NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 200,000 Max Cl/Cd: 59.1 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc10n1-il-200000-n5.txt Download as CSV file: xf-rc10n1-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6045 0.08237 0.07922 -0.0106 1.0000 0.0230
-8.500 -0.6069 0.07862 0.07546 -0.0118 1.0000 0.0234
-8.250 -0.6076 0.07492 0.07174 -0.0130 1.0000 0.0238
-8.000 -0.6072 0.07052 0.06727 -0.0149 1.0000 0.0241
-7.750 -0.6058 0.06561 0.06222 -0.0165 1.0000 0.0240
-7.500 -0.5990 0.06063 0.05706 -0.0173 1.0000 0.0206
-7.250 -0.5896 0.05573 0.05193 -0.0188 0.9266 0.0193
-7.000 -0.5823 0.05100 0.04687 -0.0182 0.8992 0.0184
-6.750 -0.5734 0.04624 0.04173 -0.0170 0.8816 0.0177
-6.500 -0.5616 0.04181 0.03687 -0.0156 0.8687 0.0173
-6.000 -0.5280 0.03633 0.03076 -0.0133 0.8486 0.0188
-5.750 -0.5091 0.03352 0.02755 -0.0119 0.8406 0.0195
-5.500 -0.4894 0.03053 0.02415 -0.0106 0.8330 0.0193
-5.250 -0.4680 0.02785 0.02103 -0.0093 0.8263 0.0193
-5.000 -0.4449 0.02552 0.01830 -0.0082 0.8196 0.0194
-4.750 -0.4208 0.02356 0.01597 -0.0072 0.8139 0.0198
-4.500 -0.3953 0.02189 0.01399 -0.0065 0.8076 0.0203
-4.250 -0.3694 0.02092 0.01273 -0.0059 0.8021 0.0215
-4.000 -0.3428 0.02005 0.01166 -0.0054 0.7963 0.0222
-3.750 -0.3170 0.01849 0.00991 -0.0049 0.7908 0.0225
-3.500 -0.2918 0.01724 0.00855 -0.0043 0.7860 0.0230
-3.250 -0.2664 0.01627 0.00755 -0.0039 0.7803 0.0237
-3.000 -0.2414 0.01554 0.00677 -0.0033 0.7752 0.0247
-2.750 -0.2163 0.01493 0.00610 -0.0028 0.7702 0.0260
-2.500 -0.1909 0.01438 0.00550 -0.0023 0.7646 0.0278
-2.250 -0.1652 0.01398 0.00501 -0.0019 0.7600 0.0305
-2.000 -0.1397 0.01349 0.00449 -0.0015 0.7548 0.0354
-1.750 -0.1133 0.01316 0.00409 -0.0012 0.7494 0.0414
-1.500 -0.0873 0.01283 0.00371 -0.0008 0.7450 0.0552
-1.250 -0.0670 0.01147 0.00329 0.0000 0.7396 0.2991
-1.000 -0.0558 0.00985 0.00323 0.0032 0.7344 0.6940
-0.750 -0.0267 0.00941 0.00337 0.0041 0.7304 0.8418
-0.500 0.0214 0.00960 0.00361 0.0009 0.7249 0.9345
-0.250 0.0617 0.00973 0.00364 -0.0014 0.7195 0.9581
0.000 0.0992 0.00978 0.00356 -0.0035 0.7145 0.9648
0.250 0.1337 0.00981 0.00353 -0.0050 0.7081 0.9734
0.500 0.1704 0.00984 0.00346 -0.0070 0.7026 0.9798
0.750 0.2059 0.00985 0.00343 -0.0087 0.6956 0.9865
1.000 0.2422 0.00986 0.00337 -0.0107 0.6890 0.9923
1.250 0.2789 0.00985 0.00334 -0.0128 0.6816 0.9974
1.500 0.3106 0.00985 0.00329 -0.0137 0.6745 1.0000
1.750 0.3368 0.00986 0.00329 -0.0136 0.6665 1.0000
2.000 0.3626 0.00986 0.00326 -0.0132 0.6592 1.0000
2.250 0.3888 0.00988 0.00330 -0.0131 0.6505 1.0000
2.500 0.4147 0.00991 0.00332 -0.0127 0.6428 1.0000
2.750 0.4407 0.00994 0.00336 -0.0125 0.6334 1.0000
3.000 0.4666 0.00998 0.00341 -0.0121 0.6231 1.0000
3.250 0.4920 0.00999 0.00341 -0.0116 0.6066 1.0000
3.500 0.5171 0.01002 0.00341 -0.0111 0.5851 1.0000
3.750 0.5425 0.01009 0.00345 -0.0106 0.5641 1.0000
4.000 0.5677 0.01019 0.00352 -0.0101 0.5419 1.0000
4.250 0.5927 0.01033 0.00364 -0.0096 0.5144 1.0000
4.500 0.6173 0.01053 0.00375 -0.0091 0.4781 1.0000
4.750 0.6412 0.01085 0.00391 -0.0085 0.4321 1.0000
5.000 0.6641 0.01133 0.00416 -0.0079 0.3759 1.0000
5.250 0.6858 0.01201 0.00454 -0.0073 0.3088 1.0000
5.500 0.7068 0.01279 0.00504 -0.0067 0.2415 1.0000
5.750 0.7275 0.01360 0.00556 -0.0060 0.1814 1.0000
6.000 0.7485 0.01433 0.00610 -0.0053 0.1414 1.0000
6.250 0.7696 0.01501 0.00666 -0.0046 0.1130 1.0000
6.500 0.7909 0.01564 0.00725 -0.0039 0.0934 1.0000
6.750 0.8122 0.01624 0.00784 -0.0031 0.0791 1.0000
7.000 0.8330 0.01687 0.00846 -0.0023 0.0671 1.0000
7.250 0.8533 0.01757 0.00913 -0.0014 0.0575 1.0000
7.500 0.8742 0.01818 0.00981 -0.0005 0.0506 1.0000
7.750 0.8936 0.01895 0.01062 0.0004 0.0445 1.0000
8.000 0.9138 0.01961 0.01138 0.0013 0.0396 1.0000
8.250 0.9315 0.02056 0.01234 0.0024 0.0358 1.0000
8.500 0.9497 0.02143 0.01333 0.0035 0.0330 1.0000
8.750 0.9685 0.02223 0.01423 0.0045 0.0300 1.0000
9.000 0.9858 0.02315 0.01519 0.0056 0.0272 1.0000
9.250 0.9994 0.02453 0.01666 0.0071 0.0254 1.0000
9.500 1.0162 0.02556 0.01786 0.0082 0.0240 1.0000
9.750 1.0324 0.02661 0.01904 0.0094 0.0221 1.0000
10.000 1.0481 0.02759 0.02013 0.0105 0.0204 1.0000
10.250 1.0614 0.02879 0.02142 0.0117 0.0191 1.0000
10.500 1.0696 0.03070 0.02341 0.0134 0.0181 1.0000
10.750 1.0800 0.03223 0.02513 0.0150 0.0175 1.0000
11.000 1.0892 0.03388 0.02700 0.0166 0.0168 1.0000
11.250 1.0970 0.03563 0.02899 0.0180 0.0159 1.0000
11.500 1.1035 0.03737 0.03092 0.0191 0.0150 1.0000
11.750 1.1091 0.03914 0.03285 0.0200 0.0142 1.0000
12.000 1.1135 0.04102 0.03486 0.0206 0.0136 1.0000
12.250 1.1155 0.04328 0.03726 0.0211 0.0132 1.0000
12.500 1.1145 0.04601 0.04015 0.0214 0.0128 1.0000
12.750 1.1093 0.04939 0.04372 0.0216 0.0125 1.0000
13.000 1.0993 0.05354 0.04808 0.0215 0.0122 1.0000
13.250 1.0906 0.05760 0.05239 0.0207 0.0121 1.0000
13.500 1.0791 0.06221 0.05725 0.0194 0.0120 1.0000
13.750 1.0651 0.06744 0.06272 0.0175 0.0119 1.0000
14.000 1.0487 0.07336 0.06886 0.0148 0.0119 1.0000
14.250 1.0305 0.08004 0.07575 0.0113 0.0119 1.0000
14.500 1.0105 0.08762 0.08352 0.0070 0.0119 1.0000
14.750 0.9890 0.09634 0.09241 0.0017 0.0120 1.0000
15.000 0.9654 0.10624 0.10246 -0.0043 0.0121 1.0000
15.250 0.9390 0.11761 0.11394 -0.0110 0.0122 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il)