DU 86-137/25 AIRFOIL (du861372-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: DU 86-137/25 AIRFOIL (du861372-il) Reynolds number: 100,000 Max Cl/Cd: 27.08 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-du861372-il-100000-n5.txt Download as CSV file: xf-du861372-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DU 86-137/25 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
------ -------- --------- --------- -------- -------- --------
-13.000 -0.7491 0.08943 0.08318 -0.0385 0.3024 0.0468
-12.750 -0.7668 0.08395 0.07762 -0.0408 0.3026 0.0465
-12.500 -0.7855 0.07896 0.07253 -0.0424 0.3027 0.0462
-12.250 -0.8079 0.07426 0.06768 -0.0431 0.3029 0.0460
-12.000 -0.8295 0.07012 0.06336 -0.0428 0.3029 0.0459
-11.750 -0.8481 0.06654 0.05957 -0.0417 0.3029 0.0458
-11.250 -0.8844 0.06032 0.05288 -0.0369 0.3028 0.0460
-11.000 -0.8979 0.05783 0.05012 -0.0336 0.3026 0.0461
-10.750 -0.9091 0.05510 0.04704 -0.0303 0.3026 0.0464
-10.500 -0.9153 0.05238 0.04387 -0.0272 0.3020 0.0467
-10.250 -0.9074 0.04965 0.04089 -0.0257 0.3013 0.0475
-10.000 -0.8926 0.04755 0.03867 -0.0249 0.3008 0.0488
-9.750 -0.8835 0.04546 0.03626 -0.0232 0.3005 0.0503
-9.500 -0.8744 0.04325 0.03362 -0.0212 0.3001 0.0514
-9.250 -0.8588 0.04097 0.03096 -0.0198 0.2998 0.0518
-9.000 -0.8370 0.03868 0.02838 -0.0192 0.2995 0.0517
-8.750 -0.8107 0.03651 0.02597 -0.0191 0.2992 0.0516
-8.500 -0.7802 0.03444 0.02372 -0.0194 0.2988 0.0511
-8.000 -0.7133 0.03095 0.01996 -0.0209 0.2983 0.0501
-7.750 -0.6824 0.02954 0.01848 -0.0213 0.2982 0.0491
-7.500 -0.6572 0.02843 0.01722 -0.0208 0.2981 0.0476
-7.250 -0.6383 0.02762 0.01629 -0.0195 0.2979 0.0466
-7.000 -0.6217 0.02685 0.01541 -0.0181 0.2979 0.0459
-6.750 -0.6048 0.02618 0.01461 -0.0166 0.2978 0.0451
-6.500 -0.5869 0.02560 0.01389 -0.0153 0.2978 0.0444
-6.250 -0.5679 0.02511 0.01325 -0.0141 0.2978 0.0436
-6.000 -0.5479 0.02469 0.01269 -0.0131 0.2978 0.0429
-5.750 -0.5273 0.02431 0.01218 -0.0122 0.2978 0.0423
-5.500 -0.5072 0.02388 0.01162 -0.0112 0.2979 0.0419
-5.250 -0.4865 0.02349 0.01112 -0.0102 0.2979 0.0414
-5.000 -0.4654 0.02313 0.01066 -0.0094 0.2980 0.0410
-4.750 -0.4442 0.02276 0.01020 -0.0085 0.2981 0.0409
-4.500 -0.4231 0.02237 0.00969 -0.0077 0.2982 0.0413
-4.250 -0.4015 0.02204 0.00927 -0.0069 0.2982 0.0415
-4.000 -0.3791 0.02176 0.00892 -0.0062 0.2984 0.0413
-3.750 -0.3562 0.02154 0.00862 -0.0055 0.2985 0.0409
-3.500 -0.3330 0.02134 0.00832 -0.0050 0.2987 0.0405
-3.250 -0.3097 0.02115 0.00807 -0.0044 0.2988 0.0401
-3.000 -0.2861 0.02097 0.00784 -0.0039 0.2990 0.0397
-2.750 -0.2623 0.02081 0.00763 -0.0034 0.2992 0.0393
-2.500 -0.2385 0.02065 0.00742 -0.0030 0.2994 0.0390
-2.250 -0.2145 0.02050 0.00721 -0.0025 0.2997 0.0387
-2.000 -0.1905 0.02034 0.00702 -0.0021 0.3000 0.0384
-1.750 -0.1662 0.02021 0.00685 -0.0017 0.3002 0.0382
-1.500 -0.1418 0.02008 0.00670 -0.0013 0.3005 0.0380
-1.000 -0.0927 0.01985 0.00640 -0.0006 0.3012 0.0377
-0.750 -0.0680 0.01974 0.00629 -0.0003 0.3016 0.0377
-0.500 -0.0432 0.01963 0.00618 0.0000 0.3020 0.0377
-0.250 -0.0183 0.01953 0.00609 0.0003 0.3025 0.0380
0.000 0.0062 0.01938 0.00604 0.0007 0.3029 0.0411
0.500 0.0555 0.01919 0.00593 0.0014 0.3034 0.0809
0.750 0.0805 0.01914 0.00592 0.0016 0.3036 0.0832
1.250 0.1195 0.01780 0.00563 0.0033 0.3040 0.3602
1.500 0.1286 0.01683 0.00613 0.0071 0.3044 0.6497
1.750 0.1525 0.01693 0.00642 0.0080 0.3049 0.6892
2.000 0.1755 0.01711 0.00672 0.0091 0.3050 0.7268
2.250 0.1991 0.01729 0.00688 0.0098 0.3042 0.7421
2.500 0.2213 0.01757 0.00715 0.0109 0.3032 0.7606
2.750 0.2427 0.01788 0.00749 0.0123 0.3026 0.7796
3.000 0.2633 0.01820 0.00786 0.0140 0.3020 0.7970
3.250 0.2859 0.01840 0.00805 0.0149 0.3014 0.8076
3.500 0.3076 0.01873 0.00836 0.0161 0.3001 0.8137
3.750 0.3246 0.01931 0.00888 0.0190 0.2986 0.8339
4.000 0.3451 0.01968 0.00925 0.0209 0.2973 0.8480
4.250 0.3617 0.02005 0.00970 0.0245 0.2966 0.8718
4.500 0.3849 0.02038 0.01004 0.0259 0.2955 0.8849
4.750 0.4187 0.02119 0.01060 0.0246 0.2917 0.8874
5.000 0.4534 0.02245 0.01169 0.0229 0.2831 0.8900
5.250 0.4826 0.02360 0.01285 0.0221 0.2724 0.8931
5.500 0.5073 0.02440 0.01380 0.0221 0.2627 0.8966
5.750 0.5324 0.02506 0.01466 0.0221 0.2524 0.8994
6.000 0.5550 0.02546 0.01531 0.0226 0.2414 0.9030
6.250 0.5712 0.02524 0.01546 0.0243 0.2295 0.9078
6.750 0.5996 0.02314 0.01446 0.0288 0.1678 0.9180
7.000 0.6125 0.02288 0.01420 0.0307 0.1446 0.9241
7.250 0.6296 0.02332 0.01457 0.0318 0.1157 0.9296
7.500 0.6499 0.02400 0.01510 0.0323 0.0929 0.9337
7.750 0.6692 0.02472 0.01574 0.0329 0.0777 0.9388
8.000 0.6883 0.02544 0.01653 0.0336 0.0671 0.9447
8.250 0.7087 0.02638 0.01749 0.0338 0.0586 0.9502
8.500 0.7289 0.02727 0.01849 0.0341 0.0527 0.9569
8.750 0.7469 0.02859 0.01985 0.0340 0.0484 0.9635
9.000 0.7699 0.02967 0.02114 0.0337 0.0445 0.9707
9.250 0.7924 0.03081 0.02241 0.0330 0.0404 0.9805
9.500 0.8094 0.03210 0.02377 0.0331 0.0381 1.0000
9.750 0.8158 0.03360 0.02526 0.0351 0.0366 1.0000
10.000 0.8327 0.03504 0.02691 0.0360 0.0354 1.0000
10.250 0.8501 0.03672 0.02881 0.0366 0.0340 1.0000
10.500 0.8652 0.03842 0.03072 0.0373 0.0324 1.0000
10.750 0.8761 0.03989 0.03234 0.0382 0.0305 1.0000
11.000 0.8864 0.04164 0.03423 0.0391 0.0294 1.0000
11.250 0.8945 0.04353 0.03623 0.0399 0.0285 1.0000
11.500 0.9004 0.04598 0.03881 0.0407 0.0277 1.0000
11.750 0.9015 0.04911 0.04213 0.0416 0.0272 1.0000
12.000 0.8980 0.05200 0.04532 0.0425 0.0270 1.0000
12.250 0.8899 0.05527 0.04888 0.0432 0.0268 1.0000
12.500 0.8784 0.05891 0.05279 0.0435 0.0267 1.0000
12.750 0.8631 0.06301 0.05715 0.0432 0.0266 1.0000
13.000 0.8448 0.06772 0.06209 0.0422 0.0266 1.0000
13.250 0.8236 0.07324 0.06782 0.0401 0.0266 1.0000
13.500 0.8007 0.07969 0.07446 0.0367 0.0267 1.0000
13.750 0.7761 0.08758 0.08250 0.0317 0.0269 1.0000
14.000 0.7479 0.09789 0.09293 0.0246 0.0271 1.0000
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