NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Reynolds number: 500,000 Max Cl/Cd: 60.01 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08b3-il-500000-n5.txt Download as CSV file: xf-rc08b3-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(B)3 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6342 0.08502 0.08279 -0.0067 1.0000 0.0106
-9.000 -0.6442 0.07828 0.07607 -0.0135 1.0000 0.0106
-8.750 -0.6563 0.07194 0.06969 -0.0179 1.0000 0.0105
-8.500 -0.6636 0.06510 0.06275 -0.0213 1.0000 0.0105
-8.250 -0.6686 0.05816 0.05566 -0.0232 1.0000 0.0106
-8.000 -0.6715 0.05117 0.04844 -0.0235 1.0000 0.0107
-7.750 -0.6737 0.04376 0.04071 -0.0224 1.0000 0.0109
-7.500 -0.6797 0.03488 0.03129 -0.0193 1.0000 0.0111
-7.250 -0.6813 0.02825 0.02402 -0.0153 1.0000 0.0114
-7.000 -0.6714 0.02545 0.02084 -0.0125 1.0000 0.0118
-6.750 -0.6517 0.02314 0.01815 -0.0115 0.9977 0.0120
-6.500 -0.6241 0.02115 0.01581 -0.0120 0.9945 0.0122
-6.250 -0.5984 0.01886 0.01313 -0.0121 0.9902 0.0125
-6.000 -0.5708 0.01698 0.01099 -0.0125 0.9865 0.0129
-5.750 -0.5416 0.01593 0.00981 -0.0131 0.9827 0.0133
-5.500 -0.5127 0.01510 0.00888 -0.0135 0.9779 0.0136
-5.250 -0.4826 0.01434 0.00802 -0.0142 0.9741 0.0140
-5.000 -0.4542 0.01365 0.00723 -0.0145 0.9689 0.0143
-4.750 -0.4258 0.01301 0.00651 -0.0147 0.9635 0.0147
-4.500 -0.3968 0.01249 0.00593 -0.0150 0.9588 0.0153
-4.250 -0.3706 0.01203 0.00541 -0.0147 0.9519 0.0159
-4.000 -0.3439 0.01153 0.00485 -0.0145 0.9459 0.0163
-3.750 -0.3184 0.01109 0.00434 -0.0139 0.9390 0.0167
-3.500 -0.2927 0.01072 0.00391 -0.0134 0.9323 0.0171
-3.250 -0.2673 0.01038 0.00353 -0.0129 0.9252 0.0176
-3.000 -0.2421 0.00998 0.00308 -0.0122 0.9182 0.0188
-2.750 -0.2163 0.00972 0.00279 -0.0117 0.9109 0.0201
-2.500 -0.1903 0.00951 0.00255 -0.0113 0.9037 0.0217
-2.250 -0.1641 0.00933 0.00234 -0.0109 0.8962 0.0238
-2.000 -0.1382 0.00910 0.00214 -0.0104 0.8885 0.0334
-1.750 -0.1125 0.00885 0.00197 -0.0099 0.8796 0.0582
-1.500 -0.0867 0.00862 0.00183 -0.0094 0.8711 0.0883
-1.250 -0.0613 0.00833 0.00171 -0.0090 0.8617 0.1394
-1.000 -0.0370 0.00790 0.00158 -0.0084 0.8514 0.2354
-0.750 -0.0173 0.00697 0.00142 -0.0071 0.8401 0.4656
-0.500 -0.0035 0.00588 0.00128 -0.0042 0.8285 0.7302
-0.250 0.0235 0.00553 0.00149 -0.0032 0.8177 0.8996
0.000 0.0594 0.00574 0.00170 -0.0044 0.8059 0.9431
0.250 0.1000 0.00605 0.00196 -0.0067 0.7922 0.9656
0.500 0.1384 0.00622 0.00206 -0.0088 0.7771 0.9747
0.750 0.1745 0.00629 0.00205 -0.0105 0.7612 0.9777
1.000 0.2070 0.00637 0.00205 -0.0115 0.7437 0.9808
1.250 0.2358 0.00646 0.00205 -0.0117 0.7189 0.9845
1.500 0.2692 0.00655 0.00203 -0.0130 0.6901 0.9860
1.750 0.3018 0.00667 0.00203 -0.0141 0.6594 0.9876
2.000 0.3331 0.00686 0.00204 -0.0149 0.6124 0.9893
2.250 0.3630 0.00715 0.00208 -0.0156 0.5481 0.9910
2.500 0.3913 0.00764 0.00217 -0.0160 0.4542 0.9925
2.750 0.4194 0.00810 0.00229 -0.0165 0.3739 0.9938
3.000 0.4486 0.00849 0.00242 -0.0172 0.3098 0.9948
3.250 0.4783 0.00889 0.00258 -0.0180 0.2484 0.9957
3.500 0.5073 0.00936 0.00276 -0.0187 0.1827 0.9967
3.750 0.5364 0.00979 0.00296 -0.0194 0.1315 0.9978
4.000 0.5656 0.01018 0.00318 -0.0201 0.0933 0.9988
4.250 0.5945 0.01056 0.00344 -0.0207 0.0649 0.9997
4.500 0.6199 0.01087 0.00368 -0.0204 0.0497 1.0000
4.750 0.6438 0.01114 0.00392 -0.0198 0.0415 1.0000
5.000 0.6675 0.01140 0.00419 -0.0191 0.0359 1.0000
5.250 0.6908 0.01170 0.00449 -0.0183 0.0314 1.0000
5.500 0.7138 0.01205 0.00486 -0.0175 0.0279 1.0000
5.750 0.7369 0.01234 0.00520 -0.0167 0.0262 1.0000
6.000 0.7597 0.01266 0.00555 -0.0158 0.0242 1.0000
6.250 0.7819 0.01305 0.00593 -0.0148 0.0223 1.0000
6.500 0.8027 0.01363 0.00658 -0.0137 0.0206 1.0000
6.750 0.8245 0.01402 0.00703 -0.0126 0.0200 1.0000
7.000 0.8458 0.01446 0.00754 -0.0115 0.0193 1.0000
7.250 0.8669 0.01491 0.00805 -0.0103 0.0185 1.0000
7.500 0.8879 0.01534 0.00853 -0.0092 0.0175 1.0000
7.750 0.9087 0.01577 0.00902 -0.0081 0.0166 1.0000
8.000 0.9284 0.01634 0.00963 -0.0068 0.0158 1.0000
8.250 0.9454 0.01725 0.01061 -0.0050 0.0148 1.0000
8.500 0.9660 0.01767 0.01112 -0.0039 0.0143 1.0000
8.750 0.9859 0.01818 0.01171 -0.0026 0.0136 1.0000
9.000 1.0049 0.01878 0.01240 -0.0012 0.0129 1.0000
9.250 1.0237 0.01938 0.01310 0.0002 0.0123 1.0000
9.500 1.0425 0.01997 0.01376 0.0015 0.0117 1.0000
9.750 1.0605 0.02062 0.01447 0.0029 0.0112 1.0000
10.000 1.0740 0.02183 0.01578 0.0049 0.0105 1.0000
10.250 1.0921 0.02248 0.01656 0.0063 0.0101 1.0000
10.500 1.1094 0.02323 0.01745 0.0077 0.0096 1.0000
10.750 1.1261 0.02405 0.01840 0.0092 0.0091 1.0000
11.000 1.1421 0.02491 0.01940 0.0107 0.0086 1.0000
11.250 1.1575 0.02579 0.02040 0.0121 0.0082 1.0000
11.500 1.1725 0.02667 0.02138 0.0136 0.0079 1.0000
11.750 1.1863 0.02760 0.02243 0.0151 0.0075 1.0000
12.000 1.1963 0.02886 0.02381 0.0170 0.0073 1.0000
12.250 1.1972 0.03068 0.02580 0.0200 0.0070 1.0000
12.500 1.2004 0.03223 0.02755 0.0226 0.0069 1.0000
12.750 1.2011 0.03409 0.02961 0.0249 0.0067 1.0000
13.000 1.1995 0.03625 0.03198 0.0269 0.0066 1.0000
13.250 1.1953 0.03881 0.03476 0.0284 0.0065 1.0000
13.500 1.1877 0.04192 0.03809 0.0293 0.0064 1.0000
13.750 1.1778 0.04561 0.04200 0.0293 0.0063 1.0000
14.000 1.1644 0.05020 0.04681 0.0281 0.0062 1.0000
14.250 1.1470 0.05603 0.05286 0.0254 0.0062 1.0000
14.500 1.1255 0.06359 0.06063 0.0206 0.0062 1.0000
14.750 1.0974 0.07382 0.07108 0.0134 0.0063 1.0000
15.000 1.0587 0.08753 0.08499 0.0045 0.0064 1.0000
15.250 1.0025 0.10533 0.10293 -0.0052 0.0067 1.0000
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Polar data table (+)
Polar graphs
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