GOE 515 AIRFOIL (goe515-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: GOE 515 AIRFOIL (goe515-il) Reynolds number: 100,000 Max Cl/Cd: 56.78 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe515-il-100000.txt Download as CSV file: xf-goe515-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4139 0.09289 0.08778 -0.0294 1.0000 0.1189
-8.250 -0.4278 0.09065 0.08565 -0.0313 1.0000 0.1227
-8.000 -0.4565 0.08895 0.08413 -0.0345 1.0000 0.1240
-7.750 -0.4778 0.08596 0.08119 -0.0392 1.0000 0.1248
-7.500 -0.4388 0.08174 0.07695 -0.0293 1.0000 0.1299
-7.250 -0.4431 0.07917 0.07444 -0.0291 1.0000 0.1344
-7.000 -0.4690 0.07672 0.07202 -0.0342 1.0000 0.1388
-6.750 -0.4725 0.07279 0.06813 -0.0329 1.0000 0.1407
-6.500 -0.4651 0.07029 0.06569 -0.0287 1.0000 0.1432
-6.250 -0.4649 0.06793 0.06334 -0.0267 1.0000 0.1471
-6.000 -0.4771 0.06440 0.05960 -0.0305 1.0000 0.1551
-5.750 -0.4702 0.06178 0.05709 -0.0266 1.0000 0.1575
-5.250 -0.4621 0.04463 0.03877 -0.0329 1.0000 0.1112
-5.000 -0.4494 0.04112 0.03511 -0.0317 1.0000 0.1058
-4.750 -0.4365 0.03671 0.03025 -0.0309 1.0000 0.1031
-4.500 -0.4211 0.03349 0.02652 -0.0299 1.0000 0.1055
-4.250 -0.4033 0.03053 0.02298 -0.0288 1.0000 0.1072
-4.000 -0.3835 0.02816 0.01992 -0.0276 1.0000 0.1094
-3.750 -0.3643 0.02648 0.01816 -0.0267 1.0000 0.1136
-3.500 -0.3447 0.02565 0.01714 -0.0256 1.0000 0.1208
-3.250 -0.3239 0.02417 0.01530 -0.0246 1.0000 0.1272
-3.000 -0.3036 0.02337 0.01438 -0.0235 1.0000 0.1354
-2.750 -0.2838 0.02270 0.01361 -0.0225 1.0000 0.1465
-2.500 -0.2637 0.02208 0.01286 -0.0215 1.0000 0.1587
-2.250 -0.2311 0.02167 0.01237 -0.0229 0.9961 0.1740
-2.000 -0.1943 0.02152 0.01219 -0.0252 0.9899 0.1921
-1.750 -0.1549 0.02141 0.01202 -0.0279 0.9838 0.2095
-1.500 -0.1205 0.02116 0.01177 -0.0296 0.9767 0.2237
-1.250 -0.0796 0.02100 0.01160 -0.0325 0.9706 0.2387
-1.000 -0.0468 0.02075 0.01141 -0.0338 0.9625 0.2537
-0.750 -0.0039 0.02058 0.01130 -0.0370 0.9561 0.2750
-0.500 0.0284 0.02022 0.01115 -0.0381 0.9467 0.3046
-0.250 0.1205 0.01771 0.01093 -0.0506 0.9474 1.0000
0.000 0.1686 0.01797 0.01090 -0.0546 0.9389 1.0000
0.250 0.2090 0.01812 0.01090 -0.0571 0.9285 1.0000
0.500 0.2462 0.01828 0.01094 -0.0590 0.9179 1.0000
0.750 0.2982 0.01833 0.01091 -0.0636 0.9113 1.0000
1.000 0.3290 0.01846 0.01100 -0.0642 0.8998 1.0000
1.250 0.3667 0.01854 0.01105 -0.0660 0.8902 1.0000
1.500 0.4172 0.01841 0.01092 -0.0701 0.8832 1.0000
1.750 0.4503 0.01846 0.01098 -0.0709 0.8720 1.0000
2.000 0.4952 0.01825 0.01081 -0.0736 0.8642 1.0000
2.250 0.5320 0.01810 0.01071 -0.0748 0.8537 1.0000
2.500 0.5626 0.01806 0.01071 -0.0747 0.8416 1.0000
2.750 0.5959 0.01790 0.01063 -0.0750 0.8298 1.0000
3.000 0.6308 0.01754 0.01034 -0.0752 0.8165 1.0000
3.250 0.6625 0.01709 0.00996 -0.0744 0.8000 1.0000
3.500 0.6927 0.01661 0.00953 -0.0733 0.7818 1.0000
3.750 0.7220 0.01621 0.00916 -0.0721 0.7637 1.0000
4.000 0.7468 0.01597 0.00897 -0.0703 0.7430 1.0000
4.250 0.7714 0.01574 0.00880 -0.0685 0.7206 1.0000
4.500 0.7954 0.01560 0.00870 -0.0668 0.6973 1.0000
4.750 0.8207 0.01547 0.00858 -0.0652 0.6731 1.0000
5.000 0.8431 0.01545 0.00860 -0.0633 0.6443 1.0000
5.250 0.8635 0.01548 0.00860 -0.0609 0.6080 1.0000
5.500 0.8826 0.01560 0.00862 -0.0584 0.5655 1.0000
5.750 0.9000 0.01585 0.00877 -0.0557 0.5181 1.0000
6.000 0.9159 0.01629 0.00899 -0.0529 0.4656 1.0000
6.250 0.9287 0.01699 0.00942 -0.0498 0.4029 1.0000
6.500 0.9381 0.01798 0.01003 -0.0463 0.3287 1.0000
6.750 0.9459 0.01921 0.01079 -0.0429 0.2506 1.0000
7.000 0.9536 0.02069 0.01177 -0.0397 0.1917 1.0000
7.250 0.9614 0.02248 0.01317 -0.0366 0.1503 1.0000
7.500 0.9722 0.02422 0.01470 -0.0339 0.1204 1.0000
7.750 0.9863 0.02593 0.01622 -0.0318 0.1005 1.0000
8.000 1.0040 0.02750 0.01776 -0.0301 0.0863 1.0000
8.250 1.0244 0.02918 0.01934 -0.0289 0.0766 1.0000
8.500 1.0457 0.03083 0.02119 -0.0276 0.0692 1.0000
8.750 1.0724 0.03341 0.02367 -0.0277 0.0641 1.0000
9.000 1.0945 0.03580 0.02643 -0.0264 0.0613 1.0000
9.250 1.1123 0.03796 0.02891 -0.0248 0.0580 1.0000
9.500 1.1291 0.04012 0.03123 -0.0234 0.0550 1.0000
9.750 1.1451 0.04326 0.03460 -0.0220 0.0537 1.0000
10.000 1.1564 0.04699 0.03866 -0.0202 0.0532 1.0000
10.250 1.1621 0.05078 0.04284 -0.0177 0.0530 1.0000
10.500 1.1633 0.05456 0.04700 -0.0150 0.0529 1.0000
10.750 1.1600 0.05859 0.05140 -0.0122 0.0527 1.0000
11.000 1.1517 0.06282 0.05594 -0.0094 0.0525 1.0000
11.250 1.1407 0.06620 0.05961 -0.0062 0.0526 1.0000
11.500 1.1246 0.06888 0.06254 -0.0025 0.0528 1.0000
11.750 1.1031 0.07158 0.06550 0.0008 0.0532 1.0000
12.000 1.0149 0.08011 0.07469 0.0013 0.0590 1.0000
12.250 0.9867 0.08671 0.08143 -0.0011 0.0605 1.0000
12.500 0.9632 0.09373 0.08853 -0.0044 0.0617 1.0000
12.750 0.9451 0.10101 0.09586 -0.0079 0.0627 1.0000
13.000 0.7137 0.12935 0.12454 -0.0254 0.0858 1.0000
13.250 0.7342 0.13152 0.12678 -0.0229 0.0888 1.0000
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Polar data table (+)
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