NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il) Reynolds number: 100,000 Max Cl/Cd: 49.44 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-nlf0215f-il-100000-n5.txt Download as CSV file: xf-nlf0215f-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY NLF(1)-0215F 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 -0.2799 0.07273 0.06788 -0.0907 0.9397 0.0330
-8.500 -0.2775 0.06818 0.06328 -0.0946 0.9266 0.0325
-8.250 -0.2803 0.06339 0.05839 -0.0987 0.9123 0.0319
-8.000 -0.2788 0.05802 0.05282 -0.1038 0.8985 0.0313
-7.750 -0.2722 0.05259 0.04706 -0.1084 0.8875 0.0306
-7.500 -0.2610 0.04756 0.04161 -0.1118 0.8771 0.0298
-7.250 -0.2488 0.04301 0.03658 -0.1138 0.8659 0.0291
-7.000 -0.2260 0.03872 0.03170 -0.1165 0.8587 0.0286
-6.750 -0.2052 0.03563 0.02813 -0.1174 0.8485 0.0285
-6.500 -0.1759 0.03284 0.02485 -0.1192 0.8415 0.0285
-6.250 -0.1499 0.03074 0.02238 -0.1199 0.8325 0.0288
-6.000 -0.1177 0.02898 0.02029 -0.1216 0.8261 0.0297
-5.750 -0.0913 0.02760 0.01864 -0.1218 0.8174 0.0311
-5.500 -0.0584 0.02613 0.01682 -0.1229 0.8117 0.0326
-5.250 -0.0325 0.02497 0.01542 -0.1226 0.8039 0.0332
-5.000 -0.0033 0.02377 0.01410 -0.1228 0.7977 0.0339
-4.750 0.0239 0.02282 0.01312 -0.1229 0.7914 0.0350
-4.500 0.0494 0.02207 0.01233 -0.1226 0.7844 0.0364
-4.250 0.0787 0.02133 0.01149 -0.1230 0.7793 0.0383
-4.000 0.1022 0.02084 0.01092 -0.1224 0.7727 0.0414
-3.750 0.1282 0.02030 0.01039 -0.1224 0.7672 0.0458
-3.500 0.1576 0.01979 0.00974 -0.1229 0.7628 0.0509
-3.250 0.1805 0.01939 0.00933 -0.1222 0.7567 0.0597
-3.000 0.2054 0.01877 0.00896 -0.1220 0.7514 0.1007
-2.750 0.2298 0.01687 0.00875 -0.1236 0.7472 0.4575
-2.500 0.2377 0.01849 0.01094 -0.1164 0.7421 0.6678
-2.000 0.2692 0.02063 0.01292 -0.1074 0.7331 0.7451
-1.750 0.2856 0.02113 0.01330 -0.1033 0.7296 0.7658
-1.500 0.3008 0.02133 0.01344 -0.1002 0.7242 0.7785
-1.250 0.3259 0.02135 0.01332 -0.1000 0.7200 0.7861
-1.000 0.3532 0.02130 0.01311 -0.1001 0.7165 0.7904
-0.750 0.3814 0.02125 0.01290 -0.1002 0.7134 0.7938
-0.500 0.4023 0.02130 0.01290 -0.0994 0.7084 0.7976
-0.250 0.4296 0.02133 0.01282 -0.0999 0.7043 0.8011
0.000 0.4594 0.02133 0.01271 -0.1009 0.7009 0.8038
0.250 0.4888 0.02128 0.01255 -0.1013 0.6981 0.8057
0.500 0.5098 0.02138 0.01264 -0.1005 0.6936 0.8080
0.750 0.5340 0.02147 0.01269 -0.1003 0.6893 0.8102
1.000 0.5624 0.02153 0.01269 -0.1009 0.6859 0.8122
1.250 0.5938 0.02157 0.01265 -0.1020 0.6829 0.8141
1.500 0.6217 0.02170 0.01275 -0.1026 0.6793 0.8160
1.750 0.6460 0.02193 0.01299 -0.1028 0.6748 0.8175
2.000 0.6728 0.02205 0.01310 -0.1030 0.6710 0.8190
2.250 0.7014 0.02211 0.01314 -0.1035 0.6678 0.8203
2.750 0.7507 0.02250 0.01359 -0.1034 0.6599 0.8225
3.000 0.7756 0.02270 0.01381 -0.1033 0.6556 0.8240
3.250 0.8053 0.02281 0.01393 -0.1040 0.6522 0.8254
3.500 0.8387 0.02286 0.01397 -0.1054 0.6493 0.8263
3.750 0.8555 0.02330 0.01452 -0.1041 0.6433 0.8274
4.000 0.8819 0.02350 0.01477 -0.1044 0.6386 0.8285
4.250 0.9145 0.02355 0.01484 -0.1056 0.6348 0.8298
4.500 0.9383 0.02383 0.01520 -0.1054 0.6294 0.8312
4.750 0.9617 0.02410 0.01554 -0.1052 0.6232 0.8322
5.000 0.9959 0.02405 0.01553 -0.1065 0.6186 0.8330
5.250 1.0151 0.02434 0.01593 -0.1055 0.6116 0.8338
5.500 1.0412 0.02437 0.01603 -0.1053 0.6049 0.8346
5.750 1.0702 0.02432 0.01602 -0.1057 0.5989 0.8356
6.000 1.0874 0.02457 0.01643 -0.1041 0.5905 0.8370
6.250 1.1201 0.02435 0.01623 -0.1049 0.5840 0.8382
6.500 1.1343 0.02465 0.01669 -0.1030 0.5742 0.8395
6.750 1.1600 0.02461 0.01674 -0.1027 0.5657 0.8405
7.000 1.1810 0.02467 0.01690 -0.1017 0.5559 0.8415
7.250 1.1965 0.02492 0.01729 -0.1000 0.5453 0.8427
7.500 1.2184 0.02492 0.01738 -0.0991 0.5347 0.8438
7.750 1.2356 0.02499 0.01756 -0.0974 0.5227 0.8450
8.000 1.2449 0.02539 0.01808 -0.0947 0.5090 0.8465
8.250 1.2574 0.02575 0.01854 -0.0926 0.4938 0.8481
8.500 1.2687 0.02609 0.01897 -0.0902 0.4772 0.8497
8.750 1.2770 0.02665 0.01960 -0.0876 0.4579 0.8513
9.000 1.2875 0.02719 0.02017 -0.0853 0.4349 0.8528
9.250 1.2974 0.02786 0.02079 -0.0831 0.4086 0.8542
9.500 1.3046 0.02878 0.02161 -0.0808 0.3798 0.8556
9.750 1.3082 0.03007 0.02276 -0.0783 0.3505 0.8571
10.000 1.3093 0.03168 0.02424 -0.0759 0.3232 0.8587
10.250 1.3094 0.03354 0.02602 -0.0738 0.2972 0.8602
10.500 1.3093 0.03557 0.02796 -0.0719 0.2738 0.8617
10.750 1.3095 0.03766 0.03000 -0.0703 0.2524 0.8631
11.000 1.3093 0.03983 0.03211 -0.0687 0.2334 0.8646
11.250 1.3098 0.04206 0.03430 -0.0674 0.2158 0.8663
11.500 1.3116 0.04426 0.03650 -0.0662 0.1999 0.8682
11.750 1.3137 0.04654 0.03876 -0.0653 0.1852 0.8702
12.000 1.3159 0.04888 0.04111 -0.0645 0.1720 0.8722
12.250 1.3194 0.05118 0.04344 -0.0638 0.1595 0.8744
12.500 1.3239 0.05348 0.04578 -0.0633 0.1485 0.8764
12.750 1.3273 0.05583 0.04817 -0.0628 0.1389 0.8783
13.000 1.3296 0.05834 0.05070 -0.0623 0.1303 0.8803
13.250 1.3350 0.06063 0.05309 -0.0619 0.1218 0.8825
13.500 1.3374 0.06328 0.05573 -0.0616 0.1152 0.8848
13.750 1.3433 0.06563 0.05824 -0.0614 0.1080 0.8873
14.000 1.3453 0.06842 0.06103 -0.0614 0.1026 0.8899
14.250 1.3516 0.07077 0.06355 -0.0612 0.0967 0.8928
14.500 1.3543 0.07351 0.06636 -0.0612 0.0919 0.8960
14.750 1.3583 0.07613 0.06907 -0.0611 0.0875 0.8998
15.000 1.3625 0.07880 0.07190 -0.0612 0.0828 0.9042
15.250 1.3642 0.08169 0.07485 -0.0613 0.0792 0.9097
15.500 1.3675 0.08436 0.07766 -0.0612 0.0757 0.9175
16.000 1.3699 0.09013 0.08374 -0.0618 0.0689 1.0000
16.250 1.3732 0.09328 0.08688 -0.0626 0.0661 1.0000
16.500 1.3763 0.09670 0.09056 -0.0636 0.0629 1.0000
16.750 1.3777 0.10035 0.09436 -0.0648 0.0599 1.0000
17.000 1.3781 0.10410 0.09817 -0.0662 0.0573 1.0000
17.250 1.3792 0.10776 0.10192 -0.0675 0.0549 1.0000
17.500 1.3774 0.11209 0.10650 -0.0693 0.0522 1.0000
17.750 1.3754 0.11646 0.11101 -0.0713 0.0498 1.0000
18.000 1.3737 0.12072 0.11537 -0.0735 0.0477 1.0000
18.250 1.3714 0.12514 0.11988 -0.0757 0.0457 1.0000
18.500 1.3656 0.13046 0.12546 -0.0785 0.0437 1.0000
18.750 1.3606 0.13563 0.13081 -0.0814 0.0418 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)