RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 200,000 Max Cl/Cd: 70.26 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf6-il-200000.txt Download as CSV file: xf-raf6-il-200000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3372 0.08516 0.08226 -0.0216 1.0000 0.0352
-7.500 -0.3486 0.08314 0.08030 -0.0195 1.0000 0.0355
-7.250 -0.3629 0.08119 0.07840 -0.0175 1.0000 0.0358
-7.000 -0.3725 0.07829 0.07553 -0.0175 0.9984 0.0360
-6.750 -0.4483 0.08152 0.07845 -0.0249 1.0000 0.0343
-6.500 -0.4490 0.07909 0.07605 -0.0225 1.0000 0.0348
-6.250 -0.4338 0.07549 0.07242 -0.0246 0.9977 0.0357
-6.000 -0.4086 0.07102 0.06788 -0.0299 0.9942 0.0370
-5.750 -0.3833 0.06649 0.06324 -0.0350 0.9895 0.0387
-5.500 -0.3337 0.06208 0.05823 -0.0451 0.9847 0.0423
-5.250 -0.3152 0.05636 0.05231 -0.0475 0.9795 0.0430
-5.000 -0.2953 0.05210 0.04814 -0.0491 0.9765 0.0441
-4.750 -0.2677 0.04902 0.04500 -0.0516 0.9740 0.0459
-4.500 -0.2433 0.04624 0.04207 -0.0529 0.9691 0.0485
-4.250 -0.2004 0.04644 0.04144 -0.0547 0.9639 0.0538
-4.000 -0.1773 0.03952 0.03472 -0.0578 0.9620 0.0567
-3.750 -0.1473 0.03743 0.03252 -0.0594 0.9574 0.0621
-3.500 -0.1126 0.03470 0.02932 -0.0611 0.9519 0.0697
-3.250 -0.0736 0.03232 0.02687 -0.0641 0.9487 0.0756
-3.000 -0.0408 0.03018 0.02443 -0.0655 0.9434 0.0863
-2.750 -0.0057 0.02833 0.02237 -0.0673 0.9379 0.1013
-2.000 0.1319 0.02136 0.01385 -0.0716 0.9227 0.0544
-1.750 0.1823 0.01899 0.01137 -0.0754 0.9197 0.0568
-1.500 0.2111 0.01793 0.01026 -0.0750 0.9106 0.0580
-1.250 0.2532 0.01675 0.00909 -0.0772 0.9075 0.0609
-1.000 0.2855 0.01609 0.00840 -0.0776 0.9007 0.0662
-0.750 0.3227 0.01489 0.00728 -0.0791 0.8956 0.0715
-0.500 0.3656 0.01398 0.00639 -0.0817 0.8921 0.0836
-0.250 0.3941 0.01337 0.00582 -0.0814 0.8825 0.1048
0.000 0.4586 0.01063 0.00555 -0.0880 0.8797 1.0000
0.250 0.4924 0.01036 0.00515 -0.0887 0.8680 1.0000
0.500 0.5286 0.01011 0.00477 -0.0900 0.8555 1.0000
0.750 0.5644 0.00991 0.00446 -0.0912 0.8407 1.0000
1.000 0.5990 0.00978 0.00419 -0.0921 0.8224 1.0000
1.250 0.6303 0.00974 0.00401 -0.0923 0.8009 1.0000
1.500 0.6554 0.00978 0.00394 -0.0913 0.7753 1.0000
1.750 0.6782 0.00985 0.00388 -0.0899 0.7448 1.0000
2.000 0.6991 0.00995 0.00381 -0.0880 0.6999 1.0000
2.250 0.7177 0.01023 0.00369 -0.0855 0.6233 1.0000
2.500 0.7300 0.01088 0.00377 -0.0820 0.5323 1.0000
2.750 0.7432 0.01159 0.00402 -0.0790 0.4674 1.0000
3.000 0.7596 0.01222 0.00431 -0.0768 0.4237 1.0000
3.250 0.7784 0.01276 0.00462 -0.0750 0.3902 1.0000
3.500 0.7985 0.01326 0.00492 -0.0736 0.3641 1.0000
3.750 0.8198 0.01370 0.00521 -0.0724 0.3437 1.0000
4.000 0.8419 0.01412 0.00551 -0.0713 0.3278 1.0000
4.250 0.8642 0.01456 0.00583 -0.0704 0.3152 1.0000
4.500 0.8873 0.01495 0.00616 -0.0695 0.3040 1.0000
4.750 0.9108 0.01534 0.00650 -0.0688 0.2942 1.0000
5.000 0.9341 0.01584 0.00688 -0.0681 0.2863 1.0000
5.250 0.9581 0.01618 0.00725 -0.0674 0.2782 1.0000
5.500 0.9821 0.01671 0.00768 -0.0669 0.2713 1.0000
5.750 1.0061 0.01706 0.00809 -0.0663 0.2644 1.0000
6.000 1.0305 0.01756 0.00852 -0.0658 0.2584 1.0000
6.250 1.0547 0.01801 0.00903 -0.0653 0.2524 1.0000
6.500 1.0786 0.01844 0.00946 -0.0647 0.2463 1.0000
6.750 1.1040 0.01907 0.01003 -0.0645 0.2410 1.0000
7.000 1.1275 0.01947 0.01056 -0.0638 0.2358 1.0000
7.250 1.1515 0.01997 0.01107 -0.0634 0.2306 1.0000
7.500 1.1767 0.02064 0.01173 -0.0632 0.2255 1.0000
7.750 1.1991 0.02106 0.01229 -0.0623 0.2202 1.0000
8.000 1.2230 0.02160 0.01286 -0.0619 0.2155 1.0000
8.250 1.2488 0.02240 0.01366 -0.0619 0.2110 1.0000
8.500 1.2703 0.02289 0.01434 -0.0609 0.2065 1.0000
8.750 1.2928 0.02340 0.01491 -0.0601 0.2018 1.0000
9.000 1.3154 0.02405 0.01557 -0.0596 0.1962 1.0000
9.250 1.3312 0.02426 0.01597 -0.0576 0.1900 1.0000
9.500 1.3502 0.02460 0.01627 -0.0564 0.1841 1.0000
9.750 1.3662 0.02497 0.01684 -0.0545 0.1785 1.0000
10.000 1.3825 0.02526 0.01724 -0.0527 0.1732 1.0000
10.250 1.4007 0.02578 0.01775 -0.0515 0.1683 1.0000
10.500 1.4131 0.02606 0.01829 -0.0490 0.1632 1.0000
10.750 1.4232 0.02625 0.01855 -0.0461 0.1578 1.0000
11.000 1.4348 0.02666 0.01909 -0.0437 0.1524 1.0000
11.250 1.4451 0.02694 0.01955 -0.0411 0.1460 1.0000
11.500 1.4552 0.02733 0.02006 -0.0387 0.1389 1.0000
11.750 1.4649 0.02767 0.02051 -0.0363 0.1279 1.0000
12.000 1.4764 0.02823 0.02113 -0.0344 0.1058 1.0000
12.250 1.4758 0.03011 0.02274 -0.0314 0.0640 1.0000
12.500 1.4658 0.03291 0.02531 -0.0277 0.0422 1.0000
12.750 1.4605 0.03535 0.02781 -0.0247 0.0368 1.0000
13.000 1.4550 0.03788 0.03049 -0.0222 0.0340 1.0000
13.250 1.4506 0.04043 0.03324 -0.0201 0.0323 1.0000
13.500 1.4432 0.04340 0.03639 -0.0184 0.0310 1.0000
13.750 1.4328 0.04686 0.04002 -0.0170 0.0300 1.0000
14.000 1.4196 0.05088 0.04420 -0.0163 0.0293 1.0000
14.250 1.4037 0.05555 0.04903 -0.0163 0.0287 1.0000
14.500 1.3865 0.06082 0.05446 -0.0170 0.0282 1.0000
14.750 1.3691 0.06657 0.06035 -0.0185 0.0279 1.0000
15.000 1.3567 0.07201 0.06597 -0.0203 0.0275 1.0000
15.250 1.3433 0.07785 0.07198 -0.0224 0.0272 1.0000
15.500 1.3300 0.08381 0.07809 -0.0247 0.0269 1.0000
15.750 1.3171 0.08970 0.08413 -0.0270 0.0265 1.0000
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Polar data table (+)
Polar graphs
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