NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 200,000 Max Cl/Cd: 54.75 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc510-il-200000-n5.txt Download as CSV file: xf-rc510-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(5)-10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5442 0.08809 0.08485 -0.0017 1.0000 0.0218
-8.500 -0.5468 0.08261 0.07941 -0.0068 1.0000 0.0217
-8.250 -0.5517 0.07664 0.07345 -0.0130 0.9919 0.0216
-8.000 -0.5511 0.07073 0.06738 -0.0180 0.9071 0.0217
-7.500 -0.5486 0.06028 0.05652 -0.0219 0.8678 0.0218
-7.250 -0.5430 0.05513 0.05112 -0.0228 0.8561 0.0215
-7.000 -0.5358 0.04970 0.04539 -0.0232 0.8465 0.0212
-6.750 -0.5273 0.04380 0.03908 -0.0229 0.8381 0.0209
-6.500 -0.5174 0.03722 0.03198 -0.0219 0.8306 0.0207
-6.250 -0.5062 0.03086 0.02489 -0.0203 0.8239 0.0208
-6.000 -0.4883 0.02702 0.02045 -0.0190 0.8176 0.0212
-5.750 -0.4669 0.02442 0.01735 -0.0180 0.8114 0.0217
-5.500 -0.4436 0.02266 0.01514 -0.0171 0.8061 0.0227
-5.250 -0.4182 0.02130 0.01341 -0.0165 0.7998 0.0235
-5.000 -0.3937 0.01999 0.01193 -0.0159 0.7942 0.0242
-4.750 -0.3682 0.01904 0.01086 -0.0155 0.7891 0.0249
-4.500 -0.3420 0.01819 0.00991 -0.0151 0.7836 0.0257
-4.250 -0.3161 0.01742 0.00902 -0.0146 0.7788 0.0266
-4.000 -0.2900 0.01671 0.00817 -0.0141 0.7740 0.0278
-3.750 -0.2637 0.01605 0.00742 -0.0137 0.7685 0.0292
-3.500 -0.2380 0.01546 0.00680 -0.0132 0.7637 0.0313
-3.250 -0.2117 0.01508 0.00639 -0.0129 0.7594 0.0344
-3.000 -0.1851 0.01462 0.00586 -0.0125 0.7544 0.0376
-2.750 -0.1590 0.01417 0.00542 -0.0121 0.7498 0.0425
-2.500 -0.1329 0.01382 0.00503 -0.0117 0.7460 0.0493
-2.250 -0.1058 0.01353 0.00479 -0.0115 0.7407 0.0601
-2.000 -0.0789 0.01330 0.00460 -0.0113 0.7360 0.0742
-1.750 -0.0521 0.01310 0.00443 -0.0110 0.7323 0.0897
-1.500 -0.0250 0.01291 0.00426 -0.0109 0.7281 0.1055
-1.250 0.0023 0.01273 0.00411 -0.0107 0.7232 0.1226
-1.000 0.0293 0.01254 0.00395 -0.0105 0.7189 0.1424
-0.750 0.0557 0.01231 0.00377 -0.0102 0.7154 0.1715
-0.500 0.0729 0.01079 0.00357 -0.0088 0.7088 0.5267
-0.250 0.0910 0.00988 0.00352 -0.0061 0.7016 0.7550
0.000 0.1414 0.00970 0.00368 -0.0095 0.6902 0.9209
0.250 0.1848 0.00977 0.00365 -0.0123 0.6812 0.9565
0.500 0.2289 0.00983 0.00361 -0.0154 0.6718 0.9770
0.750 0.2733 0.00985 0.00355 -0.0188 0.6617 0.9912
1.000 0.3159 0.00984 0.00343 -0.0219 0.6513 1.0000
1.250 0.3418 0.00982 0.00333 -0.0214 0.6404 1.0000
1.500 0.3677 0.00981 0.00327 -0.0210 0.6279 1.0000
1.750 0.3937 0.00982 0.00322 -0.0206 0.6158 1.0000
2.000 0.4196 0.00985 0.00319 -0.0202 0.6033 1.0000
2.250 0.4455 0.00988 0.00317 -0.0197 0.5891 1.0000
2.500 0.4713 0.00993 0.00316 -0.0193 0.5724 1.0000
2.750 0.4968 0.01001 0.00316 -0.0188 0.5508 1.0000
3.000 0.5221 0.01012 0.00316 -0.0182 0.5205 1.0000
3.250 0.5468 0.01032 0.00317 -0.0176 0.4697 1.0000
3.500 0.5692 0.01092 0.00328 -0.0169 0.3774 1.0000
3.750 0.5918 0.01156 0.00358 -0.0163 0.3183 1.0000
4.000 0.6153 0.01203 0.00387 -0.0158 0.2862 1.0000
4.250 0.6390 0.01244 0.00415 -0.0153 0.2654 1.0000
4.500 0.6629 0.01281 0.00445 -0.0148 0.2490 1.0000
5.000 0.7105 0.01350 0.00503 -0.0137 0.2235 1.0000
5.250 0.7343 0.01384 0.00534 -0.0131 0.2143 1.0000
5.500 0.7576 0.01422 0.00569 -0.0124 0.2053 1.0000
5.750 0.7814 0.01454 0.00602 -0.0118 0.1970 1.0000
6.250 0.8280 0.01526 0.00675 -0.0106 0.1830 1.0000
6.500 0.8508 0.01567 0.00713 -0.0099 0.1763 1.0000
6.750 0.8741 0.01603 0.00753 -0.0092 0.1701 1.0000
7.000 0.8971 0.01642 0.00793 -0.0086 0.1640 1.0000
7.250 0.9196 0.01685 0.00838 -0.0079 0.1584 1.0000
7.500 0.9428 0.01722 0.00882 -0.0073 0.1524 1.0000
7.750 0.9648 0.01770 0.00928 -0.0066 0.1472 1.0000
8.000 0.9876 0.01811 0.00977 -0.0059 0.1420 1.0000
8.250 1.0100 0.01854 0.01027 -0.0053 0.1363 1.0000
8.500 1.0313 0.01907 0.01079 -0.0046 0.1313 1.0000
8.750 1.0539 0.01949 0.01133 -0.0039 0.1259 1.0000
9.000 1.0750 0.02000 0.01186 -0.0032 0.1202 1.0000
9.250 1.0964 0.02051 0.01246 -0.0025 0.1149 1.0000
9.500 1.1174 0.02102 0.01304 -0.0018 0.1092 1.0000
9.750 1.1372 0.02163 0.01368 -0.0010 0.1042 1.0000
10.000 1.1579 0.02216 0.01433 -0.0003 0.0984 1.0000
10.250 1.1761 0.02284 0.01503 0.0006 0.0931 1.0000
10.500 1.1957 0.02343 0.01575 0.0014 0.0875 1.0000
10.750 1.2129 0.02415 0.01650 0.0023 0.0818 1.0000
11.000 1.2303 0.02487 0.01733 0.0033 0.0763 1.0000
11.250 1.2456 0.02569 0.01820 0.0044 0.0707 1.0000
11.500 1.2592 0.02654 0.01917 0.0057 0.0656 1.0000
11.750 1.2706 0.02753 0.02020 0.0070 0.0602 1.0000
12.000 1.2817 0.02862 0.02138 0.0082 0.0552 1.0000
12.250 1.2913 0.02990 0.02270 0.0092 0.0500 1.0000
12.500 1.3009 0.03125 0.02416 0.0101 0.0455 1.0000
12.750 1.3084 0.03283 0.02579 0.0109 0.0412 1.0000
13.000 1.3150 0.03454 0.02761 0.0116 0.0378 1.0000
13.250 1.3207 0.03638 0.02957 0.0122 0.0345 1.0000
13.500 1.3223 0.03868 0.03194 0.0127 0.0320 1.0000
13.750 1.3255 0.04087 0.03428 0.0130 0.0297 1.0000
14.000 1.3263 0.04339 0.03693 0.0132 0.0276 1.0000
14.250 1.3232 0.04640 0.04004 0.0131 0.0259 1.0000
14.500 1.3183 0.04974 0.04351 0.0127 0.0246 1.0000
14.750 1.3141 0.05312 0.04707 0.0122 0.0232 1.0000
15.000 1.3070 0.05702 0.05113 0.0112 0.0221 1.0000
15.250 1.2966 0.06152 0.05578 0.0098 0.0212 1.0000
15.500 1.2830 0.06675 0.06115 0.0078 0.0206 1.0000
15.750 1.2664 0.07285 0.06741 0.0051 0.0202 1.0000
16.000 1.2475 0.07985 0.07457 0.0017 0.0200 1.0000
16.250 1.2262 0.08770 0.08261 -0.0023 0.0199 1.0000
16.500 1.2034 0.09620 0.09128 -0.0067 0.0199 1.0000
16.750 1.1799 0.10493 0.10017 -0.0112 0.0200 1.0000
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Polar data table (+)
Polar graphs
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