NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Reynolds number: 500,000 Max Cl/Cd: 62.41 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc1064c-il-500000.txt Download as CSV file: xf-rc1064c-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC10-64C 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4905 0.08491 0.08267 -0.0304 1.0000 0.0315
-10.250 -0.4987 0.07966 0.07743 -0.0321 1.0000 0.0318
-10.000 -0.5117 0.07348 0.07127 -0.0346 1.0000 0.0321
-9.750 -0.5550 0.06100 0.05880 -0.0437 1.0000 0.0317
-9.500 -0.6126 0.05177 0.04944 -0.0459 1.0000 0.0311
-9.250 -0.7867 0.04435 0.04110 -0.0315 1.0000 0.0223
-9.000 -0.8034 0.03962 0.03606 -0.0265 1.0000 0.0217
-8.750 -0.8179 0.03476 0.03076 -0.0210 1.0000 0.0213
-8.500 -0.8231 0.03078 0.02632 -0.0162 1.0000 0.0212
-8.250 -0.8188 0.02784 0.02301 -0.0126 1.0000 0.0213
-8.000 -0.8088 0.02559 0.02043 -0.0097 1.0000 0.0216
-7.750 -0.7945 0.02408 0.01865 -0.0075 1.0000 0.0222
-7.500 -0.7779 0.02306 0.01741 -0.0055 1.0000 0.0227
-7.250 -0.7603 0.02231 0.01647 -0.0036 1.0000 0.0231
-7.000 -0.7412 0.01966 0.01352 -0.0025 0.9996 0.0235
-6.750 -0.7116 0.01797 0.01170 -0.0033 0.9981 0.0243
-6.500 -0.6797 0.01703 0.01069 -0.0045 0.9964 0.0251
-6.250 -0.6468 0.01627 0.00986 -0.0059 0.9948 0.0260
-6.000 -0.6149 0.01564 0.00918 -0.0070 0.9927 0.0273
-5.750 -0.5828 0.01499 0.00846 -0.0081 0.9900 0.0284
-5.500 -0.5494 0.01440 0.00779 -0.0094 0.9876 0.0292
-5.250 -0.5173 0.01328 0.00661 -0.0107 0.9856 0.0308
-5.000 -0.4865 0.01268 0.00601 -0.0115 0.9814 0.0328
-4.750 -0.4526 0.01222 0.00553 -0.0130 0.9782 0.0353
-4.500 -0.4177 0.01171 0.00498 -0.0146 0.9759 0.0387
-4.250 -0.3862 0.01122 0.00452 -0.0155 0.9723 0.0444
-4.000 -0.3532 0.01074 0.00409 -0.0168 0.9693 0.0556
-3.750 -0.3186 0.01029 0.00372 -0.0184 0.9673 0.0762
-3.500 -0.2825 0.00989 0.00344 -0.0204 0.9657 0.1015
-3.250 -0.2475 0.00947 0.00319 -0.0222 0.9633 0.1378
-3.000 -0.2106 0.00940 0.00321 -0.0242 0.9606 0.1834
-2.750 -0.1739 0.00909 0.00292 -0.0263 0.9580 0.2006
-2.250 -0.1051 0.00804 0.00239 -0.0298 0.9498 0.3215
-2.000 -0.0783 0.00722 0.00225 -0.0301 0.9439 0.5132
-1.750 -0.0606 0.00676 0.00237 -0.0280 0.9352 0.6698
-1.500 -0.0397 0.00671 0.00256 -0.0262 0.9269 0.7380
-1.250 -0.0176 0.00674 0.00270 -0.0247 0.9187 0.7765
-1.000 0.0053 0.00669 0.00274 -0.0234 0.9102 0.8067
-0.750 0.0279 0.00655 0.00271 -0.0219 0.8993 0.8365
-0.500 0.0516 0.00643 0.00268 -0.0207 0.8857 0.8676
-0.250 0.0770 0.00633 0.00262 -0.0197 0.8763 0.8916
0.000 0.1034 0.00622 0.00253 -0.0188 0.8687 0.9082
0.250 0.1311 0.00608 0.00236 -0.0182 0.8614 0.9185
0.500 0.1593 0.00605 0.00224 -0.0179 0.8503 0.9255
0.750 0.1880 0.00615 0.00226 -0.0178 0.8355 0.9314
1.000 0.2134 0.00624 0.00230 -0.0170 0.8194 0.9389
1.250 0.2432 0.00632 0.00236 -0.0172 0.8037 0.9431
1.500 0.2713 0.00638 0.00239 -0.0171 0.7861 0.9487
1.750 0.2974 0.00645 0.00241 -0.0165 0.7661 0.9548
2.000 0.3295 0.00656 0.00247 -0.0172 0.7464 0.9580
2.250 0.3591 0.00671 0.00250 -0.0175 0.7139 0.9620
2.500 0.3844 0.00689 0.00255 -0.0167 0.6789 0.9677
2.750 0.4160 0.00708 0.00262 -0.0175 0.6417 0.9704
3.000 0.4489 0.00735 0.00271 -0.0187 0.5920 0.9725
3.250 0.4799 0.00769 0.00284 -0.0195 0.5353 0.9754
3.500 0.5071 0.00817 0.00299 -0.0195 0.4575 0.9795
3.750 0.5326 0.00886 0.00323 -0.0195 0.3505 0.9832
4.000 0.5624 0.00965 0.00353 -0.0206 0.2435 0.9854
4.250 0.5927 0.01033 0.00386 -0.0217 0.1668 0.9881
4.500 0.6231 0.01094 0.00419 -0.0227 0.1103 0.9912
4.750 0.6536 0.01155 0.00458 -0.0236 0.0753 0.9941
5.000 0.6873 0.01202 0.00499 -0.0252 0.0622 0.9961
5.250 0.7215 0.01237 0.00535 -0.0269 0.0563 0.9983
5.500 0.7513 0.01296 0.00593 -0.0276 0.0507 1.0000
5.750 0.7710 0.01322 0.00624 -0.0261 0.0485 1.0000
6.000 0.7897 0.01357 0.00662 -0.0244 0.0462 1.0000
6.250 0.8076 0.01398 0.00704 -0.0226 0.0441 1.0000
6.500 0.8223 0.01468 0.00776 -0.0202 0.0417 1.0000
6.750 0.8378 0.01533 0.00847 -0.0179 0.0402 1.0000
7.000 0.8561 0.01568 0.00887 -0.0161 0.0390 1.0000
7.250 0.8736 0.01613 0.00938 -0.0142 0.0376 1.0000
7.500 0.8909 0.01663 0.00992 -0.0123 0.0363 1.0000
7.750 0.9080 0.01716 0.01048 -0.0104 0.0350 1.0000
8.000 0.9244 0.01783 0.01116 -0.0084 0.0337 1.0000
8.250 0.9378 0.01949 0.01289 -0.0061 0.0320 1.0000
8.500 0.9569 0.01987 0.01336 -0.0045 0.0313 1.0000
8.750 0.9754 0.02049 0.01406 -0.0029 0.0305 1.0000
9.000 0.9939 0.02115 0.01481 -0.0014 0.0294 1.0000
9.250 1.0123 0.02173 0.01546 0.0001 0.0283 1.0000
9.500 1.0306 0.02226 0.01603 0.0015 0.0273 1.0000
9.750 1.0482 0.02306 0.01685 0.0029 0.0263 1.0000
10.000 1.0621 0.02573 0.01969 0.0045 0.0251 1.0000
10.250 1.0797 0.02612 0.02021 0.0060 0.0245 1.0000
10.500 1.0964 0.02682 0.02105 0.0076 0.0237 1.0000
10.750 1.1124 0.02760 0.02196 0.0091 0.0227 1.0000
11.000 1.1273 0.02842 0.02288 0.0107 0.0219 1.0000
11.250 1.1415 0.02918 0.02371 0.0124 0.0213 1.0000
11.500 1.1545 0.02998 0.02456 0.0141 0.0207 1.0000
11.750 1.1597 0.03208 0.02677 0.0166 0.0199 1.0000
12.000 1.1596 0.03407 0.02902 0.0199 0.0194 1.0000
12.250 1.1640 0.03518 0.03031 0.0225 0.0189 1.0000
12.500 1.1664 0.03666 0.03199 0.0249 0.0184 1.0000
12.750 1.1679 0.03827 0.03376 0.0271 0.0178 1.0000
13.000 1.1689 0.03997 0.03559 0.0290 0.0174 1.0000
13.250 1.1684 0.04185 0.03762 0.0305 0.0170 1.0000
13.500 1.1676 0.04388 0.03977 0.0316 0.0167 1.0000
13.750 1.1678 0.04592 0.04191 0.0323 0.0164 1.0000
14.000 1.1679 0.04812 0.04418 0.0326 0.0161 1.0000
14.250 1.1619 0.05127 0.04744 0.0325 0.0158 1.0000
14.500 1.1493 0.05554 0.05187 0.0317 0.0157 1.0000
14.750 1.1299 0.06118 0.05770 0.0298 0.0155 1.0000
15.000 1.1074 0.06788 0.06461 0.0265 0.0155 1.0000
15.250 1.0835 0.07561 0.07255 0.0220 0.0155 1.0000
15.500 1.0558 0.08512 0.08228 0.0157 0.0156 1.0000
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