(Dicke 12.28%) (e171-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: (Dicke 12.28%) (e171-il) Reynolds number: 100,000 Max Cl/Cd: 41.51 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e171-il-100000-n5.txt Download as CSV file: xf-e171-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: (Dicke 12.28%)
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 -0.7521 0.09272 0.08764 -0.0167 1.0000 0.0191
-12.750 -0.7909 0.08106 0.07571 -0.0248 1.0000 0.0185
-12.500 -0.8154 0.07380 0.06820 -0.0294 1.0000 0.0181
-12.250 -0.8339 0.06825 0.06244 -0.0322 1.0000 0.0181
-12.000 -0.8504 0.06336 0.05736 -0.0340 1.0000 0.0182
-11.750 -0.8638 0.05924 0.05303 -0.0349 1.0000 0.0182
-11.500 -0.8757 0.05543 0.04903 -0.0352 1.0000 0.0184
-11.250 -0.8850 0.05209 0.04547 -0.0347 1.0000 0.0185
-11.000 -0.8916 0.04911 0.04228 -0.0336 1.0000 0.0187
-10.750 -0.8955 0.04641 0.03937 -0.0320 1.0000 0.0190
-10.500 -0.8964 0.04400 0.03675 -0.0301 1.0000 0.0194
-10.250 -0.8944 0.04174 0.03429 -0.0280 1.0000 0.0199
-10.000 -0.8887 0.03954 0.03186 -0.0261 1.0000 0.0206
-9.750 -0.8789 0.03737 0.02945 -0.0245 1.0000 0.0215
-9.500 -0.8663 0.03533 0.02718 -0.0230 1.0000 0.0226
-9.250 -0.8523 0.03378 0.02534 -0.0215 1.0000 0.0248
-9.000 -0.8394 0.03173 0.02323 -0.0201 1.0000 0.0268
-8.750 -0.8257 0.03025 0.02166 -0.0186 1.0000 0.0290
-8.500 -0.8107 0.02878 0.02006 -0.0170 1.0000 0.0313
-8.250 -0.7946 0.02761 0.01870 -0.0155 1.0000 0.0348
-8.000 -0.7827 0.02618 0.01730 -0.0137 1.0000 0.0388
-7.750 -0.7686 0.02500 0.01603 -0.0118 1.0000 0.0432
-7.500 -0.7549 0.02382 0.01477 -0.0098 1.0000 0.0480
-7.250 -0.7403 0.02280 0.01373 -0.0081 1.0000 0.0568
-7.000 -0.7269 0.02168 0.01265 -0.0061 1.0000 0.0675
-6.750 -0.7125 0.02066 0.01162 -0.0041 1.0000 0.0834
-6.500 -0.6985 0.01966 0.01078 -0.0022 1.0000 0.1097
-6.250 -0.6850 0.01867 0.01002 -0.0003 1.0000 0.1480
-6.000 -0.6723 0.01769 0.00934 0.0018 1.0000 0.1989
-5.750 -0.6603 0.01673 0.00873 0.0040 1.0000 0.2623
-5.500 -0.6486 0.01588 0.00826 0.0063 1.0000 0.3300
-5.250 -0.6358 0.01532 0.00801 0.0088 1.0000 0.3938
-5.000 -0.6155 0.01501 0.00789 0.0099 0.9961 0.4514
-4.750 -0.5789 0.01489 0.00774 0.0079 0.9827 0.5022
-4.500 -0.5430 0.01484 0.00764 0.0062 0.9687 0.5361
-4.250 -0.5067 0.01480 0.00750 0.0045 0.9550 0.5623
-4.000 -0.4711 0.01476 0.00733 0.0030 0.9413 0.5847
-3.750 -0.4357 0.01472 0.00718 0.0016 0.9278 0.6025
-3.500 -0.4009 0.01468 0.00705 0.0004 0.9143 0.6184
-3.250 -0.3673 0.01464 0.00691 -0.0006 0.9009 0.6329
-3.000 -0.3351 0.01459 0.00676 -0.0013 0.8875 0.6460
-2.750 -0.3044 0.01454 0.00657 -0.0017 0.8740 0.6581
-2.500 -0.2750 0.01450 0.00644 -0.0018 0.8611 0.6691
-2.250 -0.2462 0.01446 0.00635 -0.0018 0.8486 0.6788
-2.000 -0.2183 0.01442 0.00621 -0.0016 0.8368 0.6887
-1.750 -0.1909 0.01438 0.00608 -0.0014 0.8258 0.6989
-1.500 -0.1632 0.01435 0.00602 -0.0012 0.8148 0.7069
-1.250 -0.1361 0.01432 0.00593 -0.0010 0.8040 0.7158
-1.000 -0.1087 0.01429 0.00585 -0.0008 0.7942 0.7241
-0.750 -0.0810 0.01427 0.00579 -0.0007 0.7852 0.7322
-0.500 -0.0545 0.01425 0.00574 -0.0004 0.7752 0.7413
-0.250 -0.0270 0.01424 0.00573 -0.0002 0.7664 0.7489
0.000 0.0000 0.01423 0.00570 0.0000 0.7579 0.7579
0.250 0.0270 0.01424 0.00573 0.0002 0.7489 0.7664
0.500 0.0545 0.01425 0.00574 0.0004 0.7412 0.7752
0.750 0.0810 0.01427 0.00579 0.0007 0.7322 0.7852
1.000 0.1087 0.01429 0.00585 0.0008 0.7241 0.7942
1.250 0.1361 0.01432 0.00593 0.0010 0.7158 0.8041
1.500 0.1631 0.01435 0.00602 0.0012 0.7069 0.8148
1.750 0.1909 0.01438 0.00608 0.0014 0.6989 0.8258
2.000 0.2183 0.01442 0.00623 0.0016 0.6887 0.8368
2.250 0.2462 0.01446 0.00635 0.0018 0.6788 0.8486
2.500 0.2749 0.01450 0.00644 0.0018 0.6690 0.8611
2.750 0.3044 0.01454 0.00657 0.0017 0.6581 0.8741
3.000 0.3350 0.01459 0.00676 0.0013 0.6460 0.8875
3.250 0.3672 0.01464 0.00690 0.0006 0.6329 0.9010
3.500 0.4008 0.01468 0.00704 -0.0004 0.6184 0.9144
3.750 0.4356 0.01472 0.00718 -0.0016 0.6025 0.9278
4.000 0.4711 0.01475 0.00733 -0.0030 0.5847 0.9414
4.250 0.5067 0.01480 0.00750 -0.0045 0.5623 0.9552
4.500 0.5428 0.01484 0.00764 -0.0062 0.5362 0.9689
4.750 0.5789 0.01489 0.00774 -0.0079 0.5019 0.9829
5.000 0.6155 0.01501 0.00789 -0.0099 0.4511 0.9963
5.250 0.6355 0.01531 0.00801 -0.0087 0.3940 1.0000
5.500 0.6484 0.01587 0.00826 -0.0063 0.3311 1.0000
5.750 0.6601 0.01672 0.00872 -0.0039 0.2626 1.0000
6.000 0.6722 0.01768 0.00933 -0.0017 0.1994 1.0000
6.250 0.6848 0.01867 0.01002 0.0003 0.1475 1.0000
6.500 0.6984 0.01966 0.01078 0.0022 0.1099 1.0000
6.750 0.7125 0.02066 0.01161 0.0041 0.0835 1.0000
7.000 0.7269 0.02168 0.01265 0.0061 0.0676 1.0000
7.250 0.7404 0.02280 0.01373 0.0081 0.0567 1.0000
7.500 0.7550 0.02383 0.01478 0.0098 0.0479 1.0000
7.750 0.7687 0.02501 0.01605 0.0118 0.0430 1.0000
8.000 0.7831 0.02618 0.01730 0.0136 0.0390 1.0000
8.250 0.7949 0.02761 0.01870 0.0154 0.0349 1.0000
8.500 0.8112 0.02878 0.02007 0.0170 0.0313 1.0000
8.750 0.8262 0.03027 0.02167 0.0185 0.0290 1.0000
9.000 0.8399 0.03172 0.02322 0.0200 0.0268 1.0000
9.250 0.8529 0.03381 0.02538 0.0214 0.0247 1.0000
9.500 0.8671 0.03537 0.02722 0.0229 0.0227 1.0000
9.750 0.8795 0.03735 0.02944 0.0244 0.0215 1.0000
10.000 0.8895 0.03956 0.03188 0.0259 0.0206 1.0000
10.250 0.8952 0.04177 0.03432 0.0278 0.0199 1.0000
10.500 0.8972 0.04405 0.03681 0.0299 0.0194 1.0000
10.750 0.8962 0.04648 0.03944 0.0319 0.0190 1.0000
11.000 0.8925 0.04912 0.04229 0.0334 0.0187 1.0000
11.250 0.8857 0.05214 0.04550 0.0345 0.0184 1.0000
11.500 0.8763 0.05555 0.04912 0.0350 0.0183 1.0000
11.750 0.8648 0.05925 0.05305 0.0348 0.0183 1.0000
12.000 0.8497 0.06371 0.05769 0.0337 0.0180 1.0000
12.250 0.8345 0.06838 0.06256 0.0320 0.0180 1.0000
12.500 0.8167 0.07384 0.06822 0.0292 0.0180 1.0000
12.750 0.7906 0.08140 0.07606 0.0244 0.0185 1.0000
13.000 0.7522 0.09309 0.08802 0.0162 0.0192 1.0000
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Polar data table (+)
Polar graphs
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