lrn1007 (lrn1007-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: lrn1007 (lrn1007-il) Reynolds number: 100,000 Max Cl/Cd: 51.34 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-lrn1007-il-100000.txt Download as CSV file: xf-lrn1007-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: lrn1007
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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3964 0.10979 0.10547 -0.0166 1.0000 0.0191
-7.250 -0.4064 0.10908 0.10481 -0.0148 1.0000 0.0193
-7.000 -0.4137 0.10798 0.10375 -0.0137 1.0000 0.0194
-6.750 -0.4161 0.10651 0.10232 -0.0139 1.0000 0.0195
-6.500 -0.4160 0.10487 0.10070 -0.0149 1.0000 0.0196
-6.250 -0.4130 0.10307 0.09891 -0.0164 1.0000 0.0198
-6.000 -0.4067 0.10125 0.09709 -0.0188 1.0000 0.0199
-5.750 -0.3975 0.09928 0.09510 -0.0216 1.0000 0.0200
-5.500 -0.3857 0.09714 0.09291 -0.0244 1.0000 0.0201
-5.250 -0.3723 0.09488 0.09059 -0.0269 1.0000 0.0202
-5.000 -0.3572 0.09260 0.08824 -0.0294 1.0000 0.0203
-4.750 -0.3616 0.08531 0.08111 -0.0267 1.0000 0.0209
-4.500 -0.3646 0.08005 0.07592 -0.0238 1.0000 0.0219
-4.250 -0.3559 0.07630 0.07217 -0.0243 1.0000 0.0231
-4.000 -0.3412 0.07295 0.06878 -0.0262 1.0000 0.0244
-3.750 -0.3227 0.06969 0.06546 -0.0287 1.0000 0.0258
-3.500 -0.3009 0.06649 0.06218 -0.0316 1.0000 0.0274
-3.250 -0.2766 0.06349 0.05908 -0.0346 1.0000 0.0289
-3.000 -0.2499 0.06058 0.05604 -0.0375 1.0000 0.0299
-2.750 -0.2209 0.05818 0.05348 -0.0404 1.0000 0.0311
-2.500 -0.1824 0.05716 0.05217 -0.0438 0.9992 0.0323
-2.250 -0.1382 0.05684 0.05143 -0.0476 0.9964 0.0329
-2.000 -0.1102 0.04944 0.04406 -0.0521 0.9933 0.0346
-1.750 -0.0806 0.04605 0.04055 -0.0550 0.9902 0.0378
-1.500 -0.0423 0.04435 0.03855 -0.0578 0.9870 0.0438
-1.250 0.0005 0.04091 0.03465 -0.0620 0.9832 0.0514
-1.000 0.0383 0.03903 0.03243 -0.0645 0.9795 0.0621
-0.750 0.0695 0.03691 0.03008 -0.0663 0.9760 0.0785
-0.500 0.1064 0.03483 0.02766 -0.0690 0.9720 0.1142
-0.250 0.1440 0.03202 0.02464 -0.0730 0.9684 0.2020
0.000 0.1703 0.02995 0.02245 -0.0734 0.9642 0.2450
0.250 0.2064 0.02869 0.02098 -0.0746 0.9596 0.2658
0.500 0.2499 0.02792 0.01997 -0.0766 0.9547 0.2734
0.750 0.2818 0.02849 0.02004 -0.0754 0.9485 0.2049
1.000 0.3275 0.02911 0.02025 -0.0760 0.9431 0.1182
1.250 0.3540 0.02918 0.02006 -0.0746 0.9359 0.0836
1.500 0.3979 0.02871 0.01954 -0.0770 0.9300 0.0634
1.750 0.4212 0.02854 0.01932 -0.0762 0.9217 0.0558
2.000 0.4604 0.02862 0.01918 -0.0785 0.9145 0.0484
2.250 0.4945 0.02882 0.01913 -0.0801 0.9058 0.0462
2.500 0.5241 0.02911 0.01920 -0.0809 0.8965 0.0474
2.750 0.5835 0.02774 0.01956 -0.0876 0.8906 1.0000
3.000 0.6095 0.02820 0.01987 -0.0876 0.8793 1.0000
3.250 0.6392 0.02866 0.02022 -0.0883 0.8676 1.0000
3.500 0.6703 0.02909 0.02056 -0.0891 0.8560 1.0000
3.750 0.7026 0.02945 0.02092 -0.0900 0.8443 1.0000
4.000 0.7364 0.02973 0.02126 -0.0910 0.8326 1.0000
4.250 0.8790 0.01717 0.00912 -0.0909 0.6784 1.0000
4.500 0.8857 0.01725 0.00648 -0.0825 0.2542 1.0000
4.750 0.8893 0.02007 0.00780 -0.0790 0.0244 1.0000
5.000 0.9097 0.02109 0.00911 -0.0771 0.0214 1.0000
5.250 0.9285 0.02230 0.01066 -0.0751 0.0213 1.0000
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