EPPLER STE 871-514 AIRFOIL (ste87151-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER STE 871-514 AIRFOIL (ste87151-il) Reynolds number: 1,000,000 Max Cl/Cd: 113.92 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ste87151-il-1000000.txt Download as CSV file: xf-ste87151-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER STE 871-514 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.2879 0.08333 0.08124 -0.1043 0.9116 0.0102
-10.750 -0.2944 0.07636 0.07428 -0.1093 0.9105 0.0104
-10.500 -0.3099 0.06621 0.06411 -0.1192 0.9094 0.0102
-10.250 -0.3308 0.05911 0.05690 -0.1242 0.9082 0.0101
-10.000 -0.3490 0.05387 0.05154 -0.1261 0.9069 0.0102
-9.750 -0.3809 0.03018 0.02733 -0.1238 0.9043 0.0114
-9.500 -0.3776 0.02832 0.02539 -0.1230 0.9037 0.0115
-9.250 -0.3697 0.02670 0.02370 -0.1224 0.9031 0.0117
-9.000 -0.3584 0.02544 0.02238 -0.1219 0.9024 0.0119
-8.750 -0.3489 0.02341 0.02021 -0.1211 0.9014 0.0123
-8.250 -0.3645 0.02295 0.01834 -0.1177 0.8995 0.0116
-8.000 -0.3424 0.02130 0.01659 -0.1172 0.8989 0.0106
-7.750 -0.3216 0.01933 0.01437 -0.1165 0.8983 0.0103
-7.500 -0.2986 0.01779 0.01263 -0.1160 0.8977 0.0101
-7.250 -0.2745 0.01686 0.01157 -0.1156 0.8972 0.0103
-7.000 -0.2501 0.01603 0.01064 -0.1153 0.8967 0.0105
-6.750 -0.2263 0.01518 0.00970 -0.1148 0.8961 0.0104
-6.500 -0.2029 0.01445 0.00891 -0.1143 0.8954 0.0104
-6.250 -0.1797 0.01386 0.00827 -0.1138 0.8947 0.0106
-6.000 -0.1564 0.01340 0.00777 -0.1133 0.8941 0.0108
-5.750 -0.1321 0.01306 0.00739 -0.1129 0.8934 0.0111
-5.500 -0.1099 0.01245 0.00675 -0.1123 0.8925 0.0119
-5.250 -0.0841 0.01221 0.00650 -0.1122 0.8915 0.0128
-5.000 -0.0591 0.01211 0.00637 -0.1120 0.8902 0.0137
-4.750 -0.0363 0.01187 0.00613 -0.1114 0.8893 0.0147
-4.500 -0.0126 0.01171 0.00598 -0.1109 0.8883 0.0159
-4.250 0.0104 0.01157 0.00584 -0.1103 0.8871 0.0174
-4.000 0.0336 0.01149 0.00577 -0.1097 0.8858 0.0200
-3.750 0.0568 0.01134 0.00565 -0.1091 0.8843 0.0252
-3.500 0.0805 0.01116 0.00554 -0.1086 0.8830 0.0389
-3.250 0.1022 0.01062 0.00540 -0.1081 0.8817 0.1387
-3.000 0.1245 0.00980 0.00520 -0.1081 0.8805 0.3166
-2.750 0.1499 0.00845 0.00483 -0.1091 0.8794 0.6001
-2.500 0.1784 0.00830 0.00505 -0.1094 0.8785 0.7259
-2.250 0.2080 0.00846 0.00520 -0.1097 0.8777 0.7442
-2.000 0.2369 0.00867 0.00542 -0.1097 0.8769 0.7549
-1.750 0.2658 0.00903 0.00576 -0.1097 0.8762 0.7670
-1.500 0.2925 0.00930 0.00609 -0.1092 0.8754 0.7740
-1.250 0.3208 0.00956 0.00635 -0.1092 0.8744 0.7787
-1.000 0.3441 0.00977 0.00655 -0.1083 0.8721 0.7829
-0.750 0.3683 0.00978 0.00656 -0.1077 0.8690 0.7862
-0.500 0.3965 0.00967 0.00647 -0.1077 0.8668 0.7886
-0.250 0.4282 0.00948 0.00627 -0.1084 0.8648 0.7902
0.000 0.4626 0.00918 0.00595 -0.1097 0.8629 0.7914
0.250 0.4976 0.00886 0.00561 -0.1112 0.8609 0.7924
0.500 0.5339 0.00859 0.00530 -0.1129 0.8587 0.7933
0.750 0.5606 0.00845 0.00518 -0.1128 0.8551 0.7944
1.000 0.5878 0.00823 0.00496 -0.1126 0.8504 0.7956
1.250 0.6208 0.00789 0.00460 -0.1137 0.8461 0.7966
1.500 0.6535 0.00753 0.00421 -0.1147 0.8401 0.7974
1.750 0.6779 0.00733 0.00403 -0.1140 0.8322 0.7985
2.000 0.7082 0.00713 0.00380 -0.1146 0.8265 0.7992
2.250 0.7302 0.00707 0.00379 -0.1135 0.8175 0.7999
2.500 0.7551 0.00697 0.00369 -0.1130 0.8052 0.8007
2.750 0.7781 0.00683 0.00347 -0.1119 0.7726 0.8014
3.000 0.7858 0.00715 0.00342 -0.1076 0.7039 0.8019
3.250 0.7800 0.00774 0.00368 -0.1006 0.6413 0.8025
3.500 0.7889 0.00818 0.00392 -0.0971 0.5965 0.8036
3.750 0.7938 0.00893 0.00435 -0.0929 0.5254 0.8047
4.000 0.8020 0.00966 0.00478 -0.0895 0.4607 0.8054
4.250 0.8127 0.01036 0.00518 -0.0866 0.3965 0.8060
4.500 0.8260 0.01102 0.00557 -0.0843 0.3370 0.8065
4.750 0.8407 0.01165 0.00594 -0.0823 0.2818 0.8071
5.000 0.8566 0.01225 0.00629 -0.0805 0.2327 0.8076
5.250 0.8742 0.01278 0.00663 -0.0791 0.1923 0.8081
5.500 0.8925 0.01330 0.00698 -0.0778 0.1584 0.8085
5.750 0.9110 0.01380 0.00733 -0.0766 0.1304 0.8091
6.000 0.9303 0.01427 0.00768 -0.0754 0.1087 0.8097
6.250 0.9504 0.01469 0.00801 -0.0744 0.0931 0.8104
6.500 0.9711 0.01508 0.00836 -0.0736 0.0824 0.8109
6.750 0.9920 0.01546 0.00870 -0.0727 0.0749 0.8113
7.000 1.0127 0.01585 0.00906 -0.0719 0.0693 0.8118
7.250 1.0334 0.01623 0.00942 -0.0710 0.0649 0.8123
7.500 1.0545 0.01659 0.00979 -0.0703 0.0616 0.8127
7.750 1.0745 0.01703 0.01021 -0.0693 0.0585 0.8132
8.000 1.0958 0.01737 0.01059 -0.0686 0.0562 0.8137
8.250 1.1157 0.01780 0.01101 -0.0677 0.0540 0.8142
8.500 1.1354 0.01823 0.01147 -0.0667 0.0521 0.8146
8.750 1.1554 0.01864 0.01190 -0.0658 0.0503 0.8151
9.000 1.1736 0.01917 0.01243 -0.0647 0.0487 0.8156
9.250 1.1926 0.01964 0.01294 -0.0636 0.0474 0.8161
9.500 1.2121 0.02008 0.01342 -0.0627 0.0461 0.8165
9.750 1.2304 0.02058 0.01394 -0.0616 0.0449 0.8170
10.000 1.2455 0.02129 0.01466 -0.0601 0.0436 0.8174
10.250 1.2642 0.02177 0.01520 -0.0590 0.0427 0.8179
10.500 1.2821 0.02229 0.01577 -0.0579 0.0417 0.8183
10.750 1.2998 0.02282 0.01632 -0.0569 0.0406 0.8188
11.000 1.3139 0.02358 0.01710 -0.0553 0.0393 0.8193
11.250 1.3332 0.02402 0.01759 -0.0545 0.0381 0.8197
11.500 1.3503 0.02459 0.01821 -0.0534 0.0371 0.8202
11.750 1.3661 0.02526 0.01890 -0.0521 0.0362 0.8208
12.000 1.3789 0.02611 0.01979 -0.0505 0.0353 0.8217
12.250 1.3962 0.02665 0.02041 -0.0495 0.0344 0.8225
12.500 1.4120 0.02728 0.02109 -0.0483 0.0335 0.8232
12.750 1.4269 0.02799 0.02185 -0.0471 0.0326 0.8240
13.000 1.4376 0.02899 0.02288 -0.0453 0.0318 0.8247
13.250 1.4533 0.02965 0.02363 -0.0442 0.0310 0.8255
13.500 1.4686 0.03033 0.02438 -0.0432 0.0300 0.8262
13.750 1.4831 0.03108 0.02516 -0.0420 0.0289 0.8270
14.000 1.4950 0.03204 0.02616 -0.0406 0.0279 0.8277
14.250 1.5108 0.03272 0.02693 -0.0397 0.0268 0.8285
14.500 1.5243 0.03357 0.02782 -0.0386 0.0257 0.8293
14.750 1.5356 0.03460 0.02890 -0.0372 0.0246 0.8302
15.000 1.5490 0.03547 0.02984 -0.0362 0.0235 0.8310
15.250 1.5600 0.03653 0.03095 -0.0349 0.0223 0.8319
15.500 1.5703 0.03768 0.03217 -0.0337 0.0213 0.8327
15.750 1.5806 0.03886 0.03342 -0.0325 0.0200 0.8336
16.000 1.5887 0.04025 0.03485 -0.0312 0.0188 0.8345
16.250 1.5975 0.04162 0.03630 -0.0300 0.0175 0.8353
16.500 1.6032 0.04327 0.03799 -0.0287 0.0162 0.8361
16.750 1.6089 0.04496 0.03977 -0.0275 0.0150 0.8369
17.000 1.6116 0.04699 0.04184 -0.0262 0.0138 0.8376
17.250 1.6148 0.04903 0.04398 -0.0252 0.0128 0.8390
17.500 1.6150 0.05142 0.04645 -0.0241 0.0119 0.8404
17.750 1.6137 0.05403 0.04916 -0.0231 0.0112 0.8418
18.000 1.6128 0.05670 0.05194 -0.0223 0.0105 0.8432
18.250 1.6091 0.05977 0.05512 -0.0217 0.0100 0.8445
18.500 1.6022 0.06335 0.05881 -0.0213 0.0094 0.8458
18.750 1.5971 0.06683 0.06241 -0.0212 0.0091 0.8472
19.000 1.5909 0.07054 0.06624 -0.0214 0.0087 0.8486
19.250 1.5821 0.07469 0.07052 -0.0218 0.0084 0.8499
|
Polar data table (+)
Polar graphs
<< Back to EPPLER STE 871-514 AIRFOIL (ste87151-il)