RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 200,000 Max Cl/Cd: 68.19 at α=1.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf6-il-200000-n5.txt Download as CSV file: xf-raf6-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 6 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3291 0.11931 0.11585 -0.0280 1.0000 0.0173
-9.500 -0.3311 0.11696 0.11355 -0.0279 1.0000 0.0176
-9.250 -0.3355 0.11486 0.11150 -0.0279 1.0000 0.0178
-9.000 -0.3296 0.11156 0.10823 -0.0323 0.9979 0.0180
-8.750 -0.3183 0.10768 0.10436 -0.0377 0.9951 0.0181
-8.500 -0.3077 0.10377 0.10046 -0.0425 0.9912 0.0182
-8.250 -0.2975 0.09874 0.09546 -0.0439 0.9889 0.0185
-8.000 -0.2832 0.09484 0.09155 -0.0448 0.9872 0.0189
-7.750 -0.2727 0.09131 0.08803 -0.0472 0.9827 0.0192
-7.500 -0.2603 0.08759 0.08432 -0.0510 0.9786 0.0196
-7.250 -0.2474 0.08367 0.08042 -0.0552 0.9720 0.0200
-7.000 -0.2285 0.07927 0.07600 -0.0611 0.9672 0.0205
-6.750 -0.2121 0.07510 0.07181 -0.0663 0.9606 0.0211
-6.500 -0.1943 0.07090 0.06758 -0.0714 0.9539 0.0217
-6.000 -0.1460 0.06221 0.05869 -0.0830 0.9410 0.0237
-5.750 -0.1110 0.05844 0.05470 -0.0888 0.9369 0.0245
-5.500 -0.0911 0.05509 0.05120 -0.0903 0.9298 0.0247
-5.250 -0.0655 0.05133 0.04725 -0.0928 0.9255 0.0248
-4.250 0.0311 0.03306 0.02804 -0.0988 0.9094 0.0175
-4.000 0.0515 0.03050 0.02524 -0.0980 0.9024 0.0171
-3.750 0.0811 0.02754 0.02193 -0.0986 0.8985 0.0167
-3.500 0.1156 0.02457 0.01853 -0.0997 0.8957 0.0167
-3.250 0.1484 0.02221 0.01565 -0.0999 0.8911 0.0180
-3.000 0.1766 0.02019 0.01326 -0.0996 0.8846 0.0184
-2.750 0.2126 0.01847 0.01129 -0.1012 0.8807 0.0191
-2.500 0.2475 0.01732 0.00995 -0.1023 0.8748 0.0203
-2.250 0.2814 0.01644 0.00892 -0.1032 0.8659 0.0228
-2.000 0.3184 0.01532 0.00760 -0.1045 0.8558 0.0243
-1.750 0.3578 0.01405 0.00621 -0.1066 0.8447 0.0277
-1.500 0.3914 0.01340 0.00549 -0.1075 0.8320 0.0313
-1.250 0.4219 0.01298 0.00496 -0.1079 0.8210 0.0360
-1.000 0.4533 0.01250 0.00441 -0.1085 0.8116 0.0414
-0.750 0.4823 0.01228 0.00410 -0.1086 0.8007 0.0499
-0.500 0.5099 0.01203 0.00382 -0.1084 0.7898 0.0647
-0.250 0.5315 0.01045 0.00375 -0.1080 0.7790 0.5577
0.000 0.5808 0.00962 0.00373 -0.1114 0.7647 1.0000
0.250 0.6050 0.00969 0.00364 -0.1105 0.7480 1.0000
0.500 0.6286 0.00977 0.00357 -0.1095 0.7281 1.0000
0.750 0.6527 0.00986 0.00350 -0.1085 0.7038 1.0000
1.000 0.6762 0.00999 0.00342 -0.1074 0.6709 1.0000
1.250 0.6976 0.01023 0.00334 -0.1058 0.6232 1.0000
1.500 0.7155 0.01063 0.00336 -0.1036 0.5634 1.0000
1.750 0.7317 0.01114 0.00348 -0.1011 0.5067 1.0000
2.000 0.7494 0.01162 0.00366 -0.0991 0.4638 1.0000
2.250 0.7689 0.01204 0.00385 -0.0975 0.4329 1.0000
2.500 0.7895 0.01242 0.00405 -0.0962 0.4075 1.0000
2.750 0.8106 0.01279 0.00426 -0.0949 0.3838 1.0000
3.000 0.8317 0.01316 0.00448 -0.0937 0.3616 1.0000
3.250 0.8534 0.01350 0.00470 -0.0926 0.3409 1.0000
3.500 0.8752 0.01385 0.00493 -0.0916 0.3227 1.0000
3.750 0.8970 0.01419 0.00519 -0.0905 0.3074 1.0000
4.000 0.9188 0.01454 0.00544 -0.0895 0.2942 1.0000
4.250 0.9405 0.01490 0.00572 -0.0885 0.2822 1.0000
4.500 0.9627 0.01524 0.00602 -0.0876 0.2719 1.0000
4.750 0.9843 0.01561 0.00632 -0.0866 0.2637 1.0000
5.000 1.0067 0.01594 0.00664 -0.0857 0.2566 1.0000
5.250 1.0278 0.01634 0.00700 -0.0846 0.2503 1.0000
5.500 1.0503 0.01666 0.00734 -0.0838 0.2440 1.0000
5.750 1.0716 0.01705 0.00771 -0.0827 0.2382 1.0000
6.000 1.0930 0.01743 0.00810 -0.0818 0.2332 1.0000
6.250 1.1146 0.01777 0.00849 -0.0808 0.2281 1.0000
6.500 1.1348 0.01818 0.00889 -0.0796 0.2233 1.0000
6.750 1.1551 0.01860 0.00934 -0.0784 0.2191 1.0000
7.000 1.1765 0.01895 0.00978 -0.0774 0.2146 1.0000
7.250 1.1968 0.01937 0.01023 -0.0763 0.2101 1.0000
7.500 1.2160 0.01987 0.01074 -0.0750 0.2062 1.0000
7.750 1.2370 0.02027 0.01124 -0.0740 0.2025 1.0000
8.000 1.2578 0.02068 0.01176 -0.0730 0.1985 1.0000
8.250 1.2774 0.02114 0.01229 -0.0719 0.1944 1.0000
8.500 1.2958 0.02168 0.01287 -0.0705 0.1906 1.0000
8.750 1.3156 0.02216 0.01346 -0.0694 0.1868 1.0000
9.000 1.3344 0.02255 0.01399 -0.0682 0.1804 1.0000
9.250 1.3502 0.02302 0.01450 -0.0665 0.1720 1.0000
9.500 1.3669 0.02345 0.01500 -0.0650 0.1616 1.0000
9.750 1.3842 0.02393 0.01556 -0.0637 0.1511 1.0000
10.000 1.4003 0.02452 0.01620 -0.0622 0.1398 1.0000
10.250 1.4135 0.02531 0.01693 -0.0605 0.1212 1.0000
10.500 1.4219 0.02655 0.01801 -0.0582 0.0927 1.0000
10.750 1.4127 0.02930 0.02028 -0.0541 0.0317 1.0000
11.000 1.4125 0.03133 0.02235 -0.0509 0.0201 1.0000
11.250 1.4166 0.03302 0.02419 -0.0484 0.0169 1.0000
11.500 1.4208 0.03473 0.02607 -0.0461 0.0149 1.0000
11.750 1.4229 0.03664 0.02815 -0.0439 0.0135 1.0000
12.000 1.4261 0.03850 0.03020 -0.0419 0.0124 1.0000
12.250 1.4264 0.04068 0.03257 -0.0400 0.0117 1.0000
12.500 1.4230 0.04328 0.03535 -0.0381 0.0112 1.0000
12.750 1.4193 0.04603 0.03830 -0.0366 0.0107 1.0000
13.000 1.4153 0.04893 0.04141 -0.0353 0.0103 1.0000
13.250 1.4087 0.05225 0.04492 -0.0344 0.0099 1.0000
13.500 1.3998 0.05605 0.04894 -0.0339 0.0095 1.0000
13.750 1.3889 0.06042 0.05349 -0.0340 0.0092 1.0000
14.000 1.3754 0.06549 0.05874 -0.0348 0.0090 1.0000
14.250 1.3596 0.07132 0.06477 -0.0363 0.0089 1.0000
14.500 1.3423 0.07791 0.07154 -0.0386 0.0089 1.0000
14.750 1.3237 0.08506 0.07889 -0.0415 0.0088 1.0000
15.000 1.3043 0.09264 0.08665 -0.0446 0.0088 1.0000
15.250 1.2848 0.10032 0.09450 -0.0478 0.0089 1.0000
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