NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NLR-1T AIRFOIL (nlr1t-il) Reynolds number: 50,000 Max Cl/Cd: 28.28 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-nlr1t-il-50000-n5.txt Download as CSV file: xf-nlr1t-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5273 0.11010 0.10313 -0.0132 1.0000 0.1253
-9.250 -0.5256 0.10655 0.09963 -0.0144 1.0000 0.1297
-9.000 -0.5612 0.10372 0.09698 -0.0231 1.0000 0.1350
-8.750 -0.5291 0.09913 0.09235 -0.0174 1.0000 0.1397
-8.500 -0.5472 0.08729 0.08039 -0.0308 1.0000 0.0627
-8.250 -0.5464 0.08345 0.07657 -0.0307 1.0000 0.0612
-8.000 -0.5492 0.07950 0.07263 -0.0311 1.0000 0.0599
-7.750 -0.5532 0.07559 0.06869 -0.0312 1.0000 0.0586
-7.500 -0.5576 0.07177 0.06480 -0.0307 1.0000 0.0573
-7.250 -0.5622 0.06802 0.06093 -0.0296 1.0000 0.0559
-6.750 -0.5742 0.06096 0.05296 -0.0249 1.0000 0.0509
-6.500 -0.5687 0.05761 0.04952 -0.0228 1.0000 0.0504
-6.250 -0.5628 0.05443 0.04621 -0.0206 1.0000 0.0498
-6.000 -0.5561 0.05140 0.04297 -0.0183 1.0000 0.0491
-5.750 -0.5478 0.04848 0.03976 -0.0160 1.0000 0.0485
-5.500 -0.5376 0.04565 0.03657 -0.0136 1.0000 0.0480
-5.250 -0.5254 0.04295 0.03350 -0.0114 1.0000 0.0475
-5.000 -0.5110 0.04040 0.03057 -0.0093 1.0000 0.0472
-4.750 -0.4945 0.03801 0.02779 -0.0073 1.0000 0.0470
-4.500 -0.4759 0.03579 0.02517 -0.0055 1.0000 0.0470
-4.250 -0.4555 0.03375 0.02275 -0.0039 1.0000 0.0473
-4.000 -0.4337 0.03205 0.02064 -0.0025 1.0000 0.0489
-3.750 -0.4108 0.03071 0.01886 -0.0010 1.0000 0.0510
-3.500 -0.3873 0.02887 0.01694 -0.0003 1.0000 0.0533
-3.250 -0.3624 0.02750 0.01544 0.0006 1.0000 0.0550
-3.000 -0.3367 0.02632 0.01411 0.0015 1.0000 0.0570
-2.750 -0.3109 0.02530 0.01292 0.0023 1.0000 0.0598
-2.500 -0.2854 0.02446 0.01183 0.0031 1.0000 0.0634
-2.250 -0.2617 0.02360 0.01092 0.0037 1.0000 0.0715
-2.000 -0.1193 0.01897 0.00989 -0.0164 1.0000 1.0000
-1.750 -0.1057 0.01894 0.00946 -0.0140 1.0000 1.0000
-1.500 -0.0913 0.01892 0.00916 -0.0117 1.0000 1.0000
-1.250 -0.0763 0.01894 0.00891 -0.0096 1.0000 1.0000
-1.000 -0.0607 0.01897 0.00872 -0.0076 1.0000 1.0000
-0.750 -0.0446 0.01903 0.00858 -0.0057 1.0000 1.0000
-0.500 -0.0282 0.01912 0.00847 -0.0038 1.0000 1.0000
-0.250 -0.0114 0.01922 0.00842 -0.0021 1.0000 1.0000
0.000 0.0057 0.01934 0.00842 -0.0004 1.0000 1.0000
0.250 0.0232 0.01949 0.00845 0.0012 1.0000 1.0000
0.500 0.0409 0.01967 0.00852 0.0027 1.0000 1.0000
0.750 0.0589 0.01986 0.00865 0.0041 1.0000 1.0000
1.000 0.0770 0.02009 0.00882 0.0054 1.0000 1.0000
1.250 0.0952 0.02035 0.00905 0.0066 1.0000 1.0000
1.500 0.1570 0.02087 0.00957 -0.0009 0.9829 1.0000
1.750 0.2451 0.02108 0.00986 -0.0129 0.9499 1.0000
2.000 0.3210 0.02089 0.00981 -0.0218 0.9201 1.0000
2.250 0.3719 0.02065 0.00971 -0.0257 0.8946 1.0000
2.500 0.4099 0.02039 0.00960 -0.0268 0.8676 1.0000
2.750 0.4390 0.02012 0.00946 -0.0261 0.8357 1.0000
3.000 0.4560 0.01996 0.00944 -0.0232 0.7931 1.0000
3.250 0.4690 0.01982 0.00940 -0.0196 0.7250 1.0000
3.500 0.5331 0.01897 0.00789 -0.0225 0.5615 1.0000
3.750 0.5517 0.01982 0.00804 -0.0200 0.4563 1.0000
4.000 0.5685 0.02071 0.00852 -0.0179 0.3953 1.0000
4.250 0.5883 0.02155 0.00913 -0.0165 0.3566 1.0000
4.500 0.6098 0.02233 0.00975 -0.0154 0.3290 1.0000
4.750 0.6327 0.02306 0.01040 -0.0145 0.3076 1.0000
5.000 0.6568 0.02378 0.01110 -0.0139 0.2908 1.0000
5.250 0.6820 0.02451 0.01189 -0.0134 0.2761 1.0000
5.500 0.7075 0.02524 0.01268 -0.0130 0.2619 1.0000
5.750 0.7329 0.02600 0.01351 -0.0125 0.2485 1.0000
6.000 0.7569 0.02678 0.01437 -0.0119 0.2342 1.0000
6.250 0.7779 0.02751 0.01514 -0.0109 0.2169 1.0000
6.500 0.7976 0.02822 0.01592 -0.0097 0.1995 1.0000
6.750 0.8180 0.02904 0.01680 -0.0086 0.1840 1.0000
7.000 0.8384 0.02995 0.01779 -0.0075 0.1691 1.0000
7.250 0.8573 0.03089 0.01892 -0.0062 0.1529 1.0000
7.500 0.8767 0.03200 0.02015 -0.0050 0.1392 1.0000
7.750 0.8967 0.03334 0.02170 -0.0038 0.1277 1.0000
8.000 0.9160 0.03489 0.02355 -0.0026 0.1172 1.0000
8.250 0.9337 0.03626 0.02511 -0.0012 0.1087 1.0000
8.500 0.9506 0.03748 0.02648 0.0002 0.1016 1.0000
8.750 0.9675 0.03949 0.02880 0.0016 0.0967 1.0000
9.000 0.9827 0.04170 0.03138 0.0031 0.0926 1.0000
9.250 0.9976 0.04361 0.03352 0.0046 0.0896 1.0000
9.500 1.0108 0.04571 0.03580 0.0062 0.0869 1.0000
9.750 1.0140 0.04908 0.03976 0.0086 0.0838 1.0000
10.000 1.0164 0.05191 0.04297 0.0108 0.0806 1.0000
10.250 1.0212 0.05417 0.04546 0.0129 0.0779 1.0000
10.500 1.0253 0.05667 0.04814 0.0148 0.0762 1.0000
10.750 1.0253 0.05967 0.05129 0.0168 0.0750 1.0000
11.000 1.0061 0.06406 0.05606 0.0195 0.0746 1.0000
11.250 0.9811 0.06838 0.06062 0.0222 0.0745 1.0000
11.500 0.9533 0.07344 0.06586 0.0231 0.0745 1.0000
11.750 0.9233 0.07967 0.07222 0.0218 0.0747 1.0000
12.000 0.8920 0.08757 0.08018 0.0178 0.0750 1.0000
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Polar data table (+)
Polar graphs
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