PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: PROPFAN CRUISE MISSILE WING AIRFOIL (pfcm-il) Reynolds number: 200,000 Max Cl/Cd: 61.64 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-pfcm-il-200000.txt Download as CSV file: xf-pfcm-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: PROPFAN CRUISE MISSILE WING AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4476 0.11842 0.11498 -0.0083 1.0000 0.0342
-10.750 -0.4481 0.11471 0.11127 -0.0097 1.0000 0.0359
-7.500 -0.5657 0.06440 0.06086 -0.0367 1.0000 0.0419
-7.250 -0.5577 0.06078 0.05718 -0.0376 1.0000 0.0435
-7.000 -0.5489 0.05684 0.05311 -0.0388 1.0000 0.0455
-6.750 -0.5383 0.05280 0.04884 -0.0399 1.0000 0.0486
-6.250 -0.4965 0.03175 0.02770 -0.0382 1.0000 0.0548
-6.000 -0.4874 0.02903 0.02492 -0.0368 1.0000 0.0574
-5.750 -0.4760 0.02950 0.02465 -0.0339 1.0000 0.0651
-5.500 -0.4702 0.02321 0.01845 -0.0338 1.0000 0.0677
-5.250 -0.4577 0.02105 0.01628 -0.0324 1.0000 0.0718
-5.000 -0.4456 0.01876 0.01364 -0.0310 1.0000 0.0818
-4.750 -0.4320 0.01702 0.01173 -0.0298 1.0000 0.0950
-4.500 -0.4173 0.01523 0.00988 -0.0287 1.0000 0.1097
-4.250 -0.3703 0.02206 0.01464 -0.0284 1.0000 0.0393
-4.000 -0.3469 0.01957 0.01192 -0.0275 1.0000 0.0385
-3.750 -0.3233 0.01815 0.01030 -0.0266 1.0000 0.0388
-3.500 -0.2989 0.01622 0.00828 -0.0262 1.0000 0.0423
-3.250 -0.2748 0.01524 0.00728 -0.0256 1.0000 0.0444
-3.000 -0.2504 0.01441 0.00643 -0.0251 1.0000 0.0471
-2.750 -0.2252 0.01378 0.00574 -0.0248 1.0000 0.0513
-2.500 -0.1990 0.01304 0.00502 -0.0250 1.0000 0.0635
-2.250 -0.1649 0.01004 0.00421 -0.0278 1.0000 0.5743
-2.000 -0.1337 0.00971 0.00441 -0.0281 0.9964 0.7367
-1.750 -0.0987 0.00959 0.00441 -0.0292 0.9907 0.8014
-1.500 -0.0695 0.00942 0.00447 -0.0285 0.9846 0.8746
-1.250 -0.0335 0.00940 0.00452 -0.0289 0.9804 0.9589
-1.000 0.0204 0.00938 0.00439 -0.0348 0.9765 0.9954
-0.750 0.0672 0.00933 0.00423 -0.0396 0.9712 1.0000
-0.500 0.1093 0.00929 0.00410 -0.0432 0.9644 1.0000
-0.250 0.1525 0.00925 0.00398 -0.0469 0.9582 1.0000
0.000 0.1916 0.00922 0.00391 -0.0496 0.9509 1.0000
0.250 0.2301 0.00917 0.00384 -0.0520 0.9434 1.0000
0.500 0.2629 0.00916 0.00381 -0.0532 0.9335 1.0000
0.750 0.2958 0.00914 0.00378 -0.0542 0.9248 1.0000
1.000 0.3254 0.00913 0.00377 -0.0545 0.9147 1.0000
1.250 0.3522 0.00916 0.00383 -0.0542 0.9031 1.0000
1.500 0.3786 0.00920 0.00388 -0.0538 0.8918 1.0000
1.750 0.4047 0.00923 0.00394 -0.0532 0.8804 1.0000
2.000 0.4301 0.00925 0.00398 -0.0523 0.8679 1.0000
2.250 0.4545 0.00924 0.00403 -0.0511 0.8524 1.0000
2.500 0.4762 0.00908 0.00383 -0.0488 0.8274 1.0000
2.750 0.4973 0.00890 0.00357 -0.0462 0.7910 1.0000
3.000 0.5206 0.00889 0.00350 -0.0446 0.7553 1.0000
3.250 0.5441 0.00897 0.00353 -0.0431 0.7126 1.0000
3.500 0.5665 0.00919 0.00353 -0.0413 0.6437 1.0000
3.750 0.5836 0.01002 0.00361 -0.0388 0.4688 1.0000
4.250 0.6111 0.01459 0.00602 -0.0351 0.0551 1.0000
4.500 0.6323 0.01584 0.00722 -0.0341 0.0468 1.0000
4.750 0.6559 0.01680 0.00830 -0.0332 0.0436 1.0000
5.000 0.6795 0.01800 0.00955 -0.0323 0.0410 1.0000
5.250 0.7039 0.01934 0.01092 -0.0315 0.0388 1.0000
5.500 0.7277 0.02203 0.01362 -0.0309 0.0353 1.0000
5.750 0.7537 0.02392 0.01572 -0.0302 0.0347 1.0000
6.000 0.7791 0.02649 0.01857 -0.0294 0.0347 1.0000
6.250 0.8045 0.02912 0.02149 -0.0286 0.0358 1.0000
6.500 0.8286 0.03131 0.02402 -0.0274 0.0363 1.0000
6.750 0.8562 0.04054 0.03473 -0.0222 0.0758 1.0000
7.000 0.8680 0.04366 0.03830 -0.0206 0.0668 1.0000
7.250 0.8843 0.04661 0.04135 -0.0198 0.0629 1.0000
7.750 0.8989 0.05423 0.04971 -0.0167 0.0538 1.0000
8.000 0.9069 0.05765 0.05332 -0.0157 0.0512 1.0000
8.250 0.9163 0.06117 0.05690 -0.0150 0.0494 1.0000
8.500 0.9257 0.06820 0.06375 -0.0155 0.0477 1.0000
10.750 0.8050 0.11924 0.11592 -0.0393 0.0434 1.0000
11.000 0.8019 0.12482 0.12147 -0.0424 0.0421 1.0000
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