GOE 619 AIRFOIL (goe619-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 619 AIRFOIL (goe619-il) Reynolds number: 500,000 Max Cl/Cd: 82.04 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe619-il-500000-n5.txt Download as CSV file: xf-goe619-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 619 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
------ -------- --------- --------- -------- -------- --------
-15.000 -0.6817 0.08353 0.08035 -0.0671 1.0000 0.0254
-14.750 -0.7166 0.07470 0.07138 -0.0722 1.0000 0.0254
-14.500 -0.7473 0.06734 0.06391 -0.0764 1.0000 0.0255
-14.250 -0.7727 0.06122 0.05769 -0.0796 1.0000 0.0255
-14.000 -0.7986 0.05528 0.05166 -0.0827 1.0000 0.0255
-13.750 -0.8247 0.04960 0.04588 -0.0855 1.0000 0.0257
-13.500 -0.8518 0.04427 0.04048 -0.0877 1.0000 0.0257
-13.250 -0.8591 0.03727 0.03329 -0.0970 0.9973 0.0258
-13.000 -0.8497 0.03251 0.02826 -0.1052 0.9895 0.0260
-12.750 -0.8292 0.03028 0.02584 -0.1085 0.9842 0.0263
-12.500 -0.8101 0.02861 0.02402 -0.1099 0.9780 0.0264
-12.250 -0.7901 0.02675 0.02203 -0.1115 0.9726 0.0268
-12.000 -0.7721 0.02522 0.02042 -0.1120 0.9656 0.0272
-11.750 -0.7485 0.02398 0.01911 -0.1129 0.9602 0.0276
-11.500 -0.7239 0.02291 0.01796 -0.1137 0.9547 0.0279
-11.250 -0.6978 0.02192 0.01688 -0.1145 0.9484 0.0282
-11.000 -0.6660 0.02093 0.01579 -0.1164 0.9444 0.0286
-10.750 -0.6380 0.02007 0.01483 -0.1174 0.9377 0.0289
-10.500 -0.6067 0.01926 0.01393 -0.1188 0.9318 0.0294
-10.250 -0.5748 0.01849 0.01306 -0.1203 0.9254 0.0299
-10.000 -0.5446 0.01780 0.01227 -0.1214 0.9170 0.0303
-9.750 -0.5143 0.01714 0.01149 -0.1223 0.9087 0.0306
-9.500 -0.4865 0.01656 0.01080 -0.1227 0.8993 0.0309
-9.250 -0.4595 0.01606 0.01019 -0.1228 0.8901 0.0312
-9.000 -0.4330 0.01559 0.00962 -0.1227 0.8802 0.0314
-8.750 -0.4096 0.01501 0.00895 -0.1221 0.8699 0.0319
-8.500 -0.3851 0.01449 0.00834 -0.1217 0.8605 0.0324
-8.250 -0.3611 0.01406 0.00784 -0.1211 0.8510 0.0328
-8.000 -0.3362 0.01366 0.00736 -0.1206 0.8425 0.0335
-7.750 -0.3114 0.01332 0.00694 -0.1200 0.8332 0.0341
-7.500 -0.2862 0.01301 0.00655 -0.1195 0.8245 0.0348
-7.250 -0.2610 0.01271 0.00618 -0.1189 0.8154 0.0354
-7.000 -0.2354 0.01245 0.00583 -0.1184 0.8076 0.0359
-6.750 -0.2098 0.01219 0.00550 -0.1178 0.7991 0.0366
-6.500 -0.1840 0.01197 0.00519 -0.1173 0.7911 0.0372
-6.250 -0.1585 0.01169 0.00486 -0.1168 0.7826 0.0385
-6.000 -0.1326 0.01146 0.00457 -0.1163 0.7750 0.0400
-5.750 -0.1067 0.01123 0.00431 -0.1158 0.7669 0.0420
-5.500 -0.0808 0.01102 0.00405 -0.1152 0.7594 0.0455
-5.250 -0.0551 0.01074 0.00382 -0.1147 0.7514 0.0559
-5.000 -0.0291 0.01057 0.00367 -0.1143 0.7441 0.0700
-4.750 -0.0023 0.01046 0.00354 -0.1139 0.7366 0.0777
-4.500 0.0241 0.01034 0.00341 -0.1135 0.7290 0.0842
-4.250 0.0510 0.01027 0.00329 -0.1131 0.7219 0.0882
-4.000 0.0776 0.01017 0.00316 -0.1126 0.7141 0.0922
-3.750 0.1042 0.01008 0.00304 -0.1122 0.7068 0.0961
-3.500 0.1309 0.01000 0.00293 -0.1118 0.6989 0.0995
-3.250 0.1576 0.00996 0.00283 -0.1113 0.6912 0.1025
-3.000 0.1838 0.00986 0.00272 -0.1108 0.6817 0.1077
-2.750 0.2097 0.00980 0.00261 -0.1103 0.6704 0.1124
-2.500 0.2351 0.00977 0.00252 -0.1096 0.6569 0.1170
-2.250 0.2604 0.00972 0.00243 -0.1089 0.6437 0.1244
-2.000 0.2864 0.00963 0.00235 -0.1083 0.6333 0.1358
-1.750 0.3121 0.00954 0.00230 -0.1078 0.6251 0.1550
-1.500 0.3385 0.00945 0.00227 -0.1073 0.6166 0.1795
-1.000 0.3911 0.00937 0.00225 -0.1064 0.6018 0.2171
-0.750 0.4173 0.00937 0.00225 -0.1059 0.5939 0.2301
-0.500 0.4439 0.00936 0.00225 -0.1055 0.5860 0.2421
-0.250 0.4696 0.00938 0.00226 -0.1049 0.5765 0.2525
0.000 0.4958 0.00940 0.00227 -0.1044 0.5659 0.2621
0.250 0.5213 0.00943 0.00229 -0.1038 0.5548 0.2715
0.500 0.5466 0.00948 0.00231 -0.1031 0.5429 0.2820
0.750 0.5720 0.00952 0.00235 -0.1025 0.5301 0.2935
1.000 0.5970 0.00958 0.00239 -0.1017 0.5160 0.3028
1.250 0.6212 0.00966 0.00244 -0.1009 0.4999 0.3166
1.750 0.6677 0.00990 0.00261 -0.0988 0.4619 0.3477
2.000 0.6902 0.01005 0.00272 -0.0977 0.4406 0.3648
2.250 0.7121 0.01023 0.00286 -0.0965 0.4180 0.3878
2.500 0.7334 0.01042 0.00302 -0.0952 0.3960 0.4142
2.750 0.7547 0.01064 0.00318 -0.0939 0.3733 0.4360
3.000 0.7758 0.01085 0.00336 -0.0925 0.3524 0.4601
3.500 0.8172 0.01119 0.00371 -0.0898 0.3173 0.5266
3.750 0.8263 0.01057 0.00397 -0.0858 0.3073 0.8407
4.250 0.9136 0.01117 0.00452 -0.0929 0.2768 1.0000
4.500 0.9347 0.01140 0.00470 -0.0915 0.2678 1.0000
4.750 0.9553 0.01165 0.00489 -0.0901 0.2592 1.0000
5.000 0.9758 0.01190 0.00509 -0.0887 0.2510 1.0000
5.250 0.9959 0.01215 0.00530 -0.0871 0.2440 1.0000
5.500 1.0156 0.01238 0.00550 -0.0855 0.2380 1.0000
5.750 1.0342 0.01263 0.00571 -0.0837 0.2330 1.0000
6.000 1.0533 0.01287 0.00594 -0.0820 0.2287 1.0000
6.250 1.0729 0.01309 0.00616 -0.0805 0.2246 1.0000
6.500 1.0920 0.01335 0.00641 -0.0788 0.2203 1.0000
6.750 1.1099 0.01366 0.00669 -0.0770 0.2154 1.0000
7.000 1.1293 0.01392 0.00695 -0.0755 0.2116 1.0000
7.250 1.1482 0.01421 0.00724 -0.0740 0.2057 1.0000
7.500 1.1653 0.01458 0.00757 -0.0722 0.1998 1.0000
7.750 1.1841 0.01488 0.00788 -0.0707 0.1937 1.0000
8.000 1.2014 0.01526 0.00823 -0.0690 0.1865 1.0000
8.250 1.2188 0.01565 0.00860 -0.0673 0.1792 1.0000
8.500 1.2350 0.01610 0.00901 -0.0656 0.1704 1.0000
8.750 1.2506 0.01659 0.00946 -0.0638 0.1587 1.0000
9.000 1.2612 0.01735 0.01009 -0.0614 0.1373 1.0000
9.250 1.2591 0.01884 0.01128 -0.0575 0.0946 1.0000
9.500 1.2577 0.02040 0.01265 -0.0540 0.0649 1.0000
9.750 1.2695 0.02124 0.01347 -0.0521 0.0596 1.0000
10.000 1.2829 0.02199 0.01424 -0.0505 0.0569 1.0000
10.250 1.2959 0.02280 0.01507 -0.0489 0.0547 1.0000
10.500 1.3082 0.02368 0.01596 -0.0474 0.0522 1.0000
10.750 1.3216 0.02450 0.01683 -0.0460 0.0507 1.0000
11.000 1.3351 0.02534 0.01771 -0.0447 0.0490 1.0000
11.250 1.3475 0.02628 0.01868 -0.0434 0.0473 1.0000
11.500 1.3579 0.02738 0.01981 -0.0420 0.0450 1.0000
11.750 1.3683 0.02853 0.02098 -0.0407 0.0432 1.0000
12.000 1.3805 0.02956 0.02205 -0.0396 0.0402 1.0000
12.250 1.3894 0.03087 0.02336 -0.0384 0.0367 1.0000
12.500 1.3876 0.03309 0.02547 -0.0366 0.0173 1.0000
12.750 1.3882 0.03522 0.02764 -0.0352 0.0144 1.0000
13.000 1.3932 0.03703 0.02953 -0.0341 0.0135 1.0000
13.250 1.3970 0.03902 0.03161 -0.0331 0.0128 1.0000
13.500 1.4003 0.04112 0.03379 -0.0323 0.0122 1.0000
13.750 1.4052 0.04311 0.03587 -0.0316 0.0119 1.0000
14.000 1.4086 0.04532 0.03816 -0.0310 0.0115 1.0000
14.250 1.4113 0.04766 0.04059 -0.0305 0.0112 1.0000
14.500 1.4133 0.05014 0.04316 -0.0301 0.0109 1.0000
14.750 1.4132 0.05292 0.04603 -0.0298 0.0106 1.0000
15.000 1.4111 0.05601 0.04923 -0.0296 0.0103 1.0000
15.250 1.4088 0.05922 0.05254 -0.0297 0.0102 1.0000
15.500 1.4048 0.06272 0.05615 -0.0298 0.0099 1.0000
15.750 1.4035 0.06596 0.05949 -0.0301 0.0097 1.0000
16.000 1.3997 0.06962 0.06326 -0.0306 0.0096 1.0000
16.250 1.3958 0.07339 0.06713 -0.0312 0.0094 1.0000
16.500 1.3903 0.07746 0.07131 -0.0320 0.0092 1.0000
16.750 1.3831 0.08185 0.07582 -0.0330 0.0091 1.0000
17.000 1.3753 0.08644 0.08051 -0.0342 0.0090 1.0000
17.250 1.3669 0.09122 0.08541 -0.0355 0.0089 1.0000
17.500 1.3568 0.09634 0.09066 -0.0371 0.0088 1.0000
17.750 1.3463 0.10162 0.09605 -0.0388 0.0087 1.0000
18.000 1.3365 0.10686 0.10140 -0.0407 0.0086 1.0000
18.250 1.3259 0.11232 0.10697 -0.0427 0.0085 1.0000
18.500 1.3151 0.11787 0.11263 -0.0449 0.0084 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 619 AIRFOIL (goe619-il)