USA 45 AIRFOIL (usa45-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 45 AIRFOIL (usa45-il) Reynolds number: 500,000 Max Cl/Cd: 90.05 at α=9.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa45-il-500000.txt Download as CSV file: xf-usa45-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: USA 45 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3300 0.08899 0.08687 -0.0377 1.0000 0.0333
-9.750 -0.3345 0.08495 0.08286 -0.0393 1.0000 0.0346
-9.500 -0.6005 0.03568 0.03225 -0.0599 0.9849 0.0263
-9.250 -0.5113 0.04696 0.04439 -0.0616 0.9876 0.0315
-9.000 -0.4907 0.04338 0.04073 -0.0641 0.9819 0.0292
-8.750 -0.5095 0.03203 0.02845 -0.0654 0.9693 0.0285
-8.500 -0.4939 0.02820 0.02410 -0.0655 0.9581 0.0295
-8.250 -0.4705 0.02600 0.02143 -0.0657 0.9473 0.0305
-8.000 -0.4484 0.02304 0.01803 -0.0659 0.9353 0.0312
-7.750 -0.4216 0.02121 0.01604 -0.0663 0.9209 0.0319
-7.500 -0.3941 0.02053 0.01531 -0.0664 0.9030 0.0328
-7.250 -0.3690 0.01978 0.01443 -0.0659 0.8815 0.0337
-7.000 -0.3454 0.01894 0.01337 -0.0650 0.8596 0.0346
-6.750 -0.3229 0.01805 0.01225 -0.0638 0.8363 0.0352
-6.500 -0.3002 0.01733 0.01129 -0.0626 0.8115 0.0358
-6.250 -0.2773 0.01693 0.01066 -0.0613 0.7848 0.0365
-6.000 -0.2553 0.01585 0.00935 -0.0601 0.7587 0.0373
-5.750 -0.2328 0.01502 0.00839 -0.0589 0.7313 0.0381
-5.500 -0.2100 0.01453 0.00777 -0.0578 0.7044 0.0388
-5.250 -0.1872 0.01416 0.00725 -0.0567 0.6765 0.0396
-5.000 -0.1644 0.01384 0.00678 -0.0555 0.6471 0.0405
-4.750 -0.1413 0.01358 0.00637 -0.0544 0.6179 0.0416
-4.500 -0.1183 0.01333 0.00596 -0.0533 0.5889 0.0425
-4.250 -0.0952 0.01311 0.00558 -0.0521 0.5617 0.0432
-4.000 -0.0717 0.01292 0.00524 -0.0511 0.5368 0.0437
-3.750 -0.0505 0.01244 0.00464 -0.0497 0.5163 0.0452
-3.500 -0.0272 0.01222 0.00434 -0.0488 0.4984 0.0466
-3.250 -0.0031 0.01210 0.00412 -0.0479 0.4831 0.0486
-3.000 0.0216 0.01197 0.00391 -0.0471 0.4699 0.0505
-2.750 0.0463 0.01182 0.00368 -0.0463 0.4592 0.0526
-2.500 0.0704 0.01166 0.00347 -0.0455 0.4501 0.0567
-2.250 0.0957 0.01152 0.00331 -0.0448 0.4418 0.0631
-2.000 0.1195 0.01133 0.00315 -0.0439 0.4345 0.0848
-1.750 0.1421 0.01090 0.00302 -0.0430 0.4279 0.1599
-1.500 0.1648 0.01062 0.00295 -0.0420 0.4214 0.2365
-1.250 0.1864 0.01024 0.00290 -0.0409 0.4157 0.3401
-1.000 0.2064 0.00976 0.00285 -0.0394 0.4101 0.4664
-0.750 0.2215 0.00922 0.00285 -0.0368 0.4053 0.6352
-0.500 0.2384 0.00888 0.00293 -0.0342 0.4008 0.7642
-0.250 0.2624 0.00875 0.00305 -0.0329 0.3962 0.8495
0.000 0.2932 0.00882 0.00317 -0.0331 0.3913 0.8953
0.250 0.3300 0.00905 0.00333 -0.0346 0.3860 0.9262
0.500 0.3742 0.00925 0.00351 -0.0376 0.3813 0.9479
0.750 0.4160 0.00947 0.00364 -0.0403 0.3761 0.9581
1.000 0.4586 0.00976 0.00381 -0.0432 0.3714 0.9655
1.250 0.5040 0.00997 0.00397 -0.0467 0.3674 0.9719
1.500 0.5488 0.01013 0.00407 -0.0501 0.3630 0.9768
1.750 0.5882 0.01031 0.00417 -0.0525 0.3590 0.9829
2.000 0.6333 0.01052 0.00427 -0.0562 0.3542 0.9864
2.250 0.6777 0.01061 0.00435 -0.0597 0.3508 0.9913
2.500 0.7172 0.01069 0.00440 -0.0622 0.3472 0.9947
2.750 0.7551 0.01075 0.00441 -0.0645 0.3438 0.9976
3.000 0.7916 0.01090 0.00447 -0.0665 0.3401 1.0000
3.250 0.8123 0.01099 0.00455 -0.0653 0.3373 1.0000
3.500 0.8332 0.01105 0.00462 -0.0641 0.3347 1.0000
3.750 0.8540 0.01113 0.00470 -0.0628 0.3320 1.0000
4.000 0.8746 0.01122 0.00479 -0.0615 0.3293 1.0000
4.250 0.8949 0.01136 0.00489 -0.0601 0.3266 1.0000
4.500 0.9149 0.01161 0.00507 -0.0588 0.3233 1.0000
4.750 0.9358 0.01169 0.00520 -0.0575 0.3213 1.0000
5.000 0.9566 0.01180 0.00533 -0.0562 0.3188 1.0000
5.250 0.9773 0.01192 0.00546 -0.0549 0.3162 1.0000
5.500 0.9980 0.01206 0.00560 -0.0535 0.3135 1.0000
5.750 1.0185 0.01225 0.00576 -0.0522 0.3107 1.0000
6.000 1.0391 0.01250 0.00598 -0.0509 0.3075 1.0000
6.250 1.0600 0.01258 0.00612 -0.0497 0.3047 1.0000
6.500 1.0809 0.01268 0.00626 -0.0484 0.3013 1.0000
6.750 1.1017 0.01281 0.00639 -0.0471 0.2980 1.0000
7.000 1.1218 0.01304 0.00657 -0.0458 0.2942 1.0000
7.250 1.1425 0.01316 0.00674 -0.0445 0.2903 1.0000
7.500 1.1634 0.01324 0.00687 -0.0433 0.2861 1.0000
7.750 1.1837 0.01340 0.00702 -0.0420 0.2822 1.0000
8.000 1.2030 0.01368 0.00726 -0.0406 0.2783 1.0000
8.250 1.2242 0.01376 0.00743 -0.0394 0.2744 1.0000
8.500 1.2447 0.01389 0.00760 -0.0382 0.2698 1.0000
8.750 1.2633 0.01412 0.00780 -0.0367 0.2651 1.0000
9.000 1.2832 0.01431 0.00805 -0.0354 0.2610 1.0000
9.250 1.3030 0.01447 0.00825 -0.0341 0.2554 1.0000
9.500 1.3193 0.01474 0.00850 -0.0322 0.2487 1.0000
9.750 1.3385 0.01492 0.00872 -0.0309 0.2407 1.0000
10.000 1.3521 0.01522 0.00902 -0.0285 0.2333 1.0000
10.250 1.3650 0.01555 0.00934 -0.0261 0.2225 1.0000
10.500 1.3751 0.01605 0.00977 -0.0234 0.2066 1.0000
10.750 1.3789 0.01690 0.01046 -0.0200 0.1813 1.0000
11.000 1.3734 0.01829 0.01164 -0.0156 0.1485 1.0000
11.250 1.3632 0.02009 0.01324 -0.0112 0.1190 1.0000
11.500 1.3492 0.02230 0.01529 -0.0071 0.0905 1.0000
11.750 1.3196 0.02591 0.01871 -0.0029 0.0480 1.0000
12.000 1.3006 0.02944 0.02223 -0.0009 0.0281 1.0000
12.250 1.2954 0.03221 0.02508 -0.0001 0.0240 1.0000
12.500 1.2960 0.03464 0.02762 0.0002 0.0229 1.0000
12.750 1.2923 0.03768 0.03077 0.0002 0.0216 1.0000
13.000 1.2873 0.04106 0.03427 -0.0001 0.0209 1.0000
13.250 1.2778 0.04509 0.03843 -0.0007 0.0202 1.0000
13.500 1.2666 0.04946 0.04294 -0.0015 0.0196 1.0000
13.750 1.2567 0.05377 0.04736 -0.0024 0.0194 1.0000
14.000 1.2452 0.05833 0.05205 -0.0034 0.0191 1.0000
14.250 1.2342 0.06286 0.05669 -0.0044 0.0186 1.0000
14.500 1.2206 0.06776 0.06172 -0.0056 0.0186 1.0000
14.750 1.2076 0.07269 0.06675 -0.0068 0.0182 1.0000
15.000 1.1948 0.07763 0.07181 -0.0080 0.0179 1.0000
15.250 1.1826 0.08259 0.07687 -0.0093 0.0179 1.0000
15.500 1.1707 0.08757 0.08195 -0.0107 0.0176 1.0000
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