NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NLR-1T AIRFOIL (nlr1t-il) Reynolds number: 100,000 Max Cl/Cd: 38.93 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-nlr1t-il-100000.txt Download as CSV file: xf-nlr1t-il-100000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4155 0.10528 0.10056 -0.0216 1.0000 0.0939
-9.750 -0.4342 0.10193 0.09728 -0.0258 1.0000 0.0953
-9.500 -0.4538 0.09814 0.09356 -0.0303 1.0000 0.0957
-9.250 -0.4160 0.09327 0.08866 -0.0236 1.0000 0.1007
-9.000 -0.4162 0.08962 0.08503 -0.0242 1.0000 0.1046
-8.750 -0.4295 0.08556 0.08104 -0.0271 1.0000 0.1082
-8.500 -0.4549 0.08114 0.07670 -0.0322 1.0000 0.1094
-8.250 -0.4826 0.07750 0.07310 -0.0339 1.0000 0.1097
-8.000 -0.4433 0.07318 0.06881 -0.0289 1.0000 0.1163
-7.750 -0.4592 0.06950 0.06520 -0.0295 1.0000 0.1186
-7.500 -0.4824 0.06673 0.06248 -0.0275 1.0000 0.1201
-7.250 -0.5083 0.06394 0.05969 -0.0255 1.0000 0.1222
-7.000 -0.5449 0.06199 0.05754 -0.0230 1.0000 0.1241
-6.750 -0.5350 0.05706 0.05277 -0.0209 1.0000 0.1277
-6.500 -0.5340 0.05426 0.04997 -0.0183 1.0000 0.1331
-6.250 -0.5961 0.06364 0.05827 -0.0168 1.0000 0.1393
-6.000 -0.5802 0.05821 0.05322 -0.0151 1.0000 0.1433
-5.750 -0.5802 0.05575 0.05056 -0.0130 1.0000 0.1557
-5.500 -0.5757 0.05341 0.04808 -0.0107 1.0000 0.1697
-5.250 -0.5679 0.05074 0.04538 -0.0085 1.0000 0.1847
-5.000 -0.5583 0.04810 0.04277 -0.0062 1.0000 0.2009
-3.750 -0.4474 0.03048 0.02189 0.0026 1.0000 0.0803
-3.500 -0.4230 0.02757 0.01870 0.0037 1.0000 0.0737
-3.250 -0.3970 0.02673 0.01720 0.0058 1.0000 0.0679
-3.000 -0.3705 0.02457 0.01485 0.0064 1.0000 0.0667
-2.750 -0.3437 0.02307 0.01318 0.0071 1.0000 0.0665
-2.500 -0.3172 0.02173 0.01173 0.0077 1.0000 0.0692
-2.250 -0.2916 0.02051 0.01058 0.0083 1.0000 0.0729
-2.000 -0.2668 0.01958 0.00965 0.0092 1.0000 0.0761
-1.750 -0.1184 0.01477 0.00838 -0.0121 1.0000 1.0000
-1.500 -0.1039 0.01478 0.00806 -0.0097 1.0000 1.0000
-1.250 -0.0888 0.01481 0.00786 -0.0076 1.0000 1.0000
-1.000 -0.0729 0.01486 0.00773 -0.0055 1.0000 1.0000
-0.750 -0.0565 0.01493 0.00764 -0.0036 1.0000 1.0000
-0.500 -0.0397 0.01502 0.00759 -0.0018 1.0000 1.0000
-0.250 -0.0225 0.01514 0.00758 -0.0001 1.0000 1.0000
0.000 -0.0051 0.01527 0.00762 0.0016 1.0000 1.0000
0.250 0.0127 0.01543 0.00769 0.0031 1.0000 1.0000
0.500 0.0307 0.01561 0.00779 0.0046 1.0000 1.0000
0.750 0.0489 0.01582 0.00795 0.0059 1.0000 1.0000
1.000 0.0673 0.01606 0.00815 0.0072 1.0000 1.0000
1.250 0.0857 0.01633 0.00840 0.0084 1.0000 1.0000
1.500 0.1641 0.01689 0.00896 -0.0024 0.9835 1.0000
1.750 0.2290 0.01716 0.00929 -0.0103 0.9685 1.0000
2.000 0.2929 0.01718 0.00941 -0.0175 0.9507 1.0000
2.250 0.3647 0.01693 0.00930 -0.0258 0.9319 1.0000
2.500 0.4490 0.01618 0.00879 -0.0359 0.9123 1.0000
2.750 0.4991 0.01534 0.00812 -0.0384 0.8830 1.0000
3.000 0.5273 0.01452 0.00743 -0.0362 0.8406 1.0000
3.250 0.5380 0.01394 0.00691 -0.0308 0.7440 1.0000
3.500 0.5649 0.01451 0.00569 -0.0272 0.4518 1.0000
3.750 0.5837 0.01563 0.00617 -0.0255 0.3778 1.0000
4.000 0.6057 0.01643 0.00671 -0.0244 0.3442 1.0000
4.250 0.6289 0.01713 0.00725 -0.0235 0.3223 1.0000
4.500 0.6527 0.01782 0.00782 -0.0228 0.3061 1.0000
4.750 0.6768 0.01849 0.00844 -0.0221 0.2923 1.0000
5.000 0.7005 0.01914 0.00911 -0.0213 0.2787 1.0000
5.250 0.7231 0.01974 0.00966 -0.0205 0.2631 1.0000
5.500 0.7452 0.02032 0.01020 -0.0195 0.2472 1.0000
5.750 0.7675 0.02093 0.01083 -0.0186 0.2330 1.0000
6.000 0.7897 0.02160 0.01158 -0.0177 0.2191 1.0000
6.250 0.8112 0.02230 0.01228 -0.0167 0.2035 1.0000
6.500 0.8316 0.02306 0.01305 -0.0155 0.1854 1.0000
6.750 0.8508 0.02389 0.01401 -0.0140 0.1646 1.0000
7.000 0.8711 0.02499 0.01506 -0.0127 0.1467 1.0000
7.250 0.8915 0.02615 0.01631 -0.0114 0.1327 1.0000
7.500 0.9117 0.02732 0.01765 -0.0100 0.1218 1.0000
7.750 0.9329 0.02862 0.01895 -0.0090 0.1145 1.0000
8.000 0.9530 0.03015 0.02082 -0.0075 0.1085 1.0000
8.250 0.9742 0.03156 0.02228 -0.0065 0.1042 1.0000
8.500 0.9916 0.03386 0.02498 -0.0049 0.1010 1.0000
8.750 1.0063 0.03599 0.02756 -0.0028 0.0972 1.0000
9.000 1.0236 0.03742 0.02912 -0.0014 0.0934 1.0000
9.250 1.0405 0.03970 0.03149 -0.0001 0.0910 1.0000
9.500 1.0464 0.04316 0.03547 0.0026 0.0898 1.0000
9.750 1.0461 0.04703 0.03992 0.0059 0.0890 1.0000
10.000 1.0394 0.05127 0.04469 0.0095 0.0881 1.0000
10.250 1.0268 0.05576 0.04963 0.0130 0.0872 1.0000
10.500 1.0088 0.06049 0.05472 0.0163 0.0869 1.0000
10.750 0.9855 0.06538 0.05987 0.0192 0.0872 1.0000
11.000 0.9577 0.07001 0.06466 0.0220 0.0879 1.0000
11.250 0.9309 0.07525 0.07000 0.0228 0.0888 1.0000
11.500 0.9077 0.08113 0.07595 0.0219 0.0896 1.0000
11.750 0.8906 0.08743 0.08229 0.0200 0.0902 1.0000
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Polar data table (+)
Polar graphs
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