NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 100,000 Max Cl/Cd: 45.78 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08n1-il-100000-n5.txt Download as CSV file: xf-rc08n1-il-100000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5797 0.09666 0.09226 0.0105 1.0000 0.0403
-8.000 -0.5794 0.09218 0.08783 0.0051 1.0000 0.0417
-7.750 -0.5791 0.08751 0.08313 -0.0027 1.0000 0.0429
-7.500 -0.5758 0.08368 0.07912 -0.0084 1.0000 0.0436
-7.250 -0.5706 0.08028 0.07548 -0.0112 1.0000 0.0438
-7.000 -0.5628 0.07328 0.06870 -0.0110 1.0000 0.0451
-6.750 -0.5530 0.06930 0.06477 -0.0106 1.0000 0.0471
-6.500 -0.5417 0.06549 0.06089 -0.0118 1.0000 0.0500
-6.250 -0.5247 0.06387 0.05863 -0.0147 1.0000 0.0560
-5.750 -0.4984 0.05348 0.04827 -0.0157 1.0000 0.0591
-5.500 -0.4814 0.04999 0.04467 -0.0160 1.0000 0.0620
-5.250 -0.4621 0.04743 0.04155 -0.0163 1.0000 0.0706
-5.000 -0.4447 0.04348 0.03766 -0.0165 1.0000 0.0749
-4.500 -0.3903 0.03472 0.02768 -0.0142 0.9904 0.0344
-4.250 -0.3571 0.03165 0.02417 -0.0153 0.9670 0.0310
-4.000 -0.3253 0.02875 0.02085 -0.0161 0.9505 0.0298
-3.750 -0.2937 0.02644 0.01808 -0.0166 0.9364 0.0299
-3.500 -0.2626 0.02492 0.01607 -0.0165 0.9238 0.0321
-3.250 -0.2335 0.02305 0.01389 -0.0164 0.9125 0.0319
-3.000 -0.2052 0.02147 0.01206 -0.0160 0.9022 0.0316
-2.750 -0.1778 0.02014 0.01050 -0.0154 0.8919 0.0316
-2.500 -0.1513 0.01902 0.00925 -0.0148 0.8823 0.0317
-2.250 -0.1260 0.01806 0.00821 -0.0140 0.8737 0.0321
-2.000 -0.1017 0.01726 0.00736 -0.0131 0.8646 0.0330
-1.750 -0.0780 0.01662 0.00672 -0.0124 0.8560 0.0369
-1.500 -0.0536 0.01618 0.00614 -0.0115 0.8482 0.0396
-1.250 -0.0280 0.01579 0.00557 -0.0109 0.8398 0.0415
-1.000 -0.0024 0.01552 0.00513 -0.0102 0.8327 0.0442
-0.750 0.0238 0.01524 0.00476 -0.0098 0.8244 0.0514
-0.500 0.0490 0.01446 0.00445 -0.0094 0.8174 0.1993
-0.250 0.1354 0.01218 0.00452 -0.0202 0.8139 1.0000
0.000 0.1598 0.01226 0.00443 -0.0195 0.8063 1.0000
0.250 0.1853 0.01235 0.00439 -0.0191 0.7974 1.0000
0.500 0.2099 0.01244 0.00437 -0.0185 0.7894 1.0000
0.750 0.2347 0.01254 0.00437 -0.0180 0.7807 1.0000
1.000 0.2598 0.01265 0.00441 -0.0175 0.7715 1.0000
1.250 0.2843 0.01275 0.00444 -0.0168 0.7632 1.0000
1.500 0.3093 0.01285 0.00451 -0.0162 0.7535 1.0000
1.750 0.3343 0.01296 0.00461 -0.0157 0.7436 1.0000
2.000 0.3590 0.01305 0.00469 -0.0150 0.7342 1.0000
2.250 0.3837 0.01314 0.00477 -0.0143 0.7243 1.0000
2.500 0.4088 0.01325 0.00491 -0.0138 0.7132 1.0000
2.750 0.4338 0.01335 0.00504 -0.0132 0.7023 1.0000
3.000 0.4587 0.01343 0.00521 -0.0125 0.6908 1.0000
3.250 0.4831 0.01347 0.00529 -0.0116 0.6759 1.0000
3.500 0.5068 0.01343 0.00527 -0.0104 0.6559 1.0000
3.750 0.5308 0.01342 0.00530 -0.0094 0.6318 1.0000
4.000 0.5551 0.01345 0.00542 -0.0084 0.6079 1.0000
4.250 0.5793 0.01351 0.00551 -0.0074 0.5801 1.0000
4.500 0.6032 0.01362 0.00562 -0.0064 0.5455 1.0000
4.750 0.6267 0.01381 0.00576 -0.0054 0.4989 1.0000
5.000 0.6491 0.01418 0.00594 -0.0043 0.4331 1.0000
5.250 0.6700 0.01491 0.00635 -0.0032 0.3464 1.0000
5.500 0.6898 0.01598 0.00697 -0.0024 0.2510 1.0000
5.750 0.7096 0.01716 0.00775 -0.0018 0.1747 1.0000
6.000 0.7303 0.01822 0.00862 -0.0011 0.1333 1.0000
6.250 0.7509 0.01925 0.00952 -0.0005 0.1032 1.0000
6.500 0.7719 0.02024 0.01054 0.0003 0.0834 1.0000
6.750 0.7922 0.02131 0.01166 0.0011 0.0675 1.0000
7.000 0.8124 0.02241 0.01281 0.0019 0.0555 1.0000
7.250 0.8311 0.02381 0.01431 0.0031 0.0486 1.0000
7.500 0.8505 0.02508 0.01566 0.0040 0.0418 1.0000
7.750 0.8698 0.02652 0.01727 0.0050 0.0360 1.0000
8.000 0.8890 0.02818 0.01909 0.0061 0.0324 1.0000
8.250 0.9064 0.03033 0.02135 0.0071 0.0294 1.0000
8.500 0.9258 0.03217 0.02357 0.0081 0.0258 1.0000
8.750 0.9433 0.03448 0.02623 0.0092 0.0232 1.0000
9.000 0.9585 0.03710 0.02918 0.0103 0.0216 1.0000
9.250 0.9705 0.03995 0.03234 0.0114 0.0204 1.0000
9.500 0.9765 0.04362 0.03631 0.0125 0.0191 1.0000
9.750 0.9789 0.04715 0.04032 0.0139 0.0181 1.0000
10.000 0.9775 0.05070 0.04435 0.0152 0.0171 1.0000
10.250 0.9692 0.05457 0.04861 0.0164 0.0165 1.0000
10.500 0.9548 0.05849 0.05281 0.0171 0.0164 1.0000
10.750 0.9382 0.06302 0.05758 0.0163 0.0163 1.0000
11.000 0.9202 0.06830 0.06305 0.0141 0.0164 1.0000
11.250 0.9013 0.07445 0.06937 0.0105 0.0166 1.0000
11.500 0.8821 0.08161 0.07666 0.0056 0.0171 1.0000
11.750 0.8635 0.08980 0.08492 -0.0001 0.0176 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il)