NACA 66 (p51htip-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 66 (p51htip-il) Reynolds number: 200,000 Max Cl/Cd: 47.99 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-p51htip-il-200000-n5.txt Download as CSV file: xf-p51htip-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5289 0.06967 0.06610 -0.0680 0.9694 0.0181
-10.250 -0.5448 0.06289 0.05914 -0.0733 0.9529 0.0179
-10.000 -0.5625 0.05803 0.05408 -0.0744 0.9376 0.0178
-9.750 -0.5795 0.05445 0.05029 -0.0726 0.9246 0.0177
-9.500 -0.5953 0.05127 0.04689 -0.0693 0.9144 0.0175
-9.250 -0.6061 0.04761 0.04294 -0.0664 0.9060 0.0175
-9.000 -0.6117 0.04404 0.03903 -0.0636 0.8993 0.0177
-8.750 -0.6126 0.04063 0.03525 -0.0608 0.8931 0.0178
-8.500 -0.6098 0.03742 0.03164 -0.0580 0.8884 0.0180
-8.250 -0.6022 0.03435 0.02813 -0.0558 0.8830 0.0184
-8.000 -0.5908 0.03177 0.02504 -0.0536 0.8783 0.0193
-7.750 -0.5764 0.02973 0.02255 -0.0517 0.8748 0.0199
-7.500 -0.5582 0.02773 0.02033 -0.0509 0.8716 0.0204
-7.250 -0.5374 0.02623 0.01864 -0.0502 0.8680 0.0209
-7.000 -0.5154 0.02490 0.01711 -0.0496 0.8649 0.0215
-6.750 -0.4924 0.02367 0.01568 -0.0490 0.8622 0.0222
-6.500 -0.4685 0.02251 0.01434 -0.0485 0.8599 0.0232
-6.250 -0.4437 0.02149 0.01315 -0.0481 0.8576 0.0243
-6.000 -0.4186 0.02083 0.01233 -0.0478 0.8546 0.0258
-5.750 -0.3941 0.01961 0.01105 -0.0476 0.8519 0.0274
-5.500 -0.3706 0.01876 0.01018 -0.0471 0.8492 0.0287
-5.250 -0.3473 0.01810 0.00946 -0.0465 0.8468 0.0302
-5.000 -0.3239 0.01759 0.00889 -0.0459 0.8447 0.0326
-4.750 -0.3006 0.01705 0.00829 -0.0453 0.8425 0.0342
-4.500 -0.2775 0.01658 0.00776 -0.0447 0.8398 0.0355
-4.250 -0.2571 0.01589 0.00705 -0.0437 0.8371 0.0378
-4.000 -0.2343 0.01549 0.00663 -0.0431 0.8348 0.0414
-3.750 -0.2106 0.01517 0.00624 -0.0426 0.8328 0.0452
-3.500 -0.1870 0.01483 0.00584 -0.0419 0.8310 0.0495
-3.250 -0.1633 0.01450 0.00548 -0.0413 0.8293 0.0586
-3.000 -0.1424 0.01397 0.00522 -0.0405 0.8265 0.1082
-2.750 -0.1303 0.01276 0.00490 -0.0386 0.8236 0.3183
-2.500 -0.1244 0.01143 0.00477 -0.0349 0.8208 0.5820
-2.250 -0.1059 0.01142 0.00526 -0.0322 0.8189 0.7393
-2.000 -0.0832 0.01173 0.00566 -0.0304 0.8172 0.7873
-1.750 -0.0595 0.01202 0.00594 -0.0288 0.8157 0.8133
-1.500 -0.0344 0.01232 0.00621 -0.0275 0.8143 0.8307
-1.250 -0.0106 0.01248 0.00633 -0.0268 0.8119 0.8410
-1.000 0.0160 0.01260 0.00642 -0.0266 0.8097 0.8450
-0.750 0.0423 0.01267 0.00647 -0.0265 0.8075 0.8492
-0.500 0.0681 0.01269 0.00644 -0.0263 0.8054 0.8537
-0.250 0.0945 0.01269 0.00640 -0.0263 0.8035 0.8573
0.000 0.1223 0.01272 0.00640 -0.0265 0.8020 0.8596
0.250 0.1501 0.01272 0.00639 -0.0266 0.8005 0.8622
0.500 0.1770 0.01275 0.00641 -0.0266 0.7987 0.8653
0.750 0.2005 0.01287 0.00655 -0.0263 0.7950 0.8694
1.000 0.2260 0.01293 0.00663 -0.0261 0.7921 0.8725
1.250 0.2532 0.01296 0.00669 -0.0261 0.7896 0.8748
1.500 0.2809 0.01296 0.00671 -0.0262 0.7875 0.8775
1.750 0.3091 0.01294 0.00671 -0.0263 0.7855 0.8808
2.000 0.3326 0.01303 0.00687 -0.0259 0.7814 0.8851
2.250 0.3576 0.01310 0.00700 -0.0255 0.7771 0.8881
2.500 0.3855 0.01306 0.00702 -0.0255 0.7737 0.8907
2.750 0.4151 0.01293 0.00695 -0.0258 0.7708 0.8934
3.000 0.4376 0.01297 0.00708 -0.0249 0.7636 0.8978
3.250 0.4660 0.01248 0.00661 -0.0244 0.7520 0.9016
3.500 0.4921 0.01195 0.00613 -0.0232 0.7316 0.9045
3.750 0.5164 0.01150 0.00568 -0.0218 0.7017 0.9082
4.000 0.5365 0.01118 0.00526 -0.0197 0.6328 0.9133
4.250 0.5405 0.01187 0.00491 -0.0147 0.4460 0.9194
4.500 0.5399 0.01305 0.00543 -0.0100 0.3101 0.9281
4.750 0.5434 0.01427 0.00603 -0.0064 0.1795 0.9363
5.000 0.5507 0.01540 0.00665 -0.0035 0.0777 0.9459
5.500 0.5931 0.01661 0.00775 -0.0023 0.0435 0.9598
5.750 0.6176 0.01731 0.00847 -0.0024 0.0387 0.9660
6.000 0.6429 0.01784 0.00910 -0.0027 0.0352 0.9733
6.250 0.6700 0.01852 0.00984 -0.0035 0.0323 0.9797
6.500 0.6950 0.01937 0.01073 -0.0040 0.0303 0.9876
6.750 0.7181 0.02057 0.01197 -0.0043 0.0286 0.9982
7.000 0.7289 0.02107 0.01252 -0.0018 0.0277 1.0000
7.250 0.7427 0.02163 0.01313 0.0002 0.0263 1.0000
7.500 0.7593 0.02234 0.01388 0.0016 0.0250 1.0000
7.750 0.7780 0.02319 0.01477 0.0027 0.0239 1.0000
8.000 0.7984 0.02411 0.01574 0.0034 0.0231 1.0000
8.250 0.8209 0.02519 0.01684 0.0039 0.0224 1.0000
8.500 0.8499 0.02680 0.01850 0.0032 0.0216 1.0000
8.750 0.8797 0.02852 0.02038 0.0024 0.0209 1.0000
9.000 0.9012 0.02969 0.02177 0.0031 0.0202 1.0000
9.250 0.9220 0.03109 0.02340 0.0037 0.0193 1.0000
9.500 0.9419 0.03279 0.02533 0.0044 0.0186 1.0000
9.750 0.9588 0.03466 0.02745 0.0055 0.0181 1.0000
10.000 0.9718 0.03658 0.02962 0.0071 0.0177 1.0000
10.250 0.9816 0.03850 0.03177 0.0090 0.0174 1.0000
10.500 0.9882 0.04057 0.03406 0.0111 0.0172 1.0000
10.750 0.9908 0.04290 0.03667 0.0136 0.0170 1.0000
11.000 0.9907 0.04516 0.03916 0.0162 0.0168 1.0000
11.250 0.9894 0.04725 0.04142 0.0185 0.0166 1.0000
11.500 0.9860 0.04953 0.04388 0.0208 0.0164 1.0000
11.750 0.9804 0.05204 0.04654 0.0227 0.0161 1.0000
12.000 0.9694 0.05531 0.05000 0.0247 0.0159 1.0000
12.250 0.9549 0.05877 0.05368 0.0264 0.0159 1.0000
12.500 0.9352 0.06277 0.05795 0.0277 0.0158 1.0000
12.750 0.9153 0.06722 0.06265 0.0281 0.0157 1.0000
13.000 0.8930 0.07237 0.06802 0.0276 0.0157 1.0000
13.250 0.8703 0.07818 0.07401 0.0258 0.0157 1.0000
13.500 0.8521 0.08406 0.08001 0.0233 0.0158 1.0000
13.750 0.8280 0.09191 0.08802 0.0188 0.0158 1.0000
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Polar data table (+)
Polar graphs
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