NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 100,000 Max Cl/Cd: 48.69 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08n1-il-100000.txt Download as CSV file: xf-rc08n1-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(N)1 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5856 0.10287 0.09847 0.0102 1.0000 0.0613
-8.250 -0.5891 0.09874 0.09439 0.0028 1.0000 0.0619
-8.000 -0.5916 0.09488 0.09048 -0.0040 1.0000 0.0623
-7.750 -0.5930 0.09186 0.08727 -0.0089 1.0000 0.0626
-7.500 -0.5751 0.08502 0.08076 -0.0003 1.0000 0.0681
-7.250 -0.5694 0.08071 0.07643 -0.0037 1.0000 0.0712
-7.000 -0.5635 0.07673 0.07226 -0.0101 1.0000 0.0751
-6.750 -0.5580 0.07218 0.06750 -0.0136 1.0000 0.0772
-6.500 -0.5448 0.06747 0.06300 -0.0118 1.0000 0.0816
-6.250 -0.5332 0.06678 0.06158 -0.0159 1.0000 0.0900
-6.000 -0.4757 0.04768 0.04296 -0.0203 1.0000 0.1042
-5.750 -0.4627 0.04205 0.03767 -0.0192 1.0000 0.1096
-5.500 -0.4891 0.05238 0.04735 -0.0166 1.0000 0.1132
-5.250 -0.4730 0.04884 0.04368 -0.0171 1.0000 0.1249
-5.000 -0.4560 0.04582 0.04052 -0.0174 1.0000 0.1415
-4.750 -0.4409 0.04272 0.03731 -0.0174 1.0000 0.1636
-4.250 -0.4227 0.03797 0.03266 -0.0143 1.0000 0.2329
-4.000 -0.4225 0.03647 0.03119 -0.0108 1.0000 0.2730
-3.750 -0.4189 0.03434 0.02928 -0.0071 1.0000 0.3097
-3.500 -0.4134 0.03229 0.02731 -0.0038 1.0000 0.3472
-3.250 -0.4029 0.03037 0.02540 -0.0012 1.0000 0.3794
-3.000 -0.3800 0.02794 0.02267 -0.0016 0.9997 0.3795
-2.750 -0.2759 0.02591 0.01737 -0.0071 0.9944 0.0893
-2.500 -0.2336 0.02341 0.01450 -0.0088 0.9884 0.0742
-2.250 -0.1902 0.02167 0.01250 -0.0114 0.9830 0.0741
-2.000 -0.1502 0.02002 0.01073 -0.0134 0.9767 0.0713
-1.750 -0.1073 0.01868 0.00930 -0.0160 0.9711 0.0697
-1.500 -0.0685 0.01765 0.00829 -0.0181 0.9645 0.0705
-1.250 -0.0283 0.01687 0.00745 -0.0206 0.9577 0.0740
-1.000 0.0080 0.01627 0.00678 -0.0225 0.9498 0.0871
-0.750 0.1091 0.01278 0.00600 -0.0350 0.9595 1.0000
-0.500 0.1495 0.01288 0.00584 -0.0380 0.9497 1.0000
-0.250 0.1884 0.01299 0.00575 -0.0405 0.9399 1.0000
0.000 0.2213 0.01316 0.00579 -0.0418 0.9287 1.0000
0.250 0.2498 0.01338 0.00592 -0.0421 0.9168 1.0000
0.500 0.2761 0.01362 0.00609 -0.0419 0.9049 1.0000
0.750 0.3009 0.01389 0.00630 -0.0413 0.8932 1.0000
1.000 0.3244 0.01416 0.00653 -0.0404 0.8816 1.0000
1.250 0.3469 0.01444 0.00677 -0.0392 0.8702 1.0000
1.500 0.3688 0.01472 0.00702 -0.0378 0.8584 1.0000
1.750 0.3904 0.01502 0.00732 -0.0364 0.8460 1.0000
2.000 0.4120 0.01531 0.00765 -0.0351 0.8336 1.0000
2.250 0.4335 0.01560 0.00796 -0.0337 0.8212 1.0000
2.500 0.4550 0.01588 0.00826 -0.0322 0.8088 1.0000
2.750 0.4765 0.01614 0.00856 -0.0306 0.7960 1.0000
3.000 0.4976 0.01634 0.00881 -0.0288 0.7823 1.0000
3.250 0.5174 0.01638 0.00893 -0.0263 0.7659 1.0000
3.500 0.5366 0.01625 0.00881 -0.0232 0.7484 1.0000
3.750 0.5570 0.01615 0.00877 -0.0208 0.7288 1.0000
4.000 0.5783 0.01601 0.00869 -0.0185 0.7101 1.0000
4.250 0.5997 0.01574 0.00846 -0.0160 0.6917 1.0000
4.500 0.6219 0.01551 0.00839 -0.0140 0.6681 1.0000
4.750 0.6441 0.01518 0.00814 -0.0118 0.6419 1.0000
5.000 0.6664 0.01485 0.00788 -0.0096 0.6106 1.0000
5.250 0.6886 0.01459 0.00765 -0.0075 0.5669 1.0000
5.500 0.7094 0.01457 0.00753 -0.0054 0.4886 1.0000
5.750 0.7239 0.01575 0.00783 -0.0029 0.3197 1.0000
6.000 0.7373 0.01798 0.00911 -0.0014 0.1975 1.0000
6.250 0.7544 0.01975 0.01048 -0.0001 0.1489 1.0000
6.500 0.7734 0.02148 0.01203 0.0012 0.1209 1.0000
6.750 0.7932 0.02316 0.01351 0.0022 0.0999 1.0000
7.000 0.8153 0.02516 0.01546 0.0033 0.0864 1.0000
7.250 0.8374 0.02682 0.01728 0.0042 0.0744 1.0000
7.500 0.8609 0.02927 0.02005 0.0053 0.0673 1.0000
7.750 0.8828 0.03193 0.02276 0.0061 0.0611 1.0000
8.000 0.9019 0.03463 0.02605 0.0074 0.0560 1.0000
8.250 0.9195 0.03816 0.03013 0.0089 0.0540 1.0000
8.500 0.9333 0.04215 0.03466 0.0103 0.0530 1.0000
8.750 0.9483 0.04531 0.03780 0.0108 0.0489 1.0000
9.000 0.9517 0.05055 0.04352 0.0119 0.0476 1.0000
9.250 0.9569 0.05584 0.04922 0.0130 0.0481 1.0000
9.500 0.9400 0.06070 0.05507 0.0151 0.0526 1.0000
9.750 0.9240 0.06651 0.06123 0.0153 0.0554 1.0000
10.000 0.9090 0.07143 0.06631 0.0153 0.0571 1.0000
10.250 0.8057 0.06455 0.05971 0.0216 0.0565 1.0000
10.500 0.7849 0.07018 0.06540 0.0203 0.0579 1.0000
10.750 0.7676 0.07608 0.07134 0.0188 0.0588 1.0000
11.000 0.7571 0.08219 0.07746 0.0177 0.0598 1.0000
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Polar data table (+)
Polar graphs
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