NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NLR-1T AIRFOIL (nlr1t-il) Reynolds number: 100,000 Max Cl/Cd: 39.51 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-nlr1t-il-100000-n5.txt Download as CSV file: xf-nlr1t-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NLR-1T 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5269 0.09033 0.08549 -0.0232 1.0000 0.0581
-8.500 -0.5318 0.08574 0.08095 -0.0270 1.0000 0.0589
-8.250 -0.5403 0.08144 0.07668 -0.0297 1.0000 0.0595
-8.000 -0.5473 0.07736 0.07259 -0.0312 1.0000 0.0602
-7.750 -0.5805 0.07545 0.07028 -0.0321 1.0000 0.0640
-7.500 -0.5218 0.05604 0.05132 -0.0314 1.0000 0.0374
-7.250 -0.5740 0.06275 0.05754 -0.0284 1.0000 0.0368
-7.000 -0.5780 0.05969 0.05441 -0.0256 1.0000 0.0353
-6.750 -0.5812 0.05636 0.05092 -0.0227 1.0000 0.0337
-6.250 -0.5805 0.04822 0.04177 -0.0154 1.0000 0.0283
-6.000 -0.5746 0.04502 0.03838 -0.0129 1.0000 0.0279
-5.750 -0.5665 0.04210 0.03520 -0.0104 1.0000 0.0276
-5.500 -0.5561 0.03933 0.03212 -0.0080 1.0000 0.0274
-5.250 -0.5434 0.03673 0.02917 -0.0056 1.0000 0.0273
-5.000 -0.5280 0.03465 0.02661 -0.0033 1.0000 0.0280
-4.750 -0.5122 0.03230 0.02385 -0.0013 1.0000 0.0290
-4.500 -0.4942 0.03014 0.02143 0.0002 1.0000 0.0297
-4.250 -0.4735 0.02830 0.01923 0.0017 1.0000 0.0296
-4.000 -0.4515 0.02663 0.01727 0.0030 1.0000 0.0296
-3.750 -0.4285 0.02511 0.01552 0.0041 1.0000 0.0298
-3.500 -0.4026 0.02374 0.01394 0.0046 0.9994 0.0302
-3.250 -0.3750 0.02254 0.01260 0.0047 0.9984 0.0308
-3.000 -0.3456 0.02148 0.01145 0.0044 0.9967 0.0316
-2.750 -0.3149 0.02056 0.01048 0.0038 0.9948 0.0329
-2.500 -0.2856 0.01990 0.00975 0.0034 0.9924 0.0361
-2.250 -0.2572 0.01930 0.00906 0.0033 0.9900 0.0391
-2.000 -0.2289 0.01859 0.00830 0.0030 0.9879 0.0422
-1.750 -0.1981 0.01816 0.00778 0.0023 0.9857 0.0473
-1.500 -0.1684 0.01768 0.00729 0.0019 0.9829 0.0598
-1.250 -0.0730 0.01484 0.00787 -0.0107 0.9965 1.0000
-1.000 -0.0404 0.01491 0.00771 -0.0121 0.9926 1.0000
-0.750 -0.0066 0.01499 0.00760 -0.0137 0.9879 1.0000
-0.500 0.0295 0.01509 0.00755 -0.0158 0.9832 1.0000
-0.250 0.0672 0.01517 0.00751 -0.0182 0.9769 1.0000
0.000 0.1161 0.01518 0.00741 -0.0227 0.9670 1.0000
0.250 0.2212 0.01451 0.00664 -0.0375 0.9391 1.0000
0.500 0.2744 0.01407 0.00614 -0.0419 0.9192 1.0000
0.750 0.3034 0.01398 0.00604 -0.0419 0.9070 1.0000
1.000 0.3307 0.01390 0.00596 -0.0415 0.8944 1.0000
1.250 0.3568 0.01380 0.00587 -0.0407 0.8800 1.0000
1.500 0.3817 0.01368 0.00575 -0.0396 0.8629 1.0000
1.750 0.4033 0.01358 0.00568 -0.0379 0.8411 1.0000
2.000 0.4243 0.01348 0.00562 -0.0360 0.8131 1.0000
2.250 0.4410 0.01347 0.00565 -0.0335 0.7648 1.0000
2.500 0.4731 0.01302 0.00463 -0.0321 0.6305 1.0000
2.750 0.4926 0.01364 0.00457 -0.0298 0.5049 1.0000
3.000 0.5115 0.01435 0.00482 -0.0281 0.4133 1.0000
3.250 0.5319 0.01497 0.00509 -0.0268 0.3487 1.0000
3.500 0.5534 0.01551 0.00538 -0.0257 0.3062 1.0000
3.750 0.5754 0.01600 0.00571 -0.0247 0.2786 1.0000
4.000 0.5977 0.01645 0.00610 -0.0237 0.2603 1.0000
4.250 0.6200 0.01690 0.00649 -0.0228 0.2463 1.0000
4.500 0.6425 0.01735 0.00690 -0.0218 0.2348 1.0000
4.750 0.6647 0.01783 0.00733 -0.0209 0.2252 1.0000
5.000 0.6871 0.01827 0.00778 -0.0199 0.2149 1.0000
5.250 0.7094 0.01871 0.00827 -0.0190 0.2036 1.0000
5.500 0.7314 0.01914 0.00871 -0.0181 0.1913 1.0000
5.750 0.7531 0.01957 0.00914 -0.0171 0.1789 1.0000
6.000 0.7746 0.02002 0.00960 -0.0161 0.1682 1.0000
6.250 0.7965 0.02043 0.01014 -0.0152 0.1571 1.0000
6.500 0.8183 0.02089 0.01071 -0.0142 0.1466 1.0000
6.750 0.8395 0.02136 0.01127 -0.0131 0.1344 1.0000
7.000 0.8606 0.02182 0.01182 -0.0120 0.1196 1.0000
7.250 0.8815 0.02231 0.01239 -0.0109 0.1035 1.0000
7.500 0.9016 0.02295 0.01308 -0.0097 0.0914 1.0000
7.750 0.9206 0.02376 0.01391 -0.0083 0.0831 1.0000
8.000 0.9390 0.02464 0.01482 -0.0069 0.0771 1.0000
8.250 0.9569 0.02566 0.01592 -0.0054 0.0727 1.0000
8.500 0.9752 0.02666 0.01707 -0.0039 0.0685 1.0000
8.750 0.9923 0.02767 0.01815 -0.0024 0.0649 1.0000
9.000 1.0087 0.02891 0.01951 -0.0009 0.0621 1.0000
9.250 1.0263 0.03019 0.02103 0.0006 0.0600 1.0000
9.500 1.0430 0.03158 0.02265 0.0022 0.0580 1.0000
9.750 1.0589 0.03304 0.02433 0.0038 0.0563 1.0000
10.000 1.0737 0.03455 0.02601 0.0054 0.0545 1.0000
10.250 1.0873 0.03609 0.02765 0.0070 0.0526 1.0000
10.500 1.0983 0.03810 0.02988 0.0089 0.0507 1.0000
10.750 1.1066 0.04011 0.03235 0.0112 0.0489 1.0000
11.000 1.1117 0.04204 0.03464 0.0137 0.0465 1.0000
11.250 1.1175 0.04309 0.03581 0.0159 0.0433 1.0000
11.500 1.1206 0.04458 0.03722 0.0181 0.0403 1.0000
11.750 1.1129 0.04636 0.03944 0.0218 0.0381 1.0000
12.000 1.1048 0.04858 0.04198 0.0247 0.0360 1.0000
12.250 1.0971 0.05098 0.04461 0.0269 0.0343 1.0000
12.500 1.0903 0.05336 0.04714 0.0284 0.0328 1.0000
12.750 1.0847 0.05577 0.04961 0.0293 0.0314 1.0000
13.250 1.0539 0.06443 0.05860 0.0284 0.0300 1.0000
13.500 1.0325 0.07078 0.06519 0.0254 0.0300 1.0000
13.750 1.0098 0.07846 0.07306 0.0206 0.0302 1.0000
14.000 0.8886 0.11217 0.10710 -0.0003 0.0403 1.0000
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Polar data table (+)
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