RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 26 AIRFOIL (raf26-il) Reynolds number: 500,000 Max Cl/Cd: 69.76 at α=1.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf26-il-500000-n5.txt Download as CSV file: xf-raf26-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 26 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5008 0.09019 0.08795 -0.0183 1.0000 0.0051
-8.500 -0.5010 0.08695 0.08474 -0.0193 1.0000 0.0049
-8.250 -0.5035 0.08331 0.08112 -0.0204 1.0000 0.0048
-8.000 -0.5093 0.07964 0.07750 -0.0213 1.0000 0.0048
-7.750 -0.5173 0.07629 0.07420 -0.0218 1.0000 0.0047
-7.500 -0.5179 0.07179 0.06972 -0.0251 1.0000 0.0046
-7.250 -0.5174 0.06624 0.06418 -0.0296 1.0000 0.0045
-7.000 -0.5137 0.06023 0.05813 -0.0343 1.0000 0.0044
-6.750 -0.5083 0.05344 0.05125 -0.0385 1.0000 0.0043
-6.500 -0.4984 0.04548 0.04309 -0.0425 0.9995 0.0042
-6.250 -0.4738 0.03244 0.02942 -0.0494 0.9957 0.0039
-6.000 -0.4535 0.02190 0.01783 -0.0509 0.9915 0.0039
-5.750 -0.4258 0.01872 0.01411 -0.0517 0.9888 0.0041
-5.500 -0.3972 0.01649 0.01146 -0.0522 0.9865 0.0046
-5.250 -0.3702 0.01495 0.00957 -0.0522 0.9827 0.0050
-5.000 -0.3407 0.01374 0.00812 -0.0528 0.9797 0.0057
-4.750 -0.3128 0.01271 0.00694 -0.0530 0.9759 0.0062
-4.500 -0.2853 0.01189 0.00600 -0.0529 0.9707 0.0067
-4.250 -0.2570 0.01114 0.00513 -0.0531 0.9659 0.0074
-3.750 -0.2040 0.00988 0.00363 -0.0524 0.9490 0.0100
-3.500 -0.1725 0.00935 0.00301 -0.0531 0.9427 0.0130
-3.250 -0.1406 0.00897 0.00256 -0.0540 0.9355 0.0176
-3.000 -0.1049 0.00857 0.00217 -0.0557 0.9300 0.0339
-2.750 -0.0680 0.00822 0.00189 -0.0579 0.9247 0.0584
-2.500 -0.0286 0.00788 0.00166 -0.0606 0.9193 0.0988
-2.250 0.0139 0.00749 0.00141 -0.0641 0.9111 0.1544
-2.000 0.0558 0.00717 0.00121 -0.0675 0.8967 0.2194
-1.750 0.0888 0.00690 0.00109 -0.0688 0.8806 0.2929
-1.500 0.1166 0.00663 0.00104 -0.0689 0.8671 0.3853
-1.250 0.1433 0.00645 0.00099 -0.0687 0.8537 0.4460
-1.000 0.1693 0.00633 0.00096 -0.0683 0.8391 0.4947
-0.750 0.1947 0.00622 0.00092 -0.0678 0.8241 0.5418
-0.500 0.2183 0.00603 0.00092 -0.0668 0.8069 0.6178
-0.250 0.2382 0.00571 0.00093 -0.0649 0.7861 0.7383
0.000 0.2634 0.00535 0.00095 -0.0640 0.7653 0.8815
0.500 0.3768 0.00573 0.00098 -0.0771 0.6712 1.0000
0.750 0.3989 0.00592 0.00101 -0.0758 0.6368 1.0000
1.000 0.4212 0.00612 0.00106 -0.0745 0.6039 1.0000
1.250 0.4430 0.00635 0.00112 -0.0732 0.5572 1.0000
1.500 0.4609 0.00686 0.00122 -0.0712 0.4577 1.0000
1.750 0.4824 0.00722 0.00138 -0.0699 0.4113 1.0000
2.000 0.5051 0.00751 0.00153 -0.0689 0.3738 1.0000
2.250 0.5235 0.00818 0.00172 -0.0672 0.2549 1.0000
2.500 0.5434 0.00875 0.00200 -0.0657 0.1785 1.0000
2.750 0.5643 0.00926 0.00224 -0.0644 0.1076 1.0000
3.000 0.5873 0.00958 0.00246 -0.0635 0.0889 1.0000
3.250 0.6112 0.00982 0.00270 -0.0627 0.0810 1.0000
3.500 0.6350 0.01008 0.00299 -0.0619 0.0734 1.0000
3.750 0.6592 0.01030 0.00323 -0.0611 0.0673 1.0000
4.000 0.6833 0.01052 0.00344 -0.0604 0.0545 1.0000
4.250 0.7050 0.01101 0.00372 -0.0593 0.0200 1.0000
4.500 0.7275 0.01145 0.00417 -0.0581 0.0122 1.0000
4.750 0.7492 0.01200 0.00483 -0.0568 0.0086 1.0000
5.000 0.7717 0.01247 0.00539 -0.0557 0.0077 1.0000
5.250 0.7935 0.01303 0.00605 -0.0544 0.0069 1.0000
5.500 0.8151 0.01360 0.00669 -0.0532 0.0061 1.0000
5.750 0.8345 0.01446 0.00760 -0.0516 0.0051 1.0000
6.000 0.8555 0.01516 0.00840 -0.0503 0.0049 1.0000
6.250 0.8758 0.01602 0.00941 -0.0488 0.0045 1.0000
6.500 0.8958 0.01701 0.01053 -0.0473 0.0040 1.0000
6.750 0.9157 0.01815 0.01181 -0.0457 0.0039 1.0000
7.000 0.9356 0.01936 0.01318 -0.0443 0.0037 1.0000
7.250 0.9549 0.02075 0.01476 -0.0428 0.0035 1.0000
7.500 0.9734 0.02245 0.01669 -0.0411 0.0035 1.0000
7.750 0.9905 0.02431 0.01880 -0.0394 0.0033 1.0000
8.000 1.0052 0.02654 0.02133 -0.0373 0.0033 1.0000
8.250 1.0140 0.02964 0.02479 -0.0347 0.0031 1.0000
8.500 1.0251 0.03233 0.02793 -0.0320 0.0029 1.0000
8.750 1.0268 0.03673 0.03286 -0.0284 0.0027 1.0000
9.000 1.0194 0.04225 0.03886 -0.0241 0.0026 1.0000
9.250 1.0032 0.04927 0.04629 -0.0197 0.0026 1.0000
9.500 0.9874 0.05463 0.05190 -0.0160 0.0026 1.0000
9.750 0.9670 0.05871 0.05615 -0.0121 0.0026 1.0000
10.000 0.9504 0.06200 0.05957 -0.0101 0.0027 1.0000
10.250 0.9288 0.06702 0.06472 -0.0101 0.0027 1.0000
10.500 0.9064 0.07351 0.07133 -0.0130 0.0027 1.0000
10.750 0.8897 0.08087 0.07878 -0.0186 0.0028 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 26 AIRFOIL (raf26-il)