NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Reynolds number: 1,000,000 Max Cl/Cd: 71.51 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08b3-il-1000000.txt Download as CSV file: xf-rc08b3-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.250 -0.4891 0.14251 0.14092 0.0063 1.0000 0.0104
-13.000 -0.4890 0.13923 0.13764 0.0052 1.0000 0.0104
-12.750 -0.6393 0.14984 0.14815 0.0237 1.0000 0.0103
-12.500 -0.6358 0.14561 0.14392 0.0219 1.0000 0.0103
-12.250 -0.6326 0.14127 0.13959 0.0201 1.0000 0.0104
-12.000 -0.6297 0.13680 0.13513 0.0183 1.0000 0.0104
-6.250 -0.5968 0.02116 0.01691 -0.0127 0.9947 0.0145
-6.000 -0.5678 0.01968 0.01523 -0.0133 0.9923 0.0144
-5.750 -0.5418 0.01645 0.01159 -0.0134 0.9894 0.0146
-5.500 -0.5128 0.01426 0.00915 -0.0139 0.9870 0.0149
-5.250 -0.4821 0.01301 0.00778 -0.0147 0.9845 0.0153
-5.000 -0.4522 0.01223 0.00695 -0.0153 0.9808 0.0157
-4.750 -0.4242 0.01157 0.00622 -0.0154 0.9755 0.0161
-4.500 -0.3954 0.01095 0.00554 -0.0156 0.9708 0.0165
-4.250 -0.3703 0.01040 0.00494 -0.0150 0.9638 0.0168
-4.000 -0.3452 0.00993 0.00441 -0.0143 0.9572 0.0172
-3.750 -0.3206 0.00954 0.00398 -0.0136 0.9501 0.0178
-3.500 -0.2960 0.00922 0.00361 -0.0127 0.9432 0.0185
-3.250 -0.2710 0.00891 0.00326 -0.0121 0.9362 0.0190
-3.000 -0.2464 0.00855 0.00283 -0.0112 0.9291 0.0197
-2.750 -0.2214 0.00817 0.00239 -0.0105 0.9218 0.0213
-2.500 -0.1958 0.00795 0.00214 -0.0099 0.9143 0.0230
-2.250 -0.1694 0.00777 0.00194 -0.0095 0.9065 0.0251
-2.000 -0.1437 0.00754 0.00171 -0.0089 0.8989 0.0342
-1.750 -0.1182 0.00720 0.00155 -0.0085 0.8902 0.0802
-1.500 -0.0929 0.00691 0.00142 -0.0079 0.8814 0.1348
-1.250 -0.0683 0.00650 0.00129 -0.0074 0.8718 0.2299
-1.000 -0.0461 0.00579 0.00116 -0.0065 0.8628 0.4088
-0.750 -0.0296 0.00475 0.00101 -0.0044 0.8534 0.6781
-0.500 -0.0135 0.00403 0.00106 -0.0014 0.8431 0.8924
-0.250 0.0152 0.00412 0.00120 -0.0011 0.8331 0.9380
0.000 0.0419 0.00426 0.00130 -0.0004 0.8229 0.9575
0.250 0.0733 0.00437 0.00136 -0.0010 0.8122 0.9637
0.500 0.1074 0.00455 0.00150 -0.0021 0.8013 0.9728
0.750 0.1582 0.00486 0.00175 -0.0069 0.7871 0.9812
1.000 0.2009 0.00501 0.00184 -0.0101 0.7703 0.9856
1.250 0.2397 0.00510 0.00186 -0.0126 0.7536 0.9881
1.500 0.2750 0.00519 0.00187 -0.0142 0.7297 0.9907
1.750 0.3072 0.00529 0.00187 -0.0152 0.7023 0.9928
2.000 0.3383 0.00546 0.00187 -0.0161 0.6550 0.9945
2.250 0.3716 0.00568 0.00188 -0.0174 0.5954 0.9953
2.500 0.4051 0.00596 0.00191 -0.0189 0.5297 0.9963
2.750 0.4376 0.00636 0.00198 -0.0203 0.4464 0.9973
3.000 0.4690 0.00681 0.00209 -0.0215 0.3595 0.9982
3.250 0.4995 0.00723 0.00221 -0.0225 0.2842 0.9989
3.500 0.5301 0.00762 0.00234 -0.0235 0.2216 0.9997
3.750 0.5579 0.00802 0.00250 -0.0239 0.1637 1.0000
4.000 0.5829 0.00843 0.00269 -0.0236 0.1133 1.0000
4.250 0.6077 0.00880 0.00289 -0.0233 0.0755 1.0000
4.500 0.6325 0.00913 0.00310 -0.0229 0.0525 1.0000
4.750 0.6571 0.00941 0.00333 -0.0224 0.0408 1.0000
5.000 0.6815 0.00968 0.00357 -0.0218 0.0343 1.0000
5.250 0.7055 0.00996 0.00387 -0.0212 0.0302 1.0000
5.500 0.7294 0.01020 0.00411 -0.0205 0.0275 1.0000
5.750 0.7525 0.01061 0.00454 -0.0197 0.0245 1.0000
6.000 0.7753 0.01101 0.00501 -0.0188 0.0231 1.0000
6.250 0.7986 0.01126 0.00529 -0.0180 0.0221 1.0000
6.500 0.8215 0.01155 0.00560 -0.0172 0.0208 1.0000
6.750 0.8440 0.01189 0.00596 -0.0163 0.0197 1.0000
7.000 0.8651 0.01241 0.00652 -0.0151 0.0184 1.0000
7.250 0.8840 0.01323 0.00744 -0.0136 0.0172 1.0000
7.500 0.9065 0.01347 0.00770 -0.0127 0.0166 1.0000
7.750 0.9278 0.01385 0.00813 -0.0115 0.0159 1.0000
8.000 0.9488 0.01426 0.00859 -0.0104 0.0151 1.0000
8.250 0.9696 0.01466 0.00902 -0.0092 0.0144 1.0000
8.500 0.9891 0.01520 0.00958 -0.0079 0.0136 1.0000
8.750 1.0031 0.01648 0.01099 -0.0056 0.0126 1.0000
9.000 1.0245 0.01670 0.01126 -0.0045 0.0122 1.0000
9.250 1.0443 0.01715 0.01177 -0.0032 0.0116 1.0000
9.500 1.0635 0.01765 0.01234 -0.0018 0.0111 1.0000
9.750 1.0827 0.01811 0.01285 -0.0005 0.0106 1.0000
10.000 1.1017 0.01857 0.01335 0.0009 0.0101 1.0000
10.250 1.1163 0.01956 0.01442 0.0028 0.0096 1.0000
10.500 1.1255 0.02130 0.01636 0.0056 0.0091 1.0000
10.750 1.1431 0.02191 0.01707 0.0071 0.0089 1.0000
11.000 1.1591 0.02272 0.01799 0.0087 0.0086 1.0000
11.250 1.1742 0.02365 0.01904 0.0104 0.0083 1.0000
11.500 1.1887 0.02464 0.02017 0.0120 0.0080 1.0000
11.750 1.2035 0.02553 0.02115 0.0136 0.0077 1.0000
12.000 1.2188 0.02627 0.02198 0.0150 0.0074 1.0000
12.250 1.2339 0.02697 0.02273 0.0163 0.0072 1.0000
12.500 1.2450 0.02794 0.02378 0.0181 0.0070 1.0000
12.750 1.2438 0.02969 0.02569 0.0214 0.0067 1.0000
13.000 1.2251 0.03316 0.02946 0.0255 0.0065 1.0000
13.250 1.2145 0.03616 0.03270 0.0279 0.0064 1.0000
13.500 1.2133 0.03841 0.03511 0.0290 0.0064 1.0000
13.750 1.2079 0.04133 0.03820 0.0296 0.0063 1.0000
14.000 1.2001 0.04484 0.04188 0.0293 0.0063 1.0000
14.250 1.1873 0.04947 0.04670 0.0279 0.0062 1.0000
14.500 1.1696 0.05550 0.05293 0.0248 0.0062 1.0000
14.750 1.1472 0.06339 0.06101 0.0197 0.0062 1.0000
15.000 1.1153 0.07453 0.07236 0.0119 0.0063 1.0000
15.250 1.0513 0.09368 0.09174 0.0000 0.0064 1.0000
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Polar data table (+)
Polar graphs
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