000
17.000 1.1745 0.10489 0.10147 -0.0200 0.0049 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 515 AIRFOIL (goe515-il)