RAF 28 AIRFOIL (raf28-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 28 AIRFOIL (raf28-il) Reynolds number: 500,000 Max Cl/Cd: 90.59 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf28-il-500000.txt Download as CSV file: xf-raf28-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5722 0.07200 0.06974 -0.0381 1.0000 0.0193
-9.750 -0.6094 0.05746 0.05516 -0.0515 1.0000 0.0186
-9.500 -0.6404 0.05261 0.05021 -0.0522 1.0000 0.0182
-9.250 -0.6706 0.04688 0.04427 -0.0504 1.0000 0.0180
-9.000 -0.6966 0.04075 0.03777 -0.0475 1.0000 0.0178
-8.750 -0.7114 0.03559 0.03219 -0.0440 1.0000 0.0179
-8.500 -0.7164 0.03152 0.02769 -0.0405 1.0000 0.0182
-8.250 -0.7117 0.02888 0.02465 -0.0375 1.0000 0.0190
-8.000 -0.6964 0.02864 0.02416 -0.0353 1.0000 0.0196
-7.750 -0.6972 0.02435 0.01954 -0.0320 1.0000 0.0205
-7.500 -0.6825 0.02347 0.01861 -0.0302 0.9998 0.0212
-7.250 -0.6486 0.02233 0.01736 -0.0322 0.9971 0.0223
-7.000 -0.6144 0.02100 0.01581 -0.0341 0.9944 0.0237
-6.750 -0.5810 0.02011 0.01468 -0.0355 0.9906 0.0251
-6.500 -0.5499 0.01791 0.01218 -0.0368 0.9868 0.0268
-6.250 -0.5157 0.01663 0.01084 -0.0387 0.9841 0.0284
-6.000 -0.4840 0.01578 0.00991 -0.0397 0.9789 0.0299
-5.750 -0.4499 0.01501 0.00906 -0.0412 0.9745 0.0317
-5.500 -0.4135 0.01450 0.00846 -0.0430 0.9715 0.0334
-5.250 -0.3829 0.01367 0.00754 -0.0437 0.9647 0.0343
-5.000 -0.3517 0.01228 0.00606 -0.0447 0.9597 0.0365
-4.750 -0.3181 0.01164 0.00540 -0.0460 0.9540 0.0384
-4.500 -0.2853 0.01109 0.00481 -0.0471 0.9462 0.0400
-4.250 -0.2513 0.01061 0.00427 -0.0485 0.9384 0.0417
-4.000 -0.2164 0.01021 0.00381 -0.0500 0.9299 0.0434
-3.750 -0.1840 0.00979 0.00332 -0.0509 0.9187 0.0450
-3.500 -0.1524 0.00942 0.00289 -0.0516 0.9070 0.0485
-3.250 -0.1222 0.00916 0.00259 -0.0520 0.8950 0.0538
-3.000 -0.0945 0.00882 0.00234 -0.0520 0.8833 0.0826
-2.750 -0.0702 0.00827 0.00209 -0.0514 0.8716 0.1683
-2.500 -0.0465 0.00783 0.00197 -0.0507 0.8604 0.2688
-2.250 -0.0216 0.00757 0.00188 -0.0501 0.8505 0.3326
-2.000 0.0034 0.00735 0.00180 -0.0496 0.8408 0.3899
-1.750 0.0274 0.00712 0.00174 -0.0488 0.8306 0.4528
-1.500 0.0508 0.00686 0.00169 -0.0478 0.8211 0.5261
-1.250 0.0722 0.00653 0.00168 -0.0464 0.8117 0.6284
-1.000 0.0928 0.00624 0.00170 -0.0446 0.8015 0.7246
-0.750 0.1156 0.00605 0.00171 -0.0432 0.7919 0.7937
-0.500 0.1432 0.00588 0.00175 -0.0427 0.7822 0.8669
-0.250 0.1956 0.00592 0.00189 -0.0475 0.7720 0.9339
0.000 0.2391 0.00604 0.00197 -0.0506 0.7613 0.9578
0.250 0.2758 0.00616 0.00202 -0.0523 0.7498 0.9697
0.500 0.3177 0.00623 0.00203 -0.0552 0.7362 0.9754
0.750 0.3535 0.00631 0.00205 -0.0569 0.7195 0.9833
1.000 0.3944 0.00636 0.00203 -0.0597 0.7028 0.9879
1.250 0.4326 0.00645 0.00203 -0.0620 0.6837 0.9936
1.500 0.4726 0.00651 0.00198 -0.0646 0.6615 0.9982
1.750 0.5023 0.00659 0.00197 -0.0651 0.6420 1.0000
2.000 0.5244 0.00667 0.00200 -0.0640 0.6263 1.0000
2.250 0.5467 0.00677 0.00203 -0.0628 0.6102 1.0000
2.500 0.5691 0.00687 0.00207 -0.0617 0.5934 1.0000
2.750 0.5913 0.00699 0.00212 -0.0605 0.5748 1.0000
3.000 0.6140 0.00709 0.00219 -0.0594 0.5566 1.0000
3.250 0.6364 0.00722 0.00227 -0.0582 0.5362 1.0000
3.500 0.6587 0.00736 0.00236 -0.0571 0.5151 1.0000
3.750 0.6807 0.00753 0.00246 -0.0558 0.4899 1.0000
4.000 0.7021 0.00775 0.00259 -0.0545 0.4607 1.0000
4.250 0.7229 0.00802 0.00274 -0.0531 0.4268 1.0000
4.500 0.7425 0.00839 0.00293 -0.0515 0.3803 1.0000
4.750 0.7602 0.00891 0.00318 -0.0496 0.3160 1.0000
5.000 0.7748 0.00972 0.00358 -0.0473 0.2297 1.0000
5.250 0.7884 0.01066 0.00407 -0.0449 0.1452 1.0000
5.500 0.8039 0.01146 0.00456 -0.0428 0.0876 1.0000
5.750 0.8187 0.01235 0.00515 -0.0404 0.0425 1.0000
6.000 0.8387 0.01281 0.00566 -0.0389 0.0371 1.0000
6.250 0.8555 0.01353 0.00642 -0.0368 0.0319 1.0000
6.500 0.8753 0.01399 0.00693 -0.0353 0.0297 1.0000
6.750 0.8942 0.01451 0.00751 -0.0336 0.0275 1.0000
7.000 0.9111 0.01519 0.00821 -0.0316 0.0256 1.0000
7.250 0.9207 0.01646 0.00956 -0.0285 0.0236 1.0000
7.500 0.9414 0.01683 0.00999 -0.0272 0.0224 1.0000
7.750 0.9595 0.01741 0.01063 -0.0255 0.0212 1.0000
8.000 0.9763 0.01812 0.01137 -0.0236 0.0202 1.0000
8.250 0.9924 0.01888 0.01216 -0.0217 0.0192 1.0000
8.500 1.0058 0.01999 0.01330 -0.0195 0.0185 1.0000
8.750 1.0179 0.02178 0.01518 -0.0171 0.0176 1.0000
9.000 1.0361 0.02240 0.01590 -0.0156 0.0170 1.0000
9.250 1.0534 0.02310 0.01668 -0.0140 0.0161 1.0000
9.500 1.0697 0.02413 0.01780 -0.0123 0.0155 1.0000
9.750 1.0856 0.02516 0.01892 -0.0107 0.0150 1.0000
10.000 1.1000 0.02613 0.01995 -0.0089 0.0144 1.0000
10.250 1.1137 0.02746 0.02138 -0.0070 0.0141 1.0000
10.500 1.1276 0.02938 0.02338 -0.0056 0.0136 1.0000
10.750 1.1392 0.03324 0.02752 -0.0043 0.0132 1.0000
11.000 1.1424 0.03436 0.02885 -0.0008 0.0130 1.0000
11.250 1.1446 0.03658 0.03128 0.0022 0.0130 1.0000
11.500 1.1428 0.03822 0.03316 0.0056 0.0128 1.0000
11.750 1.1398 0.04027 0.03543 0.0086 0.0126 1.0000
12.000 1.1331 0.04302 0.03840 0.0113 0.0126 1.0000
12.250 1.1251 0.04531 0.04092 0.0136 0.0121 1.0000
12.500 1.1066 0.04961 0.04549 0.0154 0.0124 1.0000
12.750 1.0947 0.05284 0.04891 0.0163 0.0122 1.0000
13.000 1.0794 0.05679 0.05307 0.0164 0.0120 1.0000
13.250 1.0604 0.06172 0.05820 0.0157 0.0123 1.0000
13.500 1.0419 0.06692 0.06358 0.0140 0.0122 1.0000
13.750 1.0186 0.07336 0.07021 0.0111 0.0123 1.0000
14.000 1.0006 0.07950 0.07649 0.0079 0.0124 1.0000
14.250 0.9742 0.08781 0.08497 0.0029 0.0124 1.0000
14.500 0.9534 0.09585 0.09312 -0.0020 0.0127 1.0000
14.750 0.9250 0.10633 0.10372 -0.0085 0.0129 1.0000
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