NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il) Reynolds number: 1,000,000 Max Cl/Cd: 118.36 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-nlf415-il-1000000.txt Download as CSV file: xf-nlf415-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY NLF(2)-0415 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.2150 0.10283 0.10033 -0.1389 0.8869 0.0105
-13.500 -0.2138 0.09887 0.09638 -0.1406 0.8845 0.0107
-13.250 -0.2131 0.09488 0.09238 -0.1425 0.8822 0.0111
-13.000 -0.2221 0.08809 0.08560 -0.1455 0.8800 0.0106
-12.750 -0.2250 0.08311 0.08061 -0.1481 0.8778 0.0109
-12.500 -0.2437 0.07317 0.07064 -0.1544 0.8752 0.0110
-12.250 -0.2629 0.06673 0.06413 -0.1580 0.8731 0.0113
-12.000 -0.2956 0.05983 0.05713 -0.1602 0.8708 0.0108
-11.750 -0.3180 0.05531 0.05251 -0.1606 0.8686 0.0108
-11.500 -0.3301 0.05256 0.04969 -0.1605 0.8666 0.0111
-11.250 -0.3470 0.04936 0.04637 -0.1597 0.8645 0.0112
-11.000 -0.3604 0.04670 0.04360 -0.1584 0.8623 0.0115
-10.750 -0.3768 0.04382 0.04058 -0.1564 0.8599 0.0114
-10.500 -0.3853 0.04167 0.03831 -0.1541 0.8583 0.0121
-10.250 -0.3886 0.04004 0.03644 -0.1504 0.8569 0.0131
-10.000 -0.4005 0.03874 0.03501 -0.1460 0.8553 0.0131
-9.750 -0.4060 0.03678 0.03287 -0.1428 0.8536 0.0132
-9.500 -0.4076 0.03484 0.03072 -0.1399 0.8521 0.0132
-9.250 -0.4043 0.03286 0.02855 -0.1375 0.8509 0.0132
-9.000 -0.4003 0.01347 0.00910 -0.1282 0.8485 0.0140
-8.750 -0.3845 0.01238 0.00797 -0.1276 0.8477 0.0143
-8.500 -0.3695 0.01139 0.00689 -0.1267 0.8469 0.0147
-8.250 -0.3533 0.01038 0.00577 -0.1258 0.8460 0.0156
-8.000 -0.3349 0.00935 0.00458 -0.1250 0.8452 0.0165
-7.750 -0.3207 0.00796 0.00286 -0.1234 0.8444 0.0185
-7.500 -0.2971 0.00693 0.00184 -0.1236 0.8441 0.0194
-7.250 -0.2584 0.01613 0.01044 -0.1313 0.8448 0.0141
-7.000 -0.2349 0.01544 0.00973 -0.1310 0.8444 0.0144
-6.750 -0.2131 0.01491 0.00921 -0.1304 0.8440 0.0149
-6.500 -0.1935 0.01417 0.00844 -0.1292 0.8435 0.0148
-6.250 -0.1758 0.01360 0.00786 -0.1279 0.8430 0.0151
-6.000 -0.1584 0.01309 0.00732 -0.1266 0.8424 0.0154
-5.750 -0.1394 0.01267 0.00689 -0.1255 0.8418 0.0159
-5.500 -0.1210 0.01220 0.00639 -0.1244 0.8412 0.0160
-5.250 -0.1005 0.01181 0.00598 -0.1236 0.8406 0.0164
-5.000 -0.0790 0.01146 0.00559 -0.1230 0.8399 0.0166
-4.750 -0.0568 0.01111 0.00521 -0.1225 0.8392 0.0168
-4.500 -0.0330 0.01084 0.00492 -0.1222 0.8387 0.0173
-4.250 -0.0084 0.01063 0.00468 -0.1220 0.8382 0.0177
-4.000 0.0156 0.01032 0.00433 -0.1218 0.8377 0.0184
-3.750 0.0408 0.01010 0.00407 -0.1217 0.8372 0.0196
-3.500 0.0668 0.00994 0.00391 -0.1218 0.8368 0.0214
-3.250 0.0933 0.00983 0.00378 -0.1219 0.8364 0.0238
-3.000 0.1095 0.00844 0.00325 -0.1213 0.8357 0.2881
-2.750 0.1292 0.00740 0.00308 -0.1209 0.8351 0.5649
-2.500 0.1578 0.00760 0.00326 -0.1213 0.8348 0.5786
-2.250 0.1864 0.00786 0.00352 -0.1215 0.8345 0.5887
-2.000 0.2147 0.00794 0.00358 -0.1219 0.8341 0.5928
-1.750 0.2429 0.00801 0.00367 -0.1223 0.8338 0.5956
-1.500 0.2714 0.00814 0.00382 -0.1226 0.8334 0.5996
-1.250 0.3001 0.00836 0.00402 -0.1230 0.8329 0.6046
-1.000 0.3285 0.00845 0.00409 -0.1234 0.8326 0.6055
-0.750 0.3566 0.00844 0.00408 -0.1238 0.8323 0.6062
-0.500 0.3847 0.00846 0.00410 -0.1242 0.8319 0.6068
-0.250 0.4127 0.00850 0.00415 -0.1246 0.8316 0.6074
0.000 0.4405 0.00856 0.00423 -0.1250 0.8313 0.6080
0.250 0.4682 0.00864 0.00434 -0.1253 0.8309 0.6085
0.500 0.4956 0.00877 0.00448 -0.1255 0.8297 0.6091
0.750 0.5232 0.00844 0.00418 -0.1255 0.8255 0.6097
1.000 0.5547 0.00806 0.00379 -0.1262 0.8201 0.6102
1.250 0.5886 0.00788 0.00357 -0.1276 0.8165 0.6109
1.500 0.6150 0.00780 0.00353 -0.1274 0.8131 0.6116
1.750 0.6414 0.00764 0.00340 -0.1272 0.8088 0.6124
2.000 0.6715 0.00745 0.00322 -0.1278 0.8047 0.6131
2.250 0.7001 0.00730 0.00306 -0.1280 0.7993 0.6138
2.500 0.7246 0.00711 0.00291 -0.1273 0.7924 0.6145
2.750 0.7512 0.00695 0.00275 -0.1271 0.7850 0.6153
3.000 0.7730 0.00680 0.00263 -0.1258 0.7745 0.6160
3.250 0.7942 0.00671 0.00258 -0.1245 0.7604 0.6167
3.500 0.7966 0.00678 0.00240 -0.1187 0.6884 0.6175
3.750 0.7538 0.00780 0.00287 -0.1036 0.5804 0.6181
4.000 0.7211 0.00918 0.00373 -0.0914 0.4688 0.6188
4.250 0.7065 0.01037 0.00448 -0.0835 0.3703 0.6194
4.500 0.7032 0.01141 0.00514 -0.0780 0.2831 0.6199
4.750 0.7091 0.01224 0.00566 -0.0744 0.2167 0.6205
5.000 0.7212 0.01286 0.00608 -0.0720 0.1700 0.6213
5.250 0.7370 0.01333 0.00643 -0.0702 0.1393 0.6224
5.750 0.7726 0.01418 0.00712 -0.0674 0.0959 0.6244
6.000 0.7918 0.01456 0.00746 -0.0663 0.0833 0.6254
6.250 0.8109 0.01496 0.00782 -0.0651 0.0713 0.6263
6.500 0.8305 0.01535 0.00817 -0.0641 0.0620 0.6273
6.750 0.8510 0.01570 0.00852 -0.0632 0.0552 0.6284
7.000 0.8702 0.01614 0.00893 -0.0621 0.0482 0.6297
7.250 0.8914 0.01644 0.00927 -0.0613 0.0455 0.6308
7.500 0.9111 0.01686 0.00966 -0.0603 0.0408 0.6319
7.750 0.9311 0.01726 0.01008 -0.0594 0.0374 0.6330
8.000 0.9522 0.01760 0.01045 -0.0587 0.0356 0.6340
8.250 0.9723 0.01801 0.01084 -0.0578 0.0320 0.6350
8.500 0.9920 0.01845 0.01130 -0.0569 0.0292 0.6359
8.750 1.0128 0.01877 0.01167 -0.0561 0.0270 0.6376
9.000 1.0322 0.01923 0.01212 -0.0552 0.0220 0.6392
9.250 1.0511 0.01974 0.01260 -0.0542 0.0163 0.6406
9.500 1.0689 0.02039 0.01324 -0.0531 0.0120 0.6421
9.750 1.0878 0.02096 0.01384 -0.0520 0.0106 0.6437
10.000 1.1053 0.02168 0.01461 -0.0508 0.0095 0.6452
10.250 1.1240 0.02225 0.01525 -0.0498 0.0091 0.6469
10.500 1.1422 0.02289 0.01596 -0.0488 0.0086 0.6485
10.750 1.1600 0.02357 0.01670 -0.0477 0.0083 0.6500
11.000 1.1776 0.02426 0.01745 -0.0467 0.0079 0.6518
11.250 1.1942 0.02505 0.01832 -0.0455 0.0075 0.6541
11.500 1.2083 0.02610 0.01947 -0.0440 0.0071 0.6562
11.750 1.2206 0.02731 0.02080 -0.0423 0.0069 0.6584
12.000 1.2360 0.02816 0.02175 -0.0411 0.0068 0.6608
12.250 1.2497 0.02921 0.02290 -0.0396 0.0067 0.6632
12.500 1.2639 0.03019 0.02397 -0.0383 0.0065 0.6656
12.750 1.2758 0.03135 0.02527 -0.0368 0.0064 0.6688
13.000 1.2877 0.03250 0.02653 -0.0353 0.0064 0.6721
13.250 1.2981 0.03382 0.02798 -0.0337 0.0063 0.6756
13.500 1.3074 0.03523 0.02951 -0.0321 0.0062 0.6790
13.750 1.3151 0.03676 0.03118 -0.0304 0.0061 0.6827
14.000 1.3227 0.03831 0.03288 -0.0287 0.0060 0.6870
14.250 1.3285 0.04000 0.03472 -0.0270 0.0059 0.6918
14.500 1.3337 0.04180 0.03666 -0.0254 0.0058 0.6967
14.750 1.3361 0.04385 0.03887 -0.0236 0.0058 0.7027
15.000 1.3386 0.04593 0.04110 -0.0220 0.0057 0.7094
15.250 1.3373 0.04848 0.04383 -0.0203 0.0057 0.7168
15.750 1.3353 0.05356 0.04923 -0.0176 0.0055 0.7376
16.000 1.3266 0.05717 0.05307 -0.0163 0.0055 0.7502
16.250 1.3226 0.06031 0.05640 -0.0156 0.0054 0.7691
16.500 1.3135 0.06423 0.06056 -0.0150 0.0054 0.7971
16.750 1.3012 0.06801 0.06488 -0.0142 0.0054 0.9510
17.000 1.2870 0.07318 0.07022 -0.0149 0.0054 1.0000
17.250 1.2739 0.07861 0.07579 -0.0162 0.0053 1.0000
17.500 1.2550 0.08545 0.08282 -0.0185 0.0054 1.0000
17.750 1.2337 0.09335 0.09090 -0.0222 0.0054 1.0000
18.000 1.2131 0.10191 0.09963 -0.0269 0.0053 1.0000
18.250 1.1812 0.11402 0.11197 -0.0344 0.0054 1.0000
18.500 1.1338 0.13117 0.12941 -0.0451 0.0056 1.0000
18.750 1.0817 0.14986 0.14832 -0.0563 0.0057 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il)