NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 50,000 Max Cl/Cd: 33.26 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc10n1-il-50000.txt Download as CSV file: xf-rc10n1-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-10(N)1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5677 0.11553 0.10924 0.0180 1.0000 0.2478
-9.000 -0.5676 0.11227 0.10604 0.0176 1.0000 0.2629
-8.750 -0.5726 0.10937 0.10322 0.0168 1.0000 0.2784
-8.500 -0.5489 0.10488 0.09869 0.0196 1.0000 0.3042
-8.250 -0.5396 0.10157 0.09543 0.0212 1.0000 0.3322
-8.000 -0.5369 0.09890 0.09281 0.0228 1.0000 0.3635
-7.750 -0.5159 0.09535 0.08927 0.0260 1.0000 0.4038
-7.500 -0.5000 0.09210 0.08605 0.0286 1.0000 0.4433
-7.250 -0.4826 0.08894 0.08291 0.0312 1.0000 0.4862
-7.000 -0.4627 0.08570 0.07965 0.0336 1.0000 0.5315
-6.750 -0.4428 0.08237 0.07634 0.0356 1.0000 0.5760
-5.250 -0.5164 0.04771 0.04091 -0.0150 1.0000 0.2478
-5.000 -0.4982 0.04381 0.03613 -0.0151 1.0000 0.1886
-4.750 -0.4836 0.04118 0.03280 -0.0130 1.0000 0.1608
-4.500 -0.4684 0.03874 0.02987 -0.0108 1.0000 0.1455
-4.250 -0.4516 0.03690 0.02736 -0.0085 1.0000 0.1335
-4.000 -0.4341 0.03454 0.02477 -0.0069 1.0000 0.1295
-3.750 -0.4151 0.03272 0.02260 -0.0054 1.0000 0.1282
-3.500 -0.3943 0.03104 0.02059 -0.0041 1.0000 0.1275
-3.250 -0.3713 0.02949 0.01870 -0.0029 1.0000 0.1256
-3.000 -0.3468 0.02808 0.01703 -0.0019 1.0000 0.1249
-2.750 -0.3210 0.02681 0.01562 -0.0012 1.0000 0.1264
-2.500 -0.2955 0.02586 0.01448 -0.0004 1.0000 0.1332
-2.250 -0.2689 0.02475 0.01339 -0.0001 1.0000 0.1434
-2.000 -0.1087 0.01949 0.01136 -0.0167 1.0000 1.0000
-1.750 -0.1120 0.01956 0.01110 -0.0128 1.0000 1.0000
-1.500 -0.1107 0.01967 0.01090 -0.0095 1.0000 1.0000
-1.250 -0.1058 0.01982 0.01076 -0.0067 1.0000 1.0000
-1.000 -0.0983 0.02001 0.01068 -0.0044 1.0000 1.0000
-0.750 -0.0886 0.02025 0.01067 -0.0024 1.0000 1.0000
-0.500 -0.0770 0.02053 0.01072 -0.0008 1.0000 1.0000
-0.250 -0.0638 0.02085 0.01081 0.0006 1.0000 1.0000
0.000 -0.0495 0.02122 0.01098 0.0017 1.0000 1.0000
0.250 -0.0344 0.02162 0.01121 0.0026 1.0000 1.0000
0.500 -0.0185 0.02206 0.01150 0.0034 1.0000 1.0000
0.750 0.0179 0.02279 0.01205 0.0001 0.9932 1.0000
1.000 0.0610 0.02365 0.01278 -0.0044 0.9834 1.0000
1.250 0.0997 0.02445 0.01349 -0.0080 0.9728 1.0000
1.500 0.1380 0.02529 0.01426 -0.0115 0.9619 1.0000
1.750 0.1772 0.02618 0.01511 -0.0151 0.9506 1.0000
2.000 0.2178 0.02711 0.01603 -0.0188 0.9388 1.0000
2.250 0.2593 0.02805 0.01700 -0.0226 0.9264 1.0000
2.500 0.2956 0.02895 0.01795 -0.0253 0.9130 1.0000
2.750 0.3306 0.02986 0.01893 -0.0277 0.8990 1.0000
3.000 0.3675 0.03079 0.01998 -0.0303 0.8842 1.0000
3.250 0.4012 0.03172 0.02101 -0.0322 0.8686 1.0000
3.500 0.4292 0.03266 0.02206 -0.0330 0.8519 1.0000
3.750 0.4599 0.03360 0.02314 -0.0342 0.8341 1.0000
4.000 0.4999 0.03443 0.02421 -0.0364 0.8155 1.0000
4.250 0.5373 0.03517 0.02515 -0.0378 0.7958 1.0000
4.500 0.5662 0.03583 0.02601 -0.0375 0.7737 1.0000
4.750 0.5994 0.03611 0.02656 -0.0369 0.7492 1.0000
5.000 0.6367 0.03564 0.02636 -0.0354 0.7218 1.0000
5.250 0.6714 0.03460 0.02558 -0.0325 0.6931 1.0000
5.500 0.7028 0.03315 0.02443 -0.0286 0.6640 1.0000
5.750 0.7319 0.03125 0.02279 -0.0239 0.6327 1.0000
6.000 0.7631 0.02838 0.02011 -0.0182 0.5980 1.0000
6.250 0.7855 0.02605 0.01794 -0.0124 0.5456 1.0000
6.500 0.8025 0.02413 0.01572 -0.0058 0.4471 1.0000
6.750 0.8097 0.02545 0.01580 -0.0007 0.3275 1.0000
7.000 0.8221 0.02765 0.01737 0.0019 0.2596 1.0000
7.250 0.8406 0.02973 0.01912 0.0036 0.2180 1.0000
7.500 0.8636 0.03191 0.02114 0.0049 0.1898 1.0000
7.750 0.8869 0.03426 0.02342 0.0059 0.1689 1.0000
8.000 0.9097 0.03688 0.02623 0.0069 0.1542 1.0000
8.250 0.9319 0.03976 0.02921 0.0078 0.1428 1.0000
8.500 0.9503 0.04247 0.03221 0.0089 0.1325 1.0000
8.750 0.9658 0.04628 0.03648 0.0102 0.1285 1.0000
9.000 0.9741 0.05001 0.04083 0.0118 0.1247 1.0000
9.250 0.9927 0.05343 0.04418 0.0125 0.1174 1.0000
9.500 0.9953 0.05780 0.04905 0.0140 0.1161 1.0000
9.750 0.9953 0.06249 0.05422 0.0153 0.1163 1.0000
10.000 0.9946 0.06752 0.05957 0.0163 0.1168 1.0000
10.250 0.9475 0.07252 0.06525 0.0174 0.1214 1.0000
10.500 0.9123 0.07803 0.07093 0.0170 0.1243 1.0000
10.750 0.8849 0.08425 0.07722 0.0148 0.1270 1.0000
11.000 0.8710 0.09065 0.08363 0.0126 0.1293 1.0000
11.250 0.8180 0.10458 0.09749 0.0016 0.1431 1.0000
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Polar data table (+)
Polar graphs
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