GOE 417 AIRFOIL (goe417-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: GOE 417 AIRFOIL (goe417-il) Reynolds number: 100,000 Max Cl/Cd: 63.19 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe417-il-100000-n5.txt Download as CSV file: xf-goe417-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 417 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.2893 0.11356 0.10897 -0.0256 1.0000 0.0263
-8.000 -0.2981 0.11274 0.10824 -0.0237 1.0000 0.0263
-7.750 -0.3074 0.11183 0.10743 -0.0217 1.0000 0.0263
-7.500 -0.3122 0.11045 0.10612 -0.0210 1.0000 0.0264
-7.250 -0.3147 0.10878 0.10451 -0.0206 1.0000 0.0264
-7.000 -0.2978 0.10545 0.10118 -0.0255 0.9964 0.0264
-6.750 -0.2831 0.09944 0.09520 -0.0260 0.9932 0.0267
-6.500 -0.2692 0.09422 0.08999 -0.0255 0.9901 0.0273
-6.250 -0.2490 0.09006 0.08580 -0.0289 0.9858 0.0280
-6.000 -0.2289 0.08641 0.08214 -0.0328 0.9799 0.0291
-5.750 -0.2050 0.08285 0.07856 -0.0380 0.9744 0.0310
-5.500 -0.1771 0.07939 0.07506 -0.0447 0.9678 0.0333
-5.250 -0.1344 0.07665 0.07222 -0.0567 0.9602 0.0349
-5.000 -0.0895 0.07327 0.06870 -0.0673 0.9543 0.0354
-4.750 -0.0644 0.06839 0.06380 -0.0713 0.9490 0.0358
-4.500 -0.0552 0.06415 0.05961 -0.0699 0.9429 0.0370
-4.250 -0.0286 0.06065 0.05606 -0.0729 0.9388 0.0398
-4.000 0.0286 0.05846 0.05358 -0.0843 0.9314 0.0470
-3.750 0.0591 0.05369 0.04875 -0.0886 0.9263 0.0484
-3.500 0.0847 0.05007 0.04514 -0.0905 0.9231 0.0515
-3.250 0.1139 0.04761 0.04254 -0.0932 0.9142 0.0574
-3.000 0.1559 0.04417 0.03896 -0.0987 0.9101 0.0658
-2.750 0.2028 0.04168 0.03610 -0.1044 0.9020 0.0765
-2.500 0.2347 0.03840 0.03282 -0.1067 0.8958 0.0819
-2.250 0.2707 0.03593 0.03014 -0.1093 0.8852 0.0951
-1.750 0.3609 0.02917 0.02241 -0.1153 0.8684 0.0591
-1.500 0.3938 0.02712 0.02004 -0.1162 0.8593 0.0593
-1.250 0.4295 0.02500 0.01762 -0.1177 0.8525 0.0575
-1.000 0.4595 0.02335 0.01565 -0.1178 0.8419 0.0551
-0.750 0.4917 0.02184 0.01379 -0.1182 0.8322 0.0535
-0.500 0.5259 0.02050 0.01210 -0.1188 0.8234 0.0527
-0.250 0.5544 0.01971 0.01109 -0.1185 0.8108 0.0557
0.000 0.5836 0.01890 0.01005 -0.1183 0.7979 0.0566
0.250 0.6128 0.01813 0.00909 -0.1180 0.7841 0.0561
0.500 0.6416 0.01748 0.00830 -0.1177 0.7694 0.0559
0.750 0.6703 0.01692 0.00762 -0.1173 0.7533 0.0560
1.000 0.6984 0.01645 0.00706 -0.1169 0.7361 0.0564
1.250 0.7254 0.01609 0.00660 -0.1163 0.7170 0.0571
1.500 0.7535 0.01583 0.00620 -0.1158 0.6960 0.0582
1.750 0.7818 0.01565 0.00585 -0.1154 0.6746 0.0599
2.000 0.8089 0.01555 0.00562 -0.1150 0.6521 0.0660
2.250 0.8359 0.01556 0.00548 -0.1144 0.6301 0.0739
2.500 0.8618 0.01561 0.00543 -0.1137 0.6072 0.0832
2.750 0.8905 0.01429 0.00559 -0.1140 0.5846 1.0000
3.000 0.9150 0.01458 0.00567 -0.1131 0.5617 1.0000
3.250 0.9392 0.01490 0.00578 -0.1123 0.5397 1.0000
3.500 0.9630 0.01524 0.00594 -0.1114 0.5172 1.0000
3.750 0.9861 0.01561 0.00612 -0.1104 0.4936 1.0000
4.000 1.0085 0.01601 0.00635 -0.1093 0.4688 1.0000
4.250 1.0308 0.01642 0.00660 -0.1082 0.4450 1.0000
4.500 1.0532 0.01686 0.00688 -0.1072 0.4250 1.0000
4.750 1.0759 0.01729 0.00721 -0.1063 0.4088 1.0000
5.000 1.0989 0.01771 0.00760 -0.1054 0.3948 1.0000
5.250 1.1220 0.01814 0.00800 -0.1046 0.3822 1.0000
5.500 1.1450 0.01858 0.00843 -0.1038 0.3704 1.0000
5.750 1.1678 0.01904 0.00890 -0.1030 0.3594 1.0000
6.000 1.1903 0.01952 0.00938 -0.1021 0.3493 1.0000
6.250 1.2131 0.01997 0.00991 -0.1013 0.3389 1.0000
6.500 1.2355 0.02046 0.01045 -0.1005 0.3293 1.0000
6.750 1.2573 0.02099 0.01104 -0.0995 0.3197 1.0000
7.000 1.2794 0.02148 0.01166 -0.0986 0.3095 1.0000
7.250 1.3009 0.02203 0.01229 -0.0977 0.3000 1.0000
7.500 1.3220 0.02259 0.01295 -0.0966 0.2903 1.0000
7.750 1.3428 0.02314 0.01369 -0.0956 0.2795 1.0000
8.000 1.3610 0.02367 0.01432 -0.0941 0.2642 1.0000
8.250 1.3770 0.02422 0.01490 -0.0924 0.2455 1.0000
8.500 1.3936 0.02480 0.01557 -0.0908 0.2265 1.0000
8.750 1.4084 0.02548 0.01628 -0.0889 0.2050 1.0000
9.000 1.4200 0.02640 0.01710 -0.0868 0.1774 1.0000
9.250 1.4302 0.02753 0.01819 -0.0845 0.1505 1.0000
9.500 1.4349 0.02907 0.01953 -0.0816 0.1079 1.0000
10.000 1.4240 0.03394 0.02388 -0.0738 0.0327 1.0000
10.250 1.4238 0.03598 0.02602 -0.0708 0.0278 1.0000
10.500 1.4248 0.03794 0.02815 -0.0681 0.0252 1.0000
10.750 1.4244 0.04006 0.03047 -0.0656 0.0230 1.0000
11.000 1.4208 0.04255 0.03314 -0.0633 0.0213 1.0000
11.250 1.4146 0.04541 0.03620 -0.0613 0.0204 1.0000
11.500 1.4057 0.04870 0.03971 -0.0597 0.0198 1.0000
11.750 1.3943 0.05252 0.04375 -0.0586 0.0193 1.0000
12.000 1.3852 0.05638 0.04780 -0.0581 0.0190 1.0000
12.250 1.3757 0.06059 0.05228 -0.0582 0.0188 1.0000
12.500 1.3653 0.06521 0.05711 -0.0588 0.0186 1.0000
12.750 1.3550 0.07011 0.06221 -0.0598 0.0184 1.0000
13.000 1.3449 0.07521 0.06749 -0.0611 0.0181 1.0000
13.250 1.3357 0.08032 0.07279 -0.0625 0.0178 1.0000
13.500 1.3278 0.08535 0.07799 -0.0639 0.0174 1.0000
13.750 1.3213 0.09019 0.08298 -0.0652 0.0170 1.0000
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Polar data table (+)
Polar graphs
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