RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 100,000 Max Cl/Cd: 51.9 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf6-il-100000.txt Download as CSV file: xf-raf6-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3894 0.10260 0.09793 -0.0289 1.0000 0.0659
-8.000 -0.4071 0.10113 0.09657 -0.0284 1.0000 0.0662
-7.750 -0.4198 0.09900 0.09451 -0.0307 1.0000 0.0666
-7.500 -0.4301 0.09688 0.09239 -0.0327 1.0000 0.0669
-7.250 -0.4377 0.09481 0.09025 -0.0338 1.0000 0.0671
-7.000 -0.4296 0.08834 0.08400 -0.0283 1.0000 0.0686
-6.750 -0.4273 0.08546 0.08117 -0.0249 1.0000 0.0702
-6.500 -0.4280 0.08282 0.07856 -0.0234 1.0000 0.0718
-6.250 -0.4286 0.08009 0.07584 -0.0228 1.0000 0.0737
-6.000 -0.4277 0.07718 0.07292 -0.0231 1.0000 0.0762
-5.750 -0.4207 0.07465 0.07014 -0.0278 1.0000 0.0808
-5.500 -0.4150 0.07068 0.06598 -0.0294 1.0000 0.0826
-5.250 -0.4097 0.06685 0.06233 -0.0268 1.0000 0.0845
-5.000 -0.4016 0.06422 0.05972 -0.0254 1.0000 0.0882
-4.750 -0.3824 0.06166 0.05665 -0.0292 1.0000 0.0978
-4.500 -0.3748 0.05792 0.05315 -0.0272 1.0000 0.1012
-4.250 -0.3546 0.05571 0.05052 -0.0292 1.0000 0.1135
-4.000 -0.3426 0.05236 0.04736 -0.0279 1.0000 0.1180
-3.750 -0.3229 0.04973 0.04451 -0.0289 1.0000 0.1312
-3.500 -0.3038 0.04736 0.04201 -0.0293 1.0000 0.1464
-3.250 -0.2854 0.04505 0.03964 -0.0293 1.0000 0.1634
-3.000 -0.2660 0.04322 0.03767 -0.0296 1.0000 0.1925
-1.750 -0.0727 0.03184 0.02377 -0.0378 0.9896 0.1109
-1.500 -0.0206 0.02989 0.02131 -0.0411 0.9844 0.0948
-1.250 0.0233 0.02833 0.01950 -0.0437 0.9771 0.0916
-1.000 0.0715 0.02714 0.01805 -0.0469 0.9697 0.0890
-0.750 0.1283 0.02619 0.01693 -0.0517 0.9592 0.0949
-0.500 0.1859 0.02475 0.01556 -0.0565 0.9473 0.1006
-0.250 0.2333 0.02372 0.01458 -0.0596 0.9347 0.1139
0.000 0.2739 0.02288 0.01381 -0.0617 0.9234 0.1388
0.250 0.3186 0.02005 0.01332 -0.0634 0.9164 1.0000
0.500 0.3673 0.01973 0.01267 -0.0667 0.9061 1.0000
0.750 0.4073 0.01938 0.01215 -0.0684 0.8931 1.0000
1.000 0.4485 0.01890 0.01156 -0.0702 0.8801 1.0000
1.250 0.4910 0.01828 0.01086 -0.0720 0.8670 1.0000
1.500 0.5357 0.01750 0.01002 -0.0741 0.8539 1.0000
1.750 0.5825 0.01665 0.00912 -0.0764 0.8389 1.0000
2.000 0.6298 0.01580 0.00823 -0.0787 0.8207 1.0000
2.250 0.6686 0.01521 0.00759 -0.0795 0.7939 1.0000
2.500 0.7062 0.01473 0.00700 -0.0801 0.7563 1.0000
2.750 0.7311 0.01453 0.00665 -0.0783 0.6968 1.0000
3.000 0.7583 0.01461 0.00620 -0.0769 0.6034 1.0000
3.250 0.7791 0.01534 0.00625 -0.0749 0.5225 1.0000
3.500 0.7992 0.01622 0.00663 -0.0733 0.4725 1.0000
3.750 0.8215 0.01706 0.00708 -0.0722 0.4392 1.0000
4.000 0.8448 0.01778 0.00760 -0.0714 0.4138 1.0000
4.250 0.8693 0.01850 0.00811 -0.0708 0.3948 1.0000
4.500 0.8941 0.01918 0.00866 -0.0703 0.3789 1.0000
4.750 0.9188 0.01981 0.00922 -0.0698 0.3650 1.0000
5.000 0.9436 0.02044 0.00982 -0.0694 0.3531 1.0000
5.250 0.9689 0.02111 0.01043 -0.0690 0.3426 1.0000
5.500 0.9947 0.02178 0.01103 -0.0688 0.3329 1.0000
5.750 1.0187 0.02242 0.01175 -0.0682 0.3235 1.0000
6.000 1.0452 0.02319 0.01242 -0.0682 0.3156 1.0000
6.250 1.0684 0.02383 0.01321 -0.0674 0.3070 1.0000
6.500 1.0944 0.02465 0.01398 -0.0673 0.2998 1.0000
6.750 1.1175 0.02538 0.01485 -0.0666 0.2923 1.0000
7.000 1.1441 0.02628 0.01568 -0.0666 0.2858 1.0000
7.250 1.1653 0.02705 0.01670 -0.0656 0.2785 1.0000
7.500 1.1921 0.02798 0.01755 -0.0657 0.2724 1.0000
7.750 1.2125 0.02897 0.01883 -0.0646 0.2664 1.0000
8.000 1.2357 0.02989 0.01987 -0.0640 0.2604 1.0000
8.250 1.2594 0.03105 0.02110 -0.0636 0.2550 1.0000
8.500 1.2781 0.03210 0.02243 -0.0623 0.2490 1.0000
8.750 1.3022 0.03317 0.02356 -0.0620 0.2440 1.0000
9.000 1.3219 0.03466 0.02529 -0.0610 0.2397 1.0000
9.250 1.3376 0.03611 0.02708 -0.0593 0.2350 1.0000
9.500 1.3589 0.03735 0.02847 -0.0586 0.2306 1.0000
9.750 1.3811 0.03880 0.02999 -0.0581 0.2262 1.0000
10.000 1.3883 0.04016 0.03181 -0.0552 0.2204 1.0000
10.250 1.4186 0.04012 0.03150 -0.0558 0.2120 1.0000
10.500 1.4210 0.04114 0.03299 -0.0521 0.2053 1.0000
10.750 1.4511 0.04089 0.03251 -0.0527 0.1977 1.0000
11.000 1.4514 0.04211 0.03422 -0.0488 0.1921 1.0000
11.250 1.4747 0.04146 0.03348 -0.0482 0.1843 1.0000
11.500 1.4784 0.04236 0.03472 -0.0448 0.1784 1.0000
11.750 1.4955 0.04129 0.03357 -0.0431 0.1700 1.0000
12.000 1.4910 0.04171 0.03434 -0.0385 0.1629 1.0000
12.250 1.4962 0.04061 0.03313 -0.0349 0.1547 1.0000
12.500 1.4838 0.04129 0.03422 -0.0294 0.1476 1.0000
12.750 1.4764 0.04138 0.03446 -0.0249 0.1387 1.0000
13.000 1.4659 0.04208 0.03533 -0.0210 0.1269 1.0000
13.250 1.4545 0.04376 0.03721 -0.0179 0.1091 1.0000
13.500 1.4396 0.04662 0.04000 -0.0156 0.0922 1.0000
13.750 1.4233 0.05042 0.04383 -0.0141 0.0795 1.0000
14.000 1.4061 0.05474 0.04821 -0.0132 0.0722 1.0000
14.250 1.3881 0.05954 0.05311 -0.0130 0.0678 1.0000
14.500 1.3699 0.06482 0.05854 -0.0137 0.0646 1.0000
14.750 1.3497 0.07082 0.06467 -0.0153 0.0626 1.0000
15.000 1.3292 0.07737 0.07134 -0.0177 0.0614 1.0000
15.250 1.3090 0.08419 0.07823 -0.0204 0.0602 1.0000
15.500 1.2912 0.09095 0.08517 -0.0231 0.0589 1.0000
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Polar data table (+)
Polar graphs
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