NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 100,000 Max Cl/Cd: 46.6 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc510-il-100000.txt Download as CSV file: xf-rc510-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3954 0.09636 0.09208 -0.0069 1.0000 0.0964
-8.750 -0.3922 0.09259 0.08834 -0.0078 1.0000 0.0991
-8.500 -0.5443 0.09719 0.09277 -0.0136 1.0000 0.0923
-8.250 -0.5092 0.09330 0.08891 -0.0042 1.0000 0.0953
-8.000 -0.5029 0.08959 0.08524 -0.0057 1.0000 0.0980
-7.750 -0.5026 0.08546 0.08116 -0.0096 1.0000 0.1008
-7.500 -0.5110 0.08114 0.07674 -0.0190 1.0000 0.1054
-7.250 -0.5137 0.07646 0.07191 -0.0237 1.0000 0.1073
-7.000 -0.4945 0.07187 0.06748 -0.0220 1.0000 0.1096
-6.750 -0.4812 0.06845 0.06409 -0.0226 1.0000 0.1134
-6.500 -0.4865 0.06807 0.06307 -0.0273 1.0000 0.1219
-6.250 -0.4757 0.06185 0.05717 -0.0261 1.0000 0.1236
-6.000 -0.4804 0.05979 0.05521 -0.0226 1.0000 0.1253
-5.750 -0.4858 0.05810 0.05353 -0.0193 1.0000 0.1277
-5.500 -0.4759 0.05574 0.05074 -0.0210 0.9951 0.1385
-5.250 -0.4471 0.05171 0.04682 -0.0231 0.9889 0.1468
-5.000 -0.4164 0.04803 0.04294 -0.0266 0.9822 0.1610
-4.750 -0.3878 0.04477 0.03950 -0.0292 0.9744 0.1764
-4.500 -0.3579 0.04174 0.03622 -0.0320 0.9681 0.2018
-4.250 -0.3055 0.03391 0.02669 -0.0331 0.9614 0.0887
-4.000 -0.2671 0.03041 0.02271 -0.0353 0.9569 0.0843
-3.750 -0.2356 0.02827 0.02018 -0.0359 0.9498 0.0854
-3.500 -0.1986 0.02600 0.01746 -0.0372 0.9442 0.0856
-3.250 -0.1620 0.02429 0.01536 -0.0384 0.9389 0.0889
-3.000 -0.1318 0.02303 0.01396 -0.0389 0.9320 0.0966
-2.750 -0.0945 0.02194 0.01264 -0.0402 0.9271 0.1070
-2.500 -0.0665 0.02114 0.01182 -0.0403 0.9199 0.1219
-2.250 -0.0358 0.02003 0.01091 -0.0409 0.9141 0.1428
-2.000 -0.0073 0.01924 0.01022 -0.0410 0.9088 0.1696
-1.750 0.0150 0.01857 0.00975 -0.0402 0.9016 0.2009
-1.500 0.1686 0.01518 0.00894 -0.0599 0.9124 1.0000
-1.250 0.1916 0.01537 0.00894 -0.0597 0.9038 1.0000
-1.000 0.2161 0.01556 0.00896 -0.0595 0.8968 1.0000
-0.750 0.2386 0.01582 0.00909 -0.0591 0.8888 1.0000
-0.500 0.2620 0.01606 0.00920 -0.0585 0.8817 1.0000
-0.250 0.2829 0.01634 0.00938 -0.0575 0.8717 1.0000
0.000 0.3028 0.01654 0.00947 -0.0557 0.8601 1.0000
0.250 0.3215 0.01668 0.00950 -0.0533 0.8475 1.0000
0.500 0.3397 0.01677 0.00949 -0.0506 0.8347 1.0000
0.750 0.3586 0.01690 0.00954 -0.0483 0.8212 1.0000
1.000 0.3779 0.01701 0.00959 -0.0462 0.8081 1.0000
1.250 0.3982 0.01716 0.00968 -0.0443 0.7964 1.0000
1.500 0.4190 0.01726 0.00972 -0.0424 0.7869 1.0000
1.750 0.4403 0.01737 0.00979 -0.0406 0.7760 1.0000
2.000 0.4617 0.01750 0.00991 -0.0391 0.7643 1.0000
2.250 0.4830 0.01754 0.00992 -0.0372 0.7529 1.0000
2.500 0.5044 0.01745 0.00978 -0.0350 0.7421 1.0000
2.750 0.5260 0.01737 0.00970 -0.0330 0.7292 1.0000
3.000 0.5478 0.01724 0.00957 -0.0311 0.7149 1.0000
3.250 0.5698 0.01704 0.00936 -0.0291 0.7000 1.0000
3.500 0.5921 0.01672 0.00905 -0.0269 0.6839 1.0000
3.750 0.6146 0.01631 0.00861 -0.0247 0.6662 1.0000
4.000 0.6370 0.01595 0.00828 -0.0227 0.6416 1.0000
4.250 0.6594 0.01546 0.00778 -0.0205 0.6106 1.0000
4.500 0.6812 0.01503 0.00725 -0.0181 0.5612 1.0000
4.750 0.7008 0.01504 0.00686 -0.0157 0.4691 1.0000
5.000 0.7186 0.01585 0.00698 -0.0138 0.3947 1.0000
5.250 0.7386 0.01678 0.00753 -0.0125 0.3569 1.0000
5.500 0.7599 0.01767 0.00815 -0.0114 0.3331 1.0000
5.750 0.7824 0.01849 0.00882 -0.0106 0.3138 1.0000
6.000 0.8054 0.01931 0.00953 -0.0098 0.2980 1.0000
6.250 0.8289 0.02016 0.01029 -0.0092 0.2849 1.0000
6.500 0.8526 0.02105 0.01108 -0.0086 0.2731 1.0000
6.750 0.8762 0.02186 0.01189 -0.0080 0.2616 1.0000
7.000 0.8997 0.02274 0.01285 -0.0074 0.2513 1.0000
7.250 0.9238 0.02374 0.01380 -0.0069 0.2421 1.0000
7.500 0.9469 0.02459 0.01474 -0.0063 0.2324 1.0000
7.750 0.9701 0.02565 0.01589 -0.0057 0.2240 1.0000
8.000 0.9935 0.02662 0.01685 -0.0052 0.2156 1.0000
8.250 1.0151 0.02771 0.01817 -0.0045 0.2071 1.0000
8.500 1.0389 0.02882 0.01921 -0.0041 0.1997 1.0000
8.750 1.0588 0.02998 0.02070 -0.0032 0.1915 1.0000
9.000 1.0826 0.03119 0.02177 -0.0029 0.1841 1.0000
9.250 1.1001 0.03237 0.02341 -0.0017 0.1762 1.0000
9.500 1.1231 0.03359 0.02451 -0.0013 0.1690 1.0000
9.750 1.1385 0.03486 0.02621 0.0000 0.1610 1.0000
10.000 1.1602 0.03608 0.02736 0.0005 0.1541 1.0000
10.250 1.1736 0.03745 0.02916 0.0019 0.1463 1.0000
10.500 1.1935 0.03870 0.03038 0.0026 0.1396 1.0000
10.750 1.2046 0.04012 0.03217 0.0041 0.1319 1.0000
11.000 1.2217 0.04142 0.03348 0.0050 0.1253 1.0000
11.250 1.2303 0.04285 0.03523 0.0067 0.1180 1.0000
11.500 1.2443 0.04421 0.03662 0.0077 0.1117 1.0000
11.750 1.2483 0.04585 0.03860 0.0097 0.1052 1.0000
12.000 1.2614 0.04708 0.03979 0.0107 0.0991 1.0000
12.250 1.2539 0.04960 0.04273 0.0132 0.0944 1.0000
12.500 1.2744 0.04999 0.04285 0.0138 0.0877 1.0000
12.750 1.2530 0.05321 0.04654 0.0169 0.0857 1.0000
13.000 1.2330 0.05673 0.05039 0.0188 0.0836 1.0000
13.250 1.2301 0.05869 0.05243 0.0196 0.0797 1.0000
13.500 1.2365 0.06053 0.05418 0.0203 0.0755 1.0000
13.750 1.2093 0.06562 0.05959 0.0201 0.0750 1.0000
14.000 1.1798 0.07162 0.06588 0.0187 0.0750 1.0000
14.250 1.1481 0.07864 0.07312 0.0161 0.0753 1.0000
14.500 1.1148 0.08674 0.08140 0.0124 0.0759 1.0000
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Polar data table (+)
Polar graphs
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