RAF 38 AIRFOIL (raf38-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 38 AIRFOIL (raf38-il) Reynolds number: 50,000 Max Cl/Cd: 31.62 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf38-il-50000.txt Download as CSV file: xf-raf38-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 38 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3689 0.10634 0.09892 -0.0175 1.0000 0.3230
-9.250 -0.3721 0.10373 0.09638 -0.0174 1.0000 0.3341
-9.000 -0.3918 0.10304 0.09582 -0.0171 1.0000 0.3470
-8.750 -0.3702 0.09875 0.09151 -0.0162 1.0000 0.3618
-8.500 -0.3652 0.09595 0.08875 -0.0152 1.0000 0.3781
-8.250 -0.3513 0.09239 0.08520 -0.0144 1.0000 0.3923
-8.000 -0.3432 0.08913 0.08198 -0.0137 1.0000 0.4033
-7.750 -0.3410 0.08646 0.07938 -0.0126 1.0000 0.4179
-7.500 -0.3400 0.08394 0.07693 -0.0112 1.0000 0.4330
-7.250 -0.5680 0.06270 0.05592 -0.0334 1.0000 0.1834
-7.000 -0.5741 0.05941 0.05260 -0.0309 1.0000 0.1795
-6.750 -0.6092 0.05426 0.04675 -0.0280 1.0000 0.1694
-6.500 -0.6107 0.05123 0.04350 -0.0254 1.0000 0.1690
-6.250 -0.6115 0.04833 0.04023 -0.0229 1.0000 0.1693
-6.000 -0.6055 0.04564 0.03741 -0.0209 1.0000 0.1711
-5.750 -0.5978 0.04310 0.03459 -0.0189 1.0000 0.1721
-5.500 -0.5872 0.04110 0.03242 -0.0171 1.0000 0.1753
-5.250 -0.5760 0.03915 0.03012 -0.0153 1.0000 0.1793
-5.000 -0.5630 0.03714 0.02765 -0.0137 1.0000 0.1826
-4.750 -0.5478 0.03533 0.02543 -0.0123 1.0000 0.1867
-4.500 -0.5308 0.03407 0.02415 -0.0110 1.0000 0.1935
-4.250 -0.5131 0.03274 0.02242 -0.0098 1.0000 0.1996
-4.000 -0.4946 0.03148 0.02105 -0.0087 1.0000 0.2064
-3.750 -0.4760 0.03056 0.01986 -0.0075 1.0000 0.2168
-3.500 -0.4567 0.02955 0.01891 -0.0065 1.0000 0.2272
-3.250 -0.4366 0.02863 0.01796 -0.0055 1.0000 0.2402
-3.000 -0.4165 0.02783 0.01718 -0.0046 1.0000 0.2617
-2.750 -0.3940 0.02693 0.01646 -0.0041 1.0000 0.2952
-2.500 -0.3742 0.02567 0.01586 -0.0028 1.0000 0.3799
-2.250 -0.3639 0.02420 0.01577 0.0010 1.0000 0.5675
-2.000 -0.1989 0.02585 0.01822 -0.0183 1.0000 1.0000
-1.750 -0.2070 0.02543 0.01753 -0.0150 1.0000 1.0000
-1.500 -0.1747 0.02587 0.01759 -0.0182 0.9916 1.0000
-1.250 -0.1335 0.02656 0.01792 -0.0227 0.9799 1.0000
-1.000 -0.0934 0.02728 0.01833 -0.0266 0.9684 1.0000
-0.750 -0.0554 0.02798 0.01877 -0.0300 0.9566 1.0000
-0.500 -0.0245 0.02858 0.01916 -0.0320 0.9444 1.0000
-0.250 0.0071 0.02926 0.01965 -0.0340 0.9325 1.0000
0.000 0.0427 0.03003 0.02024 -0.0366 0.9211 1.0000
0.250 0.0784 0.03079 0.02085 -0.0391 0.9096 1.0000
0.500 0.1007 0.03145 0.02139 -0.0392 0.8973 1.0000
0.750 0.1274 0.03223 0.02206 -0.0401 0.8859 1.0000
1.000 0.1657 0.03307 0.02279 -0.0428 0.8749 1.0000
1.250 0.1854 0.03384 0.02348 -0.0424 0.8629 1.0000
1.500 0.2050 0.03469 0.02426 -0.0420 0.8514 1.0000
1.750 0.2374 0.03559 0.02510 -0.0436 0.8409 1.0000
2.000 0.2589 0.03649 0.02595 -0.0435 0.8297 1.0000
2.250 0.2736 0.03748 0.02690 -0.0424 0.8185 1.0000
2.500 0.3053 0.03843 0.02783 -0.0437 0.8083 1.0000
2.750 0.3228 0.03944 0.02881 -0.0430 0.7972 1.0000
3.000 0.3365 0.04054 0.02989 -0.0418 0.7861 1.0000
3.250 0.3676 0.04152 0.03089 -0.0429 0.7755 1.0000
3.500 0.3873 0.04256 0.03193 -0.0424 0.7638 1.0000
3.750 0.3981 0.04375 0.03313 -0.0408 0.7518 1.0000
4.000 0.4212 0.04480 0.03420 -0.0407 0.7397 1.0000
4.250 0.4565 0.04560 0.03506 -0.0418 0.7272 1.0000
4.500 0.4763 0.04661 0.03612 -0.0411 0.7140 1.0000
4.750 0.4867 0.04788 0.03742 -0.0394 0.7005 1.0000
5.000 0.5030 0.04907 0.03865 -0.0385 0.6871 1.0000
5.250 0.5243 0.05015 0.03980 -0.0379 0.6735 1.0000
5.500 0.5504 0.05108 0.04082 -0.0377 0.6600 1.0000
5.750 0.5844 0.05170 0.04155 -0.0380 0.6465 1.0000
6.000 0.6049 0.05273 0.04270 -0.0371 0.6325 1.0000
6.250 0.6159 0.05416 0.04420 -0.0356 0.6176 1.0000
6.500 0.6292 0.05553 0.04566 -0.0343 0.6027 1.0000
6.750 0.6422 0.05700 0.04721 -0.0331 0.5878 1.0000
7.000 0.6570 0.05840 0.04872 -0.0320 0.5728 1.0000
7.250 0.6707 0.05992 0.05035 -0.0308 0.5580 1.0000
7.500 0.6877 0.06117 0.05171 -0.0297 0.5425 1.0000
7.750 0.7048 0.06240 0.05307 -0.0285 0.5269 1.0000
8.000 0.7790 0.05708 0.04811 -0.0258 0.4998 1.0000
8.250 0.9996 0.03630 0.02801 -0.0262 0.4523 1.0000
8.500 1.0399 0.03351 0.02510 -0.0237 0.4087 1.0000
8.750 1.0476 0.03313 0.02453 -0.0188 0.3585 1.0000
9.000 1.0484 0.03361 0.02455 -0.0135 0.3055 1.0000
9.250 1.0502 0.03501 0.02550 -0.0094 0.2643 1.0000
9.500 1.0619 0.03681 0.02697 -0.0070 0.2338 1.0000
9.750 1.0821 0.03875 0.02862 -0.0058 0.2100 1.0000
10.000 1.0988 0.04085 0.03071 -0.0044 0.1929 1.0000
10.250 1.1180 0.04304 0.03292 -0.0034 0.1787 1.0000
10.500 1.1393 0.04535 0.03518 -0.0027 0.1663 1.0000
10.750 1.1469 0.04778 0.03791 -0.0005 0.1584 1.0000
11.000 1.1607 0.05049 0.04072 0.0007 0.1506 1.0000
11.250 1.1624 0.05325 0.04378 0.0032 0.1454 1.0000
11.500 1.1838 0.05644 0.04697 0.0034 0.1388 1.0000
11.750 1.1677 0.05949 0.05042 0.0072 0.1371 1.0000
12.000 1.1481 0.06271 0.05395 0.0108 0.1359 1.0000
12.250 1.1255 0.06640 0.05790 0.0136 0.1353 1.0000
12.500 1.0983 0.07072 0.06248 0.0153 0.1353 1.0000
12.750 1.0664 0.07605 0.06803 0.0158 0.1360 1.0000
13.000 1.0328 0.08237 0.07450 0.0148 0.1373 1.0000
13.250 0.9994 0.08975 0.08200 0.0126 0.1386 1.0000
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Polar data table (+)
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