RAF-48 AIRFOIL (raf48-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAF-48 AIRFOIL (raf48-il) Reynolds number: 200,000 Max Cl/Cd: 68.54 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf48-il-200000.txt Download as CSV file: xf-raf48-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAF-48 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
------ -------- --------- --------- -------- -------- --------
-12.750 -0.6386 0.06675 0.06262 -0.0731 1.0000 0.0516
-12.500 -0.6693 0.05990 0.05566 -0.0772 1.0000 0.0512
-12.250 -0.7046 0.05436 0.04995 -0.0791 1.0000 0.0510
-12.000 -0.7354 0.05039 0.04580 -0.0780 1.0000 0.0508
-11.750 -0.7624 0.04750 0.04274 -0.0744 1.0000 0.0508
-11.500 -0.7906 0.04529 0.04036 -0.0687 1.0000 0.0509
-11.250 -0.8122 0.04305 0.03789 -0.0635 1.0000 0.0511
-11.000 -0.8325 0.04135 0.03599 -0.0576 1.0000 0.0514
-10.750 -0.8584 0.04052 0.03505 -0.0502 1.0000 0.0515
-10.500 -0.8462 0.03724 0.03146 -0.0508 0.9940 0.0527
-10.250 -0.8124 0.03575 0.03000 -0.0534 0.9876 0.0545
-10.000 -0.7808 0.03370 0.02772 -0.0559 0.9816 0.0562
-9.750 -0.7546 0.03152 0.02516 -0.0572 0.9725 0.0581
-9.500 -0.7264 0.02954 0.02271 -0.0586 0.9650 0.0599
-9.250 -0.6919 0.02766 0.02086 -0.0609 0.9592 0.0621
-9.000 -0.6537 0.02631 0.01938 -0.0636 0.9553 0.0644
-8.750 -0.6224 0.02514 0.01798 -0.0648 0.9472 0.0670
-8.500 -0.5865 0.02354 0.01619 -0.0669 0.9416 0.0697
-8.250 -0.5517 0.02249 0.01515 -0.0687 0.9347 0.0724
-8.000 -0.5191 0.02168 0.01422 -0.0698 0.9267 0.0760
-7.750 -0.4868 0.02064 0.01303 -0.0709 0.9195 0.0793
-7.500 -0.4588 0.01985 0.01227 -0.0711 0.9102 0.0828
-7.250 -0.4316 0.01932 0.01165 -0.0710 0.9017 0.0874
-7.000 -0.4069 0.01858 0.01089 -0.0705 0.8926 0.0922
-6.750 -0.3834 0.01817 0.01046 -0.0697 0.8841 0.0978
-6.500 -0.3603 0.01759 0.00984 -0.0687 0.8754 0.1046
-6.250 -0.3381 0.01729 0.00949 -0.0675 0.8665 0.1130
-6.000 -0.3161 0.01678 0.00901 -0.0663 0.8572 0.1230
-5.750 -0.2956 0.01636 0.00859 -0.0648 0.8475 0.1348
-5.500 -0.2727 0.01598 0.00820 -0.0636 0.8395 0.1489
-5.250 -0.2511 0.01569 0.00791 -0.0624 0.8315 0.1633
-5.000 -0.2281 0.01541 0.00762 -0.0614 0.8244 0.1783
-4.750 -0.2043 0.01515 0.00736 -0.0605 0.8182 0.1937
-4.500 -0.1823 0.01492 0.00716 -0.0593 0.8103 0.2095
-4.250 -0.1583 0.01462 0.00689 -0.0585 0.8043 0.2275
-4.000 -0.1363 0.01438 0.00674 -0.0573 0.7973 0.2470
-3.750 -0.1136 0.01411 0.00654 -0.0562 0.7905 0.2707
-3.250 -0.0701 0.01350 0.00618 -0.0537 0.7771 0.3361
-3.000 -0.0491 0.01311 0.00600 -0.0523 0.7709 0.3861
-2.750 -0.0301 0.01276 0.00591 -0.0504 0.7646 0.4531
-2.500 -0.0126 0.01239 0.00586 -0.0482 0.7573 0.5274
-2.250 0.0075 0.01205 0.00581 -0.0462 0.7517 0.6095
-2.000 0.0273 0.01189 0.00591 -0.0441 0.7442 0.6829
-1.750 0.0518 0.01178 0.00594 -0.0427 0.7372 0.7415
-1.500 0.0789 0.01177 0.00598 -0.0418 0.7312 0.7874
-1.250 0.1046 0.01181 0.00607 -0.0408 0.7232 0.8231
-1.000 0.1383 0.01184 0.00606 -0.0413 0.7170 0.8532
-0.750 0.1719 0.01192 0.00614 -0.0420 0.7090 0.8790
-0.500 0.2169 0.01202 0.00617 -0.0448 0.7013 0.8981
-0.250 0.2623 0.01220 0.00628 -0.0478 0.6933 0.9134
0.000 0.3073 0.01238 0.00638 -0.0508 0.6849 0.9267
0.250 0.3624 0.01266 0.00656 -0.0557 0.6766 0.9347
0.500 0.4072 0.01282 0.00664 -0.0587 0.6672 0.9454
0.750 0.4468 0.01296 0.00672 -0.0608 0.6581 0.9561
1.000 0.4945 0.01302 0.00668 -0.0647 0.6487 0.9629
1.250 0.5373 0.01312 0.00674 -0.0677 0.6384 0.9726
1.500 0.5873 0.01316 0.00666 -0.0721 0.6293 0.9830
1.750 0.6371 0.01308 0.00658 -0.0768 0.6177 0.9919
2.000 0.6821 0.01294 0.00636 -0.0805 0.6078 0.9987
2.250 0.7065 0.01290 0.00627 -0.0802 0.5983 1.0000
2.500 0.7250 0.01293 0.00628 -0.0787 0.5892 1.0000
2.750 0.7442 0.01294 0.00622 -0.0772 0.5804 1.0000
3.000 0.7622 0.01298 0.00626 -0.0756 0.5707 1.0000
3.250 0.7821 0.01301 0.00620 -0.0742 0.5624 1.0000
3.500 0.7996 0.01308 0.00629 -0.0724 0.5524 1.0000
3.750 0.8197 0.01315 0.00627 -0.0710 0.5445 1.0000
4.000 0.8372 0.01324 0.00640 -0.0692 0.5350 1.0000
4.250 0.8574 0.01337 0.00645 -0.0678 0.5274 1.0000
4.500 0.8753 0.01350 0.00661 -0.0660 0.5184 1.0000
4.750 0.8959 0.01366 0.00670 -0.0646 0.5108 1.0000
5.000 0.9144 0.01383 0.00691 -0.0629 0.5019 1.0000
5.250 0.9354 0.01402 0.00703 -0.0617 0.4944 1.0000
5.500 0.9542 0.01421 0.00726 -0.0600 0.4855 1.0000
5.750 0.9758 0.01444 0.00743 -0.0588 0.4781 1.0000
6.000 0.9946 0.01465 0.00771 -0.0572 0.4694 1.0000
6.250 1.0165 0.01491 0.00790 -0.0560 0.4618 1.0000
6.500 1.0347 0.01514 0.00821 -0.0543 0.4528 1.0000
6.750 1.0560 0.01542 0.00843 -0.0531 0.4446 1.0000
7.000 1.0740 0.01567 0.00875 -0.0513 0.4353 1.0000
7.250 1.0945 0.01597 0.00903 -0.0499 0.4270 1.0000
7.500 1.1122 0.01623 0.00935 -0.0481 0.4174 1.0000
7.750 1.1303 0.01653 0.00967 -0.0464 0.4076 1.0000
8.000 1.1480 0.01680 0.00989 -0.0445 0.3965 1.0000
8.250 1.1610 0.01704 0.01023 -0.0419 0.3839 1.0000
8.500 1.1743 0.01732 0.01054 -0.0393 0.3708 1.0000
8.750 1.1865 0.01761 0.01084 -0.0365 0.3572 1.0000
9.000 1.1969 0.01794 0.01116 -0.0335 0.3429 1.0000
9.250 1.2055 0.01832 0.01154 -0.0302 0.3281 1.0000
9.500 1.2108 0.01876 0.01196 -0.0264 0.3127 1.0000
9.750 1.2128 0.01925 0.01242 -0.0221 0.2984 1.0000
10.000 1.2153 0.01986 0.01298 -0.0182 0.2841 1.0000
10.250 1.2193 0.02057 0.01367 -0.0148 0.2701 1.0000
10.500 1.2238 0.02139 0.01448 -0.0117 0.2567 1.0000
10.750 1.2288 0.02230 0.01538 -0.0090 0.2445 1.0000
11.000 1.2334 0.02333 0.01639 -0.0065 0.2330 1.0000
11.250 1.2372 0.02450 0.01750 -0.0041 0.2227 1.0000
11.500 1.2430 0.02563 0.01867 -0.0022 0.2114 1.0000
11.750 1.2481 0.02689 0.01995 -0.0003 0.2011 1.0000
12.000 1.2507 0.02837 0.02139 0.0016 0.1915 1.0000
12.250 1.2557 0.02978 0.02286 0.0030 0.1810 1.0000
12.500 1.2589 0.03138 0.02450 0.0044 0.1703 1.0000
12.750 1.2595 0.03325 0.02638 0.0058 0.1593 1.0000
13.000 1.2580 0.03539 0.02850 0.0070 0.1476 1.0000
13.250 1.2546 0.03779 0.03089 0.0081 0.1345 1.0000
13.500 1.2495 0.04048 0.03359 0.0090 0.1196 1.0000
13.750 1.2422 0.04348 0.03656 0.0098 0.1060 1.0000
14.000 1.2337 0.04675 0.03978 0.0104 0.0956 1.0000
14.250 1.2269 0.04998 0.04299 0.0108 0.0870 1.0000
14.500 1.2225 0.05307 0.04610 0.0110 0.0804 1.0000
14.750 1.2159 0.05647 0.04947 0.0111 0.0759 1.0000
15.000 1.2137 0.05947 0.05253 0.0112 0.0717 1.0000
15.250 1.2111 0.06258 0.05567 0.0111 0.0684 1.0000
15.500 1.2065 0.06586 0.05890 0.0111 0.0655 1.0000
15.750 1.2073 0.06869 0.06184 0.0109 0.0627 1.0000
16.000 1.2074 0.07157 0.06477 0.0108 0.0604 1.0000
16.250 1.2067 0.07443 0.06758 0.0107 0.0580 1.0000
16.500 1.2096 0.07688 0.07008 0.0109 0.0561 1.0000
16.750 1.2111 0.07972 0.07302 0.0106 0.0541 1.0000
17.000 1.2140 0.08227 0.07562 0.0104 0.0523 1.0000
17.250 1.2191 0.08432 0.07760 0.0107 0.0503 1.0000
17.500 1.2257 0.08624 0.07956 0.0111 0.0488 1.0000
17.750 1.2239 0.08972 0.08321 0.0102 0.0472 1.0000
18.000 1.2270 0.09235 0.08592 0.0099 0.0460 1.0000
18.250 1.2318 0.09467 0.08826 0.0098 0.0446 1.0000
18.500 1.2453 0.09540 0.08890 0.0108 0.0430 1.0000
18.750 1.2454 0.09854 0.09217 0.0103 0.0421 1.0000
19.000 1.2393 0.10282 0.09665 0.0087 0.0413 1.0000
19.250 1.2352 0.10677 0.10076 0.0074 0.0403 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF-48 AIRFOIL (raf48-il)