NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 500,000 Max Cl/Cd: 84.3 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc10n1-il-500000.txt Download as CSV file: xf-rc10n1-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5995 0.08705 0.08510 0.0029 1.0000 0.0158
-8.500 -0.6048 0.08065 0.07872 -0.0044 1.0000 0.0158
-8.250 -0.6120 0.07520 0.07323 -0.0090 1.0000 0.0159
-8.000 -0.6121 0.07027 0.06825 -0.0121 1.0000 0.0162
-7.750 -0.6067 0.06538 0.06322 -0.0151 0.9407 0.0167
-7.500 -0.6057 0.06148 0.05908 -0.0148 0.8974 0.0173
-7.250 -0.5986 0.05720 0.05453 -0.0148 0.8785 0.0183
-7.000 -0.5797 0.05412 0.05103 -0.0141 0.8655 0.0195
-6.500 -0.5701 0.04281 0.03910 -0.0124 0.8466 0.0204
-6.250 -0.5549 0.04022 0.03637 -0.0119 0.8392 0.0208
-6.000 -0.5374 0.03790 0.03390 -0.0114 0.8318 0.0215
-5.750 -0.5188 0.03565 0.03145 -0.0107 0.8255 0.0224
-5.500 -0.4982 0.03329 0.02884 -0.0099 0.8189 0.0240
-5.250 -0.4688 0.03379 0.02895 -0.0083 0.8132 0.0270
-5.000 -0.4555 0.02795 0.02258 -0.0066 0.8078 0.0280
-4.750 -0.4339 0.02550 0.02004 -0.0063 0.8022 0.0292
-4.500 -0.4104 0.02408 0.01848 -0.0059 0.7972 0.0307
-4.250 -0.3850 0.02301 0.01726 -0.0054 0.7917 0.0338
-4.000 -0.3554 0.02456 0.01856 -0.0046 0.7865 0.0371
-3.750 -0.3318 0.01692 0.01033 -0.0025 0.7827 0.0238
-3.500 -0.3047 0.01527 0.00849 -0.0019 0.7775 0.0233
-3.250 -0.2782 0.01419 0.00730 -0.0014 0.7728 0.0236
-3.000 -0.2518 0.01345 0.00648 -0.0010 0.7683 0.0244
-2.750 -0.2251 0.01286 0.00584 -0.0006 0.7631 0.0256
-2.500 -0.1993 0.01218 0.00509 -0.0001 0.7584 0.0263
-2.250 -0.1731 0.01175 0.00459 0.0003 0.7540 0.0270
-2.000 -0.1490 0.01082 0.00361 0.0011 0.7488 0.0294
-1.750 -0.1224 0.01047 0.00322 0.0014 0.7440 0.0327
-1.500 -0.0951 0.01029 0.00299 0.0015 0.7396 0.0371
-1.250 -0.0684 0.00988 0.00257 0.0018 0.7341 0.0495
-1.000 -0.0564 0.00748 0.00212 0.0039 0.7294 0.5962
-0.750 -0.0360 0.00688 0.00203 0.0057 0.7247 0.7359
-0.500 -0.0139 0.00641 0.00215 0.0077 0.7192 0.8894
-0.250 0.0216 0.00649 0.00225 0.0066 0.7142 0.9443
0.000 0.0649 0.00668 0.00238 0.0037 0.7087 0.9655
0.250 0.1171 0.00690 0.00253 -0.0013 0.7025 0.9811
0.500 0.1555 0.00696 0.00250 -0.0036 0.6972 0.9858
0.750 0.1948 0.00697 0.00247 -0.0062 0.6903 0.9903
1.000 0.2321 0.00700 0.00244 -0.0083 0.6843 0.9949
1.250 0.2718 0.00699 0.00241 -0.0110 0.6775 0.9983
1.500 0.3047 0.00697 0.00236 -0.0123 0.6711 1.0000
1.750 0.3317 0.00696 0.00233 -0.0123 0.6644 1.0000
2.000 0.3586 0.00694 0.00229 -0.0122 0.6572 1.0000
2.250 0.3856 0.00694 0.00228 -0.0121 0.6496 1.0000
2.500 0.4122 0.00692 0.00222 -0.0120 0.6382 1.0000
2.750 0.4387 0.00690 0.00217 -0.0117 0.6238 1.0000
3.000 0.4654 0.00691 0.00216 -0.0116 0.6117 1.0000
3.250 0.4919 0.00693 0.00218 -0.0114 0.5984 1.0000
3.500 0.5184 0.00697 0.00221 -0.0112 0.5841 1.0000
3.750 0.5448 0.00703 0.00225 -0.0110 0.5689 1.0000
4.000 0.5712 0.00710 0.00231 -0.0108 0.5493 1.0000
4.250 0.5972 0.00722 0.00239 -0.0106 0.5261 1.0000
4.500 0.6230 0.00739 0.00248 -0.0104 0.4918 1.0000
4.750 0.6482 0.00769 0.00261 -0.0101 0.4435 1.0000
5.000 0.6726 0.00815 0.00284 -0.0098 0.3853 1.0000
5.250 0.6963 0.00873 0.00316 -0.0095 0.3202 1.0000
5.500 0.7189 0.00944 0.00355 -0.0091 0.2482 1.0000
5.750 0.7407 0.01021 0.00399 -0.0086 0.1786 1.0000
6.000 0.7624 0.01093 0.00445 -0.0080 0.1298 1.0000
6.250 0.7848 0.01149 0.00489 -0.0074 0.1012 1.0000
6.500 0.8072 0.01200 0.00532 -0.0067 0.0814 1.0000
6.750 0.8294 0.01254 0.00579 -0.0060 0.0662 1.0000
7.000 0.8517 0.01303 0.00626 -0.0053 0.0561 1.0000
7.250 0.8738 0.01354 0.00680 -0.0045 0.0484 1.0000
7.500 0.8939 0.01429 0.00753 -0.0035 0.0409 1.0000
7.750 0.9165 0.01469 0.00799 -0.0028 0.0373 1.0000
8.000 0.9365 0.01538 0.00869 -0.0018 0.0329 1.0000
8.250 0.9557 0.01615 0.00954 -0.0007 0.0299 1.0000
8.500 0.9770 0.01667 0.01011 0.0002 0.0273 1.0000
8.750 0.9964 0.01740 0.01088 0.0012 0.0250 1.0000
9.000 1.0094 0.01888 0.01246 0.0030 0.0227 1.0000
9.250 1.0307 0.01937 0.01304 0.0038 0.0214 1.0000
9.500 1.0505 0.02002 0.01374 0.0046 0.0198 1.0000
9.750 1.0690 0.02079 0.01456 0.0056 0.0185 1.0000
10.000 1.0817 0.02225 0.01609 0.0072 0.0172 1.0000
10.250 1.0941 0.02389 0.01790 0.0088 0.0162 1.0000
10.500 1.1119 0.02464 0.01876 0.0098 0.0154 1.0000
10.750 1.1280 0.02551 0.01973 0.0108 0.0144 1.0000
11.000 1.1422 0.02643 0.02073 0.0120 0.0135 1.0000
11.250 1.1523 0.02753 0.02190 0.0136 0.0129 1.0000
11.500 1.1575 0.02932 0.02379 0.0154 0.0123 1.0000
11.750 1.1541 0.03274 0.02747 0.0177 0.0118 1.0000
12.000 1.1615 0.03425 0.02916 0.0187 0.0115 1.0000
12.250 1.1659 0.03626 0.03136 0.0197 0.0112 1.0000
12.500 1.1676 0.03863 0.03394 0.0205 0.0109 1.0000
12.750 1.1668 0.04133 0.03684 0.0211 0.0106 1.0000
13.000 1.1636 0.04435 0.04006 0.0214 0.0104 1.0000
13.250 1.1585 0.04765 0.04355 0.0213 0.0102 1.0000
13.500 1.1520 0.05122 0.04730 0.0209 0.0100 1.0000
13.750 1.1442 0.05511 0.05135 0.0200 0.0098 1.0000
14.000 1.1342 0.05946 0.05588 0.0187 0.0097 1.0000
14.250 1.1223 0.06433 0.06091 0.0168 0.0096 1.0000
14.500 1.1078 0.06988 0.06663 0.0143 0.0095 1.0000
14.750 1.0877 0.07683 0.07377 0.0108 0.0095 1.0000
15.000 1.0613 0.08569 0.08285 0.0059 0.0096 1.0000
15.250 1.0155 0.09993 0.09738 -0.0026 0.0101 1.0000
15.500 0.9686 0.11617 0.11381 -0.0122 0.0106 1.0000
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Polar data table (+)
Polar graphs
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