RAF 28 AIRFOIL (raf28-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 28 AIRFOIL (raf28-il) Reynolds number: 500,000 Max Cl/Cd: 79.52 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf28-il-500000-n5.txt Download as CSV file: xf-raf28-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 28 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.9100 0.04350 0.04069 -0.0613 1.0000 0.0065
-12.500 -0.9348 0.03854 0.03551 -0.0630 1.0000 0.0065
-12.250 -0.9519 0.03586 0.03265 -0.0602 1.0000 0.0066
-12.000 -0.9615 0.03347 0.03005 -0.0571 1.0000 0.0067
-11.750 -0.9627 0.03131 0.02767 -0.0547 1.0000 0.0068
-11.500 -0.9611 0.02922 0.02535 -0.0522 1.0000 0.0069
-11.250 -0.9553 0.02743 0.02333 -0.0499 1.0000 0.0070
-11.000 -0.9462 0.02587 0.02154 -0.0478 1.0000 0.0073
-10.750 -0.9360 0.02438 0.01986 -0.0457 1.0000 0.0075
-10.500 -0.9229 0.02319 0.01855 -0.0439 1.0000 0.0079
-10.250 -0.9067 0.02236 0.01762 -0.0423 1.0000 0.0084
-10.000 -0.8905 0.02147 0.01661 -0.0407 1.0000 0.0087
-9.750 -0.8739 0.02060 0.01561 -0.0391 1.0000 0.0092
-9.500 -0.8571 0.01973 0.01457 -0.0375 1.0000 0.0097
-9.250 -0.8404 0.01884 0.01352 -0.0357 1.0000 0.0100
-9.000 -0.8233 0.01799 0.01254 -0.0340 1.0000 0.0106
-8.750 -0.7929 0.01727 0.01175 -0.0352 0.9980 0.0115
-8.500 -0.7588 0.01683 0.01127 -0.0369 0.9951 0.0124
-8.250 -0.7268 0.01636 0.01072 -0.0381 0.9916 0.0137
-8.000 -0.6945 0.01579 0.01005 -0.0394 0.9878 0.0147
-7.750 -0.6618 0.01552 0.00976 -0.0407 0.9837 0.0161
-7.500 -0.6313 0.01522 0.00941 -0.0415 0.9784 0.0174
-7.250 -0.6002 0.01489 0.00902 -0.0424 0.9736 0.0190
-7.000 -0.5709 0.01461 0.00866 -0.0428 0.9665 0.0203
-6.750 -0.5406 0.01421 0.00817 -0.0435 0.9597 0.0213
-6.500 -0.5098 0.01349 0.00741 -0.0446 0.9522 0.0230
-6.250 -0.4763 0.01311 0.00697 -0.0460 0.9437 0.0245
-6.000 -0.4403 0.01270 0.00650 -0.0479 0.9353 0.0262
-5.750 -0.4043 0.01229 0.00601 -0.0498 0.9238 0.0277
-5.500 -0.3677 0.01193 0.00556 -0.0518 0.9103 0.0288
-5.250 -0.3337 0.01162 0.00514 -0.0532 0.8947 0.0295
-5.000 -0.3037 0.01133 0.00474 -0.0537 0.8790 0.0300
-4.750 -0.2774 0.01088 0.00416 -0.0535 0.8648 0.0308
-4.500 -0.2518 0.01054 0.00370 -0.0530 0.8525 0.0317
-4.250 -0.2262 0.01026 0.00333 -0.0525 0.8419 0.0329
-4.000 -0.2004 0.01005 0.00304 -0.0521 0.8322 0.0343
-3.750 -0.1746 0.00986 0.00278 -0.0516 0.8229 0.0355
-3.500 -0.1486 0.00970 0.00255 -0.0512 0.8142 0.0369
-3.000 -0.0961 0.00943 0.00217 -0.0504 0.7980 0.0402
-2.750 -0.0700 0.00930 0.00201 -0.0500 0.7904 0.0463
-2.500 -0.0446 0.00905 0.00187 -0.0495 0.7819 0.0733
-2.250 -0.0187 0.00888 0.00174 -0.0491 0.7738 0.0960
-2.000 0.0069 0.00869 0.00162 -0.0487 0.7648 0.1239
-1.750 0.0313 0.00834 0.00150 -0.0481 0.7555 0.1924
-1.500 0.0562 0.00808 0.00142 -0.0476 0.7456 0.2541
-1.250 0.0814 0.00789 0.00136 -0.0471 0.7343 0.3030
-1.000 0.1065 0.00769 0.00130 -0.0465 0.7222 0.3506
-0.750 0.1310 0.00749 0.00125 -0.0459 0.7084 0.4087
-0.500 0.1545 0.00722 0.00121 -0.0450 0.6940 0.4901
-0.250 0.1764 0.00691 0.00119 -0.0438 0.6797 0.5849
0.000 0.1969 0.00661 0.00121 -0.0421 0.6656 0.6946
0.250 0.2176 0.00635 0.00125 -0.0404 0.6527 0.7899
0.500 0.2435 0.00622 0.00131 -0.0397 0.6405 0.8573
0.750 0.2817 0.00626 0.00140 -0.0417 0.6276 0.9070
1.000 0.3219 0.00639 0.00147 -0.0443 0.6099 0.9319
1.250 0.3586 0.00652 0.00153 -0.0463 0.5929 0.9479
1.500 0.3953 0.00663 0.00159 -0.0482 0.5794 0.9586
1.750 0.4285 0.00676 0.00166 -0.0493 0.5631 0.9672
2.000 0.4588 0.00691 0.00173 -0.0498 0.5434 0.9757
2.250 0.4922 0.00705 0.00181 -0.0511 0.5255 0.9818
2.500 0.5272 0.00721 0.00189 -0.0527 0.5011 0.9885
2.750 0.5662 0.00747 0.00199 -0.0554 0.4627 0.9959
3.000 0.6029 0.00770 0.00209 -0.0575 0.4299 1.0000
3.250 0.6250 0.00788 0.00220 -0.0564 0.4083 1.0000
3.500 0.6465 0.00813 0.00234 -0.0551 0.3783 1.0000
3.750 0.6677 0.00841 0.00250 -0.0539 0.3439 1.0000
4.000 0.6879 0.00878 0.00270 -0.0524 0.3009 1.0000
4.250 0.7071 0.00924 0.00295 -0.0509 0.2516 1.0000
4.500 0.7249 0.00983 0.00328 -0.0491 0.1941 1.0000
4.750 0.7429 0.01042 0.00362 -0.0474 0.1426 1.0000
5.000 0.7627 0.01086 0.00394 -0.0460 0.1155 1.0000
5.250 0.7831 0.01126 0.00424 -0.0446 0.0944 1.0000
5.500 0.8000 0.01195 0.00467 -0.0427 0.0458 1.0000
5.750 0.8201 0.01238 0.00506 -0.0413 0.0356 1.0000
6.000 0.8416 0.01269 0.00541 -0.0401 0.0315 1.0000
6.250 0.8624 0.01308 0.00580 -0.0388 0.0277 1.0000
6.500 0.8830 0.01346 0.00622 -0.0375 0.0243 1.0000
6.750 0.9044 0.01378 0.00660 -0.0363 0.0216 1.0000
7.000 0.9246 0.01420 0.00702 -0.0350 0.0186 1.0000
7.250 0.9439 0.01469 0.00754 -0.0334 0.0165 1.0000
7.500 0.9642 0.01509 0.00800 -0.0321 0.0151 1.0000
7.750 0.9841 0.01552 0.00844 -0.0308 0.0135 1.0000
8.000 1.0023 0.01608 0.00901 -0.0292 0.0122 1.0000
8.250 1.0190 0.01676 0.00978 -0.0273 0.0113 1.0000
8.500 1.0371 0.01730 0.01039 -0.0257 0.0108 1.0000
8.750 1.0546 0.01788 0.01104 -0.0240 0.0101 1.0000
9.000 1.0717 0.01847 0.01169 -0.0224 0.0095 1.0000
9.250 1.0883 0.01906 0.01233 -0.0207 0.0090 1.0000
9.500 1.1030 0.01977 0.01308 -0.0187 0.0083 1.0000
9.750 1.1134 0.02074 0.01413 -0.0161 0.0079 1.0000
10.000 1.1253 0.02145 0.01492 -0.0136 0.0077 1.0000
10.250 1.1354 0.02223 0.01583 -0.0109 0.0075 1.0000
10.500 1.1450 0.02309 0.01678 -0.0083 0.0071 1.0000
10.750 1.1542 0.02402 0.01781 -0.0058 0.0069 1.0000
11.000 1.1629 0.02502 0.01890 -0.0035 0.0067 1.0000
11.250 1.1710 0.02612 0.02010 -0.0013 0.0065 1.0000
11.500 1.1796 0.02719 0.02125 0.0007 0.0063 1.0000
11.750 1.1856 0.02854 0.02272 0.0027 0.0062 1.0000
12.000 1.1933 0.02977 0.02402 0.0043 0.0060 1.0000
12.250 1.1963 0.03148 0.02582 0.0061 0.0058 1.0000
12.500 1.1984 0.03335 0.02780 0.0075 0.0056 1.0000
12.750 1.1977 0.03565 0.03025 0.0090 0.0055 1.0000
13.000 1.1982 0.03793 0.03271 0.0101 0.0055 1.0000
13.250 1.2039 0.03969 0.03461 0.0107 0.0053 1.0000
13.500 1.2036 0.04223 0.03732 0.0114 0.0051 1.0000
13.750 1.2000 0.04523 0.04048 0.0117 0.0051 1.0000
14.000 1.2004 0.04787 0.04326 0.0116 0.0050 1.0000
14.250 1.1940 0.05153 0.04710 0.0112 0.0049 1.0000
14.500 1.1826 0.05599 0.05176 0.0104 0.0049 1.0000
14.750 1.1773 0.05985 0.05576 0.0092 0.0047 1.0000
15.000 1.1658 0.06474 0.06082 0.0076 0.0048 1.0000
15.250 1.1554 0.06976 0.06600 0.0055 0.0048 1.0000
15.500 1.1419 0.07562 0.07203 0.0029 0.0047 1.0000
15.750 1.1196 0.08329 0.07990 -0.0008 0.0047 1.0000
16.000 1.1069 0.08970 0.08645 -0.0042 0.0047 1.0000
16.250 1.0893 0.09738 0.09428 -0.0084 0.0046 1.0000
16.500 1.0676 0.10631 0.10337 -0.0133 0.0046 1.0000
16.750 1.0354 0.11819 0.11544 -0.0199 0.0048 1.0000
17.000 1.0090 0.12942 0.12682 -0.0260 0.0049 1.0000
17.250 0.9708 0.14465 0.14221 -0.0341 0.0050 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 28 AIRFOIL (raf28-il)