Cambered plate C=10% T=5% R=1.3 (cp-100-050-gn) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: Cambered plate C=10% T=5% R=1.3 (cp-100-050-gn) Reynolds number: 500,000 Max Cl/Cd: 46.01 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-cp-100-050-gn-500000-n5.txt Download as CSV file: xf-cp-100-050-gn-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: Cambered plate C=10% T=5% R=1.3
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.0302 0.11617 0.11308 -0.1254 0.9591 0.0283
-10.250 -0.0199 0.11408 0.11100 -0.1261 0.9566 0.0268
-10.000 -0.0116 0.11170 0.10861 -0.1270 0.9531 0.0273
-9.500 0.0137 0.10630 0.10320 -0.1313 0.9493 0.0290
-9.250 0.0303 0.10409 0.10099 -0.1338 0.9478 0.0310
-9.000 0.0457 0.10097 0.09786 -0.1370 0.9466 0.0313
-8.750 0.0626 0.09788 0.09475 -0.1406 0.9455 0.0313
-8.500 0.0809 0.09480 0.09165 -0.1443 0.9445 0.0313
-8.250 0.0868 0.09304 0.08990 -0.1437 0.9394 0.0299
-7.750 0.1170 0.08711 0.08394 -0.1500 0.9337 0.0314
-7.250 0.1579 0.08187 0.07867 -0.1571 0.9301 0.0307
-7.000 0.1686 0.07965 0.07644 -0.1583 0.9256 0.0310
-6.750 0.1796 0.07758 0.07437 -0.1595 0.9204 0.0320
-6.500 0.1955 0.07477 0.07153 -0.1626 0.9167 0.0318
-6.250 0.2174 0.07295 0.06970 -0.1654 0.9141 0.0324
-6.000 0.2270 0.07155 0.06831 -0.1654 0.9087 0.0335
-5.750 0.2360 0.06962 0.06638 -0.1659 0.9030 0.0339
-5.500 0.2504 0.06739 0.06411 -0.1680 0.8987 0.0338
-5.250 0.2564 0.06562 0.06235 -0.1678 0.8931 0.0338
-5.000 0.2598 0.06403 0.06075 -0.1671 0.8864 0.0342
-4.500 0.2862 0.06009 0.05678 -0.1702 0.8769 0.0353
-4.250 0.2966 0.05884 0.05554 -0.1699 0.8714 0.0358
-4.000 0.3131 0.05729 0.05397 -0.1713 0.8668 0.0372
-3.750 0.3324 0.05509 0.05174 -0.1740 0.8630 0.0370
-3.500 0.3416 0.05339 0.05005 -0.1740 0.8572 0.0368
-3.250 0.3572 0.05150 0.04814 -0.1754 0.8520 0.0369
-3.000 0.3779 0.04940 0.04600 -0.1779 0.8479 0.0372
-2.750 0.4000 0.04657 0.04311 -0.1819 0.8434 0.0385
-2.500 0.4146 0.04569 0.04226 -0.1809 0.8376 0.0389
-2.250 0.4365 0.04438 0.04090 -0.1821 0.8307 0.0398
-2.000 0.4575 0.04253 0.03902 -0.1837 0.8245 0.0400
-1.750 0.4782 0.04062 0.03704 -0.1850 0.8145 0.0398
-1.500 0.5009 0.03877 0.03513 -0.1865 0.8061 0.0399
-1.250 0.5239 0.03704 0.03331 -0.1878 0.7950 0.0402
-1.000 0.5473 0.03526 0.03147 -0.1890 0.7853 0.0411
-0.750 0.5803 0.03224 0.02826 -0.1931 0.7738 0.0421
-0.500 0.5952 0.03192 0.02792 -0.1910 0.7601 0.0425
-0.250 0.6160 0.03064 0.02656 -0.1908 0.7450 0.0428
0.000 0.6368 0.02943 0.02525 -0.1905 0.7294 0.0431
0.250 0.6543 0.02822 0.02387 -0.1893 0.7065 0.0432
0.500 0.6624 0.02741 0.02286 -0.1857 0.6714 0.0435
0.750 0.6640 0.02684 0.02206 -0.1807 0.6329 0.0442
1.000 0.6596 0.02645 0.02135 -0.1744 0.5755 0.0446
1.250 0.6399 0.02700 0.02134 -0.1652 0.4648 0.0448
1.500 0.6096 0.02882 0.02222 -0.1546 0.2465 0.0447
1.750 0.6074 0.02934 0.02207 -0.1497 0.0527 0.0447
2.000 0.6369 0.02771 0.02024 -0.1506 0.0448 0.0452
2.250 0.6694 0.02562 0.01792 -0.1520 0.0422 0.0462
2.500 0.6893 0.02553 0.01782 -0.1508 0.0396 0.0466
2.750 0.7115 0.02518 0.01741 -0.1499 0.0383 0.0471
3.000 0.7367 0.02435 0.01646 -0.1495 0.0373 0.0476
3.250 0.7627 0.02328 0.01522 -0.1492 0.0367 0.0480
3.500 0.7884 0.02209 0.01379 -0.1486 0.0361 0.0485
3.750 0.8158 0.01995 0.01105 -0.1481 0.0353 0.0502
4.000 0.8351 0.01995 0.01101 -0.1464 0.0349 0.0506
4.250 0.8539 0.02000 0.01100 -0.1446 0.0343 0.0508
4.500 0.8720 0.02012 0.01106 -0.1428 0.0337 0.0511
4.750 0.8904 0.02025 0.01114 -0.1409 0.0334 0.0514
5.000 0.9092 0.02032 0.01114 -0.1391 0.0330 0.0518
5.250 0.9270 0.02052 0.01130 -0.1372 0.0328 0.0524
5.500 0.9446 0.02076 0.01150 -0.1352 0.0325 0.0531
5.750 0.9623 0.02098 0.01165 -0.1332 0.0322 0.0537
6.000 0.9791 0.02128 0.01191 -0.1312 0.0320 0.0540
6.250 0.9954 0.02164 0.01225 -0.1291 0.0318 0.0543
6.500 1.0114 0.02206 0.01264 -0.1269 0.0315 0.0546
6.750 1.0275 0.02250 0.01307 -0.1248 0.0312 0.0548
7.000 1.0426 0.02296 0.01355 -0.1226 0.0310 0.0550
7.250 1.0576 0.02348 0.01409 -0.1205 0.0308 0.0555
7.500 1.0727 0.02404 0.01469 -0.1183 0.0306 0.0559
7.750 1.0874 0.02465 0.01532 -0.1162 0.0304 0.0565
8.000 1.1019 0.02529 0.01597 -0.1141 0.0300 0.0572
8.250 1.1165 0.02593 0.01664 -0.1119 0.0299 0.0576
8.500 1.1308 0.02663 0.01735 -0.1098 0.0296 0.0581
8.750 1.1440 0.02743 0.01816 -0.1077 0.0293 0.0583
9.000 1.1569 0.02833 0.01908 -0.1055 0.0290 0.0590
9.250 1.1717 0.02911 0.01989 -0.1035 0.0288 0.0594
9.500 1.1875 0.02984 0.02065 -0.1017 0.0287 0.0599
9.750 1.2032 0.03057 0.02143 -0.1000 0.0286 0.0604
10.000 1.2193 0.03133 0.02225 -0.0982 0.0284 0.0614
10.250 1.2357 0.03210 0.02306 -0.0966 0.0282 0.0623
10.500 1.2527 0.03289 0.02390 -0.0951 0.0280 0.0637
10.750 1.2700 0.03366 0.02471 -0.0936 0.0278 0.0656
11.000 1.2880 0.03445 0.02556 -0.0922 0.0276 0.0685
11.250 1.3063 0.03525 0.02640 -0.0908 0.0274 0.0725
11.750 1.3682 0.03640 0.02886 -0.0936 0.0268 1.0000
12.000 1.3856 0.03726 0.02974 -0.0922 0.0266 1.0000
12.250 1.4024 0.03814 0.03063 -0.0907 0.0263 1.0000
12.500 1.4191 0.03905 0.03156 -0.0893 0.0261 1.0000
12.750 1.4345 0.03999 0.03253 -0.0878 0.0258 1.0000
13.000 1.4508 0.04096 0.03354 -0.0863 0.0257 1.0000
13.250 1.4675 0.04197 0.03458 -0.0850 0.0255 1.0000
13.500 1.4826 0.04302 0.03566 -0.0835 0.0253 1.0000
13.750 1.4979 0.04411 0.03677 -0.0821 0.0252 1.0000
14.000 1.5145 0.04525 0.03795 -0.0808 0.0250 1.0000
14.250 1.5325 0.04649 0.03923 -0.0797 0.0249 1.0000
14.500 1.5505 0.04795 0.04077 -0.0786 0.0247 1.0000
14.750 1.5607 0.04945 0.04241 -0.0768 0.0245 1.0000
15.000 1.5715 0.05109 0.04420 -0.0751 0.0244 1.0000
15.250 1.5807 0.05296 0.04625 -0.0733 0.0242 1.0000
15.500 1.5888 0.05509 0.04857 -0.0716 0.0240 1.0000
15.750 1.5944 0.05741 0.05109 -0.0697 0.0238 1.0000
16.000 1.5972 0.05988 0.05376 -0.0678 0.0234 1.0000
16.250 1.5978 0.06257 0.05664 -0.0659 0.0232 1.0000
16.500 1.5965 0.06527 0.05952 -0.0640 0.0229 1.0000
16.750 1.5933 0.06824 0.06267 -0.0622 0.0228 1.0000
17.000 1.5879 0.07151 0.06611 -0.0606 0.0226 1.0000
17.250 1.5812 0.07492 0.06970 -0.0591 0.0225 1.0000
17.500 1.5726 0.07860 0.07355 -0.0579 0.0223 1.0000
17.750 1.5630 0.08244 0.07756 -0.0569 0.0222 1.0000
18.000 1.5515 0.08666 0.08197 -0.0563 0.0221 1.0000
18.250 1.5396 0.09102 0.08649 -0.0560 0.0220 1.0000
18.500 1.5250 0.09594 0.09159 -0.0562 0.0219 1.0000
18.750 1.5103 0.10100 0.09681 -0.0568 0.0219 1.0000
19.000 1.4926 0.10680 0.10279 -0.0581 0.0218 1.0000
19.250 1.4734 0.11310 0.10927 -0.0601 0.0217 1.0000
19.500 1.4500 0.12051 0.11688 -0.0631 0.0217 1.0000
19.750 1.4245 0.12881 0.12537 -0.0672 0.0216 1.0000
20.000 1.3766 0.14281 0.13968 -0.0760 0.0216 1.0000
|
Polar data table (+)
Polar graphs
<< Back to Cambered plate C=10% T=5% R=1.3 (cp-100-050-gn)