NASA/LANGLEY RC-10(B)3 AIRFOIL (rc10b3-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY RC-10(B)3 AIRFOIL (rc10b3-il) Reynolds number: 100,000 Max Cl/Cd: 49.28 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc10b3-il-100000.txt Download as CSV file: xf-rc10b3-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-10(B)3 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3883 0.10112 0.09629 -0.0219 1.0000 0.1052
-9.500 -0.4009 0.09747 0.09270 -0.0248 1.0000 0.1088
-9.250 -0.4256 0.09356 0.08886 -0.0297 1.0000 0.1100
-9.000 -0.4087 0.08873 0.08404 -0.0271 1.0000 0.1131
-8.750 -0.3975 0.08551 0.08082 -0.0259 1.0000 0.1177
-8.500 -0.4055 0.08152 0.07687 -0.0277 1.0000 0.1218
-8.250 -0.4326 0.07677 0.07221 -0.0328 1.0000 0.1241
-8.000 -0.4668 0.07317 0.06863 -0.0341 1.0000 0.1248
-7.750 -0.5014 0.07042 0.06582 -0.0328 1.0000 0.1254
-7.500 -0.4523 0.06504 0.06059 -0.0309 1.0000 0.1316
-7.250 -0.4653 0.06199 0.05758 -0.0291 1.0000 0.1350
-7.000 -0.4927 0.05953 0.05511 -0.0262 1.0000 0.1386
-6.750 -0.5413 0.05925 0.05453 -0.0207 1.0000 0.1409
-6.500 -0.5199 0.05341 0.04896 -0.0198 1.0000 0.1441
-6.250 -0.5200 0.05093 0.04651 -0.0167 1.0000 0.1481
-6.000 -0.5524 0.05007 0.04525 -0.0122 1.0000 0.1568
-5.750 -0.5636 0.05617 0.05099 -0.0115 1.0000 0.1604
-5.250 -0.5588 0.05164 0.04634 -0.0058 1.0000 0.1776
-5.000 -0.5567 0.04953 0.04410 -0.0030 1.0000 0.1910
-4.750 -0.5518 0.04749 0.04198 -0.0003 1.0000 0.2061
-4.500 -0.5448 0.04537 0.03982 0.0023 1.0000 0.2222
-4.250 -0.5078 0.03451 0.02655 0.0053 1.0000 0.0795
-4.000 -0.4906 0.03210 0.02386 0.0071 1.0000 0.0781
-3.750 -0.4715 0.03004 0.02146 0.0088 1.0000 0.0774
-3.500 -0.4499 0.02804 0.01909 0.0103 1.0000 0.0760
-3.250 -0.4262 0.02629 0.01698 0.0114 1.0000 0.0751
-3.000 -0.4021 0.02488 0.01530 0.0124 1.0000 0.0757
-2.750 -0.3780 0.02375 0.01396 0.0133 1.0000 0.0779
-2.500 -0.3546 0.02284 0.01285 0.0142 1.0000 0.0826
-2.250 -0.3315 0.02182 0.01193 0.0148 1.0000 0.0888
-2.000 -0.0935 0.01675 0.01038 -0.0212 1.0000 1.0000
-1.750 -0.0871 0.01690 0.01035 -0.0180 1.0000 1.0000
-1.500 -0.0800 0.01708 0.01035 -0.0149 1.0000 1.0000
-1.250 -0.0721 0.01728 0.01041 -0.0120 1.0000 1.0000
-1.000 -0.0635 0.01750 0.01050 -0.0093 1.0000 1.0000
-0.750 -0.0432 0.01775 0.01061 -0.0089 0.9979 1.0000
-0.500 0.0043 0.01803 0.01070 -0.0137 0.9900 1.0000
-0.250 0.0528 0.01831 0.01085 -0.0186 0.9823 1.0000
0.000 0.0996 0.01853 0.01097 -0.0231 0.9733 1.0000
0.250 0.1449 0.01873 0.01109 -0.0272 0.9635 1.0000
0.500 0.1972 0.01888 0.01119 -0.0326 0.9550 1.0000
0.750 0.2452 0.01896 0.01125 -0.0370 0.9447 1.0000
1.000 0.2905 0.01903 0.01131 -0.0408 0.9338 1.0000
1.250 0.3543 0.01886 0.01117 -0.0479 0.9270 1.0000
1.500 0.3963 0.01877 0.01112 -0.0507 0.9144 1.0000
1.750 0.4357 0.01863 0.01103 -0.0527 0.9009 1.0000
2.000 0.4719 0.01845 0.01091 -0.0538 0.8870 1.0000
2.250 0.5044 0.01824 0.01076 -0.0540 0.8726 1.0000
2.500 0.5334 0.01801 0.01060 -0.0534 0.8575 1.0000
2.750 0.5597 0.01773 0.01038 -0.0521 0.8413 1.0000
3.000 0.5843 0.01738 0.01008 -0.0502 0.8240 1.0000
3.250 0.6078 0.01701 0.00977 -0.0480 0.8063 1.0000
3.500 0.6306 0.01665 0.00945 -0.0457 0.7881 1.0000
3.750 0.6508 0.01643 0.00929 -0.0432 0.7650 1.0000
4.000 0.6732 0.01602 0.00888 -0.0407 0.7422 1.0000
4.250 0.6940 0.01574 0.00863 -0.0381 0.7131 1.0000
4.500 0.7147 0.01552 0.00840 -0.0356 0.6788 1.0000
4.750 0.7351 0.01540 0.00820 -0.0332 0.6374 1.0000
5.000 0.7549 0.01544 0.00810 -0.0308 0.5877 1.0000
5.250 0.7737 0.01570 0.00816 -0.0284 0.5310 1.0000
5.500 0.7913 0.01618 0.00836 -0.0262 0.4699 1.0000
5.750 0.8073 0.01687 0.00872 -0.0238 0.4069 1.0000
6.000 0.8217 0.01772 0.00924 -0.0215 0.3409 1.0000
6.250 0.8337 0.01883 0.00994 -0.0189 0.2710 1.0000
6.500 0.8439 0.02025 0.01088 -0.0161 0.2063 1.0000
6.750 0.8552 0.02172 0.01194 -0.0134 0.1662 1.0000
7.000 0.8700 0.02296 0.01305 -0.0112 0.1418 1.0000
7.250 0.8869 0.02425 0.01424 -0.0094 0.1273 1.0000
7.500 0.9054 0.02552 0.01548 -0.0078 0.1162 1.0000
7.750 0.9257 0.02698 0.01681 -0.0066 0.1079 1.0000
8.000 0.9468 0.02828 0.01826 -0.0053 0.1015 1.0000
8.250 0.9699 0.03016 0.02002 -0.0048 0.0956 1.0000
8.500 0.9889 0.03152 0.02169 -0.0031 0.0909 1.0000
8.750 1.0093 0.03326 0.02360 -0.0018 0.0874 1.0000
9.000 1.0305 0.03545 0.02578 -0.0011 0.0839 1.0000
9.250 1.0445 0.03771 0.02841 0.0010 0.0813 1.0000
9.500 1.0557 0.03987 0.03102 0.0035 0.0791 1.0000
9.750 1.0650 0.04252 0.03407 0.0061 0.0779 1.0000
10.000 1.0710 0.04534 0.03727 0.0090 0.0769 1.0000
10.250 1.0747 0.04808 0.04034 0.0119 0.0756 1.0000
10.500 1.0804 0.05054 0.04299 0.0143 0.0738 1.0000
10.750 1.0900 0.05332 0.04583 0.0159 0.0721 1.0000
11.000 1.0827 0.05679 0.04967 0.0195 0.0722 1.0000
11.250 1.0461 0.06102 0.05450 0.0256 0.0739 1.0000
11.500 1.0090 0.06514 0.05896 0.0308 0.0754 1.0000
11.750 0.9740 0.06973 0.06378 0.0336 0.0767 1.0000
12.000 0.9394 0.07525 0.06949 0.0340 0.0780 1.0000
12.250 0.9041 0.08203 0.07641 0.0320 0.0795 1.0000
12.500 0.8762 0.08954 0.08398 0.0287 0.0810 1.0000
12.750 0.8677 0.09566 0.09012 0.0271 0.0823 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY RC-10(B)3 AIRFOIL (rc10b3-il)