NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il) Reynolds number: 100,000 Max Cl/Cd: 27.15 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-nlf415-il-100000.txt Download as CSV file: xf-nlf415-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY NLF(2)-0415 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6564 0.10599 0.10183 -0.0357 1.0000 0.1209
-9.000 -0.6902 0.10134 0.09718 -0.0363 1.0000 0.1211
-8.750 -0.7214 0.09765 0.09349 -0.0353 1.0000 0.1210
-8.500 -0.7525 0.09453 0.09037 -0.0330 1.0000 0.1208
-8.250 -0.7882 0.09112 0.08682 -0.0315 1.0000 0.1214
-8.000 -0.8271 0.08870 0.08404 -0.0297 1.0000 0.1224
-7.750 -0.6616 0.07937 0.07562 -0.0284 1.0000 0.1599
-7.500 -0.8136 0.07951 0.07502 -0.0273 1.0000 0.1346
-7.250 -0.8217 0.07504 0.07045 -0.0263 1.0000 0.1412
-5.500 -0.7967 0.05191 0.04692 -0.0181 1.0000 0.2881
-5.250 -0.7829 0.04871 0.04383 -0.0165 1.0000 0.3094
-5.000 -0.7601 0.04466 0.03932 -0.0184 1.0000 0.3122
-4.750 -0.6664 0.03974 0.03106 -0.0185 1.0000 0.0747
-4.500 -0.6394 0.03894 0.02955 -0.0167 1.0000 0.0659
-4.250 -0.6168 0.03651 0.02698 -0.0158 1.0000 0.0648
-4.000 -0.5940 0.03514 0.02540 -0.0148 1.0000 0.0651
-3.750 -0.5713 0.03375 0.02388 -0.0137 1.0000 0.0651
-3.500 -0.5490 0.03242 0.02252 -0.0125 1.0000 0.0645
-3.250 -0.5271 0.03143 0.02149 -0.0112 1.0000 0.0646
-3.000 -0.5058 0.03064 0.02068 -0.0099 1.0000 0.0649
-2.750 -0.4847 0.03000 0.02004 -0.0088 1.0000 0.0658
-2.500 -0.4648 0.02909 0.01928 -0.0077 1.0000 0.0685
-2.250 -0.4427 0.02862 0.01883 -0.0073 1.0000 0.0752
-2.000 -0.4185 0.02813 0.01822 -0.0074 1.0000 0.0827
-1.750 -0.3928 0.02764 0.01768 -0.0079 1.0000 0.0977
-1.500 -0.3788 0.02800 0.02040 -0.0041 1.0000 0.6514
-1.250 -0.3805 0.02999 0.02261 0.0055 1.0000 0.6969
-1.000 -0.3841 0.03128 0.02403 0.0150 1.0000 0.7294
-0.750 -0.3876 0.03211 0.02493 0.0240 1.0000 0.7596
-0.500 -0.3822 0.03244 0.02523 0.0297 1.0000 0.7794
-0.250 -0.3808 0.03265 0.02543 0.0365 1.0000 0.8029
0.000 -0.3634 0.03276 0.02543 0.0379 1.0000 0.8143
0.250 -0.3434 0.03264 0.02520 0.0383 1.0000 0.8166
0.500 -0.3174 0.03265 0.02507 0.0369 1.0000 0.8166
0.750 -0.2907 0.03272 0.02502 0.0352 1.0000 0.8162
1.000 -0.2641 0.03282 0.02503 0.0336 1.0000 0.8159
1.250 -0.2381 0.03297 0.02508 0.0321 1.0000 0.8157
1.500 -0.2132 0.03314 0.02519 0.0309 1.0000 0.8159
1.750 -0.1892 0.03334 0.02535 0.0300 1.0000 0.8165
2.000 -0.1648 0.03360 0.02556 0.0289 1.0000 0.8168
2.250 -0.1403 0.03390 0.02582 0.0277 1.0000 0.8170
2.500 -0.1162 0.03424 0.02614 0.0266 1.0000 0.8173
2.750 -0.0925 0.03462 0.02651 0.0256 1.0000 0.8177
3.000 -0.0609 0.03551 0.02740 0.0230 0.9967 0.8182
3.250 -0.0297 0.03650 0.02840 0.0205 0.9935 0.8189
3.500 0.0021 0.03745 0.02937 0.0179 0.9877 0.8200
3.750 0.0343 0.03876 0.03072 0.0152 0.9840 0.8214
4.000 0.0640 0.03944 0.03145 0.0129 0.9761 0.8226
4.250 0.0942 0.04055 0.03263 0.0106 0.9706 0.8237
4.500 0.1295 0.04177 0.03391 0.0073 0.9616 0.8249
4.750 0.1572 0.04244 0.03468 0.0053 0.9511 0.8260
5.000 0.1891 0.04346 0.03579 0.0027 0.9394 0.8273
5.250 0.2341 0.04478 0.03722 -0.0019 0.9210 0.8288
5.500 0.3841 0.04266 0.03538 -0.0172 0.8291 0.8306
5.750 0.4247 0.04168 0.03460 -0.0187 0.8067 0.8324
6.000 0.4839 0.04028 0.03344 -0.0225 0.7890 0.8348
6.250 0.5377 0.03830 0.03174 -0.0250 0.7697 0.8374
6.500 0.5952 0.03524 0.02898 -0.0271 0.7467 0.8403
7.000 0.6950 0.02566 0.01749 -0.0220 0.2776 0.8453
7.250 0.7108 0.02709 0.01812 -0.0205 0.2002 0.8475
7.500 0.7354 0.02810 0.01889 -0.0203 0.1674 0.8502
7.750 0.7655 0.02908 0.01973 -0.0210 0.1455 0.8533
8.000 0.8105 0.03030 0.02078 -0.0238 0.1288 0.8565
8.250 0.8656 0.03197 0.02234 -0.0284 0.1132 0.8596
8.500 0.8977 0.03306 0.02377 -0.0291 0.1037 0.8633
8.750 0.9440 0.03513 0.02602 -0.0324 0.0925 0.8674
9.000 0.9799 0.03732 0.02839 -0.0344 0.0819 0.8721
9.250 1.0126 0.03974 0.03088 -0.0361 0.0717 0.8768
9.500 1.0239 0.04130 0.03301 -0.0332 0.0659 0.8831
9.750 1.0594 0.04547 0.03727 -0.0358 0.0595 0.8887
10.000 1.0578 0.04697 0.03938 -0.0308 0.0572 0.8964
10.250 1.0605 0.04967 0.04261 -0.0271 0.0548 0.9051
10.750 1.0542 0.05630 0.05020 -0.0191 0.0538 0.9341
11.000 1.0455 0.06032 0.05469 -0.0156 0.0544 1.0000
11.250 1.0334 0.06475 0.05948 -0.0125 0.0552 1.0000
11.500 1.0185 0.06919 0.06421 -0.0098 0.0561 1.0000
11.750 1.0026 0.07372 0.06896 -0.0075 0.0571 1.0000
12.000 0.9892 0.07852 0.07393 -0.0061 0.0581 1.0000
12.250 0.9887 0.08378 0.07933 -0.0058 0.0608 1.0000
12.500 0.8977 0.09235 0.08844 -0.0037 0.0655 1.0000
12.750 0.8649 0.10011 0.09632 -0.0061 0.0680 1.0000
13.000 0.6958 0.09819 0.09488 0.0045 0.0683 1.0000
13.250 0.6727 0.10446 0.10119 0.0015 0.0706 1.0000
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Polar data table (+)
Polar graphs
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