RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL (r1082-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL (r1082-il) Reynolds number: 500,000 Max Cl/Cd: 102.23 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-r1082-il-500000-n5.txt Download as CSV file: xf-r1082-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RONCZ 1082 VOYAGER ROOT OUTER AFT WING 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.000 -0.6409 0.08889 0.08582 -0.0705 1.0000 0.0093
-15.750 -0.6960 0.07391 0.07047 -0.0796 1.0000 0.0091
-15.500 -0.7149 0.06681 0.06321 -0.0840 1.0000 0.0091
-15.250 -0.7293 0.06043 0.05667 -0.0884 0.9995 0.0091
-15.000 -0.7433 0.05468 0.05075 -0.0910 0.9691 0.0091
-14.750 -0.7192 0.04825 0.04405 -0.1032 0.9343 0.0092
-14.500 -0.6721 0.04232 0.03775 -0.1188 0.9057 0.0093
-14.250 -0.6473 0.03862 0.03360 -0.1263 0.8525 0.0094
-14.000 -0.6453 0.03681 0.03151 -0.1260 0.8167 0.0094
-13.500 -0.6436 0.03358 0.02789 -0.1236 0.7758 0.0096
-13.250 -0.6406 0.03217 0.02631 -0.1222 0.7609 0.0097
-13.000 -0.6374 0.03072 0.02469 -0.1207 0.7484 0.0097
-12.750 -0.6314 0.02955 0.02337 -0.1192 0.7373 0.0098
-12.500 -0.6245 0.02846 0.02214 -0.1177 0.7268 0.0100
-12.250 -0.6169 0.02737 0.02092 -0.1161 0.7176 0.0101
-12.000 -0.6088 0.02632 0.01973 -0.1144 0.7089 0.0102
-11.750 -0.5987 0.02541 0.01870 -0.1128 0.7011 0.0102
-11.500 -0.5879 0.02456 0.01773 -0.1113 0.6936 0.0104
-11.250 -0.5769 0.02369 0.01675 -0.1096 0.6874 0.0106
-11.000 -0.5635 0.02300 0.01595 -0.1082 0.6814 0.0108
-10.750 -0.5493 0.02240 0.01522 -0.1067 0.6757 0.0111
-10.500 -0.5365 0.02163 0.01436 -0.1051 0.6707 0.0111
-10.250 -0.5219 0.02099 0.01363 -0.1036 0.6657 0.0114
-10.000 -0.5089 0.02027 0.01285 -0.1019 0.6610 0.0115
-9.750 -0.4966 0.01955 0.01206 -0.1000 0.6568 0.0118
-9.500 -0.4821 0.01893 0.01140 -0.0984 0.6529 0.0120
-9.250 -0.4666 0.01840 0.01081 -0.0968 0.6489 0.0122
-9.000 -0.4511 0.01790 0.01025 -0.0951 0.6450 0.0124
-8.750 -0.4355 0.01743 0.00971 -0.0934 0.6415 0.0127
-8.500 -0.4196 0.01698 0.00920 -0.0917 0.6383 0.0130
-8.250 -0.4036 0.01656 0.00873 -0.0899 0.6349 0.0133
-8.000 -0.3875 0.01621 0.00833 -0.0881 0.6315 0.0136
-7.750 -0.3713 0.01588 0.00794 -0.0863 0.6283 0.0138
-7.500 -0.3538 0.01557 0.00756 -0.0846 0.6255 0.0142
-7.250 -0.3357 0.01519 0.00714 -0.0832 0.6229 0.0148
-7.000 -0.3156 0.01487 0.00679 -0.0820 0.6207 0.0152
-6.750 -0.2950 0.01457 0.00646 -0.0809 0.6183 0.0160
-6.500 -0.2738 0.01430 0.00616 -0.0799 0.6158 0.0171
-6.250 -0.2529 0.01401 0.00585 -0.0789 0.6136 0.0183
-6.000 -0.2314 0.01375 0.00556 -0.0779 0.6113 0.0202
-5.750 -0.2101 0.01349 0.00528 -0.0769 0.6092 0.0231
-5.500 -0.1887 0.01323 0.00503 -0.0759 0.6073 0.0291
-5.250 -0.1669 0.01296 0.00479 -0.0750 0.6056 0.0387
-5.000 -0.1447 0.01270 0.00458 -0.0742 0.6037 0.0505
-4.750 -0.1220 0.01246 0.00439 -0.0735 0.6018 0.0639
-4.500 -0.0995 0.01221 0.00421 -0.0727 0.5997 0.0798
-4.250 -0.0766 0.01198 0.00405 -0.0720 0.5977 0.0982
-4.000 -0.0542 0.01172 0.00388 -0.0713 0.5959 0.1229
-3.750 -0.0322 0.01144 0.00372 -0.0705 0.5943 0.1563
-3.500 -0.0113 0.01109 0.00355 -0.0695 0.5927 0.2058
-3.250 0.0092 0.01070 0.00338 -0.0685 0.5912 0.2667
-3.000 0.0284 0.01023 0.00320 -0.0673 0.5898 0.3477
-2.750 0.0483 0.00976 0.00302 -0.0662 0.5885 0.4273
-2.500 0.0676 0.00924 0.00285 -0.0649 0.5871 0.5192
-2.250 0.0873 0.00878 0.00278 -0.0636 0.5857 0.6147
-2.000 0.1114 0.00870 0.00292 -0.0628 0.5843 0.6819
-1.750 0.1387 0.00876 0.00300 -0.0627 0.5829 0.7071
-1.500 0.1665 0.00882 0.00304 -0.0627 0.5813 0.7217
-1.250 0.1945 0.00888 0.00304 -0.0628 0.5797 0.7328
-1.000 0.2223 0.00897 0.00311 -0.0628 0.5782 0.7421
-0.750 0.2502 0.00907 0.00317 -0.0629 0.5769 0.7520
-0.500 0.2780 0.00920 0.00328 -0.0629 0.5757 0.7623
-0.250 0.3058 0.00933 0.00339 -0.0629 0.5744 0.7699
0.000 0.3340 0.00945 0.00349 -0.0630 0.5732 0.7754
0.250 0.3621 0.00950 0.00353 -0.0632 0.5721 0.7796
0.500 0.3904 0.00955 0.00355 -0.0635 0.5710 0.7830
0.750 0.4183 0.00959 0.00360 -0.0637 0.5696 0.7852
1.000 0.4463 0.00964 0.00365 -0.0638 0.5681 0.7874
1.250 0.4743 0.00969 0.00371 -0.0640 0.5666 0.7896
1.500 0.5025 0.00974 0.00376 -0.0643 0.5654 0.7918
1.750 0.5306 0.00979 0.00380 -0.0645 0.5641 0.7939
2.000 0.5588 0.00984 0.00384 -0.0648 0.5626 0.7957
2.250 0.5870 0.00989 0.00387 -0.0651 0.5612 0.7974
2.500 0.6153 0.00996 0.00391 -0.0655 0.5597 0.7989
2.750 0.6439 0.01003 0.00396 -0.0659 0.5582 0.8002
3.000 0.6718 0.01009 0.00403 -0.0661 0.5566 0.8015
3.250 0.6988 0.01013 0.00411 -0.0662 0.5548 0.8027
3.500 0.7258 0.01017 0.00419 -0.0662 0.5527 0.8039
3.750 0.7529 0.01021 0.00427 -0.0663 0.5504 0.8051
4.000 0.7799 0.01025 0.00433 -0.0664 0.5481 0.8063
4.250 0.8068 0.01030 0.00440 -0.0664 0.5458 0.8078
4.500 0.8340 0.01036 0.00446 -0.0665 0.5433 0.8093
4.750 0.8613 0.01044 0.00454 -0.0667 0.5407 0.8108
5.000 0.8868 0.01048 0.00464 -0.0665 0.5378 0.8122
5.250 0.9125 0.01051 0.00473 -0.0663 0.5344 0.8136
5.500 0.9381 0.01056 0.00480 -0.0662 0.5308 0.8150
5.750 0.9637 0.01061 0.00485 -0.0660 0.5271 0.8163
6.000 0.9886 0.01066 0.00495 -0.0657 0.5231 0.8176
6.250 1.0124 0.01069 0.00506 -0.0651 0.5179 0.8189
6.500 1.0356 0.01074 0.00514 -0.0644 0.5125 0.8205
6.750 1.0588 0.01080 0.00525 -0.0638 0.5070 0.8222
7.000 1.0810 0.01086 0.00537 -0.0629 0.4998 0.8238
7.250 1.1020 0.01095 0.00548 -0.0619 0.4926 0.8255
7.500 1.1222 0.01104 0.00562 -0.0606 0.4828 0.8273
7.750 1.1388 0.01114 0.00574 -0.0587 0.4711 0.8293
8.000 1.1516 0.01131 0.00589 -0.0560 0.4557 0.8315
8.250 1.1585 0.01166 0.00617 -0.0524 0.4345 0.8338
8.500 1.1609 0.01217 0.00660 -0.0480 0.4094 0.8363
8.750 1.1591 0.01287 0.00721 -0.0432 0.3849 0.8393
9.000 1.1553 0.01373 0.00798 -0.0383 0.3601 0.8430
9.250 1.1503 0.01475 0.00892 -0.0336 0.3365 0.8469
9.500 1.1470 0.01586 0.00996 -0.0296 0.3158 0.8504
9.750 1.1407 0.01719 0.01125 -0.0254 0.2955 0.8542
10.000 1.1357 0.01863 0.01265 -0.0218 0.2756 0.8585
10.250 1.1323 0.02017 0.01415 -0.0188 0.2579 0.8632
10.500 1.1296 0.02179 0.01574 -0.0160 0.2413 0.8678
10.750 1.1264 0.02349 0.01743 -0.0133 0.2250 0.8733
11.000 1.1249 0.02519 0.01912 -0.0110 0.2103 0.8795
11.250 1.1240 0.02689 0.02082 -0.0089 0.1965 0.8866
11.500 1.1231 0.02863 0.02256 -0.0069 0.1828 0.8961
11.750 1.1228 0.03041 0.02435 -0.0051 0.1692 0.9083
12.000 1.1273 0.03219 0.02615 -0.0043 0.1567 0.9260
12.250 1.1409 0.03439 0.02834 -0.0060 0.1391 0.9498
12.500 1.1515 0.03614 0.03010 -0.0065 0.1306 1.0000
12.750 1.1543 0.03807 0.03199 -0.0055 0.1187 1.0000
13.000 1.1599 0.03981 0.03371 -0.0048 0.1091 1.0000
13.250 1.1637 0.04174 0.03561 -0.0040 0.0993 1.0000
13.500 1.1678 0.04367 0.03749 -0.0033 0.0900 1.0000
13.750 1.1723 0.04560 0.03940 -0.0027 0.0820 1.0000
14.000 1.1785 0.04742 0.04120 -0.0023 0.0745 1.0000
14.250 1.1834 0.04939 0.04317 -0.0018 0.0680 1.0000
14.500 1.1897 0.05125 0.04502 -0.0014 0.0626 1.0000
14.750 1.1963 0.05312 0.04691 -0.0011 0.0580 1.0000
15.000 1.2020 0.05511 0.04891 -0.0009 0.0536 1.0000
15.250 1.2088 0.05702 0.05083 -0.0007 0.0497 1.0000
15.500 1.2150 0.05902 0.05287 -0.0005 0.0464 1.0000
15.750 1.2201 0.06116 0.05502 -0.0004 0.0428 1.0000
16.000 1.2269 0.06313 0.05703 -0.0004 0.0401 1.0000
16.250 1.2309 0.06546 0.05938 -0.0003 0.0373 1.0000
16.500 1.2373 0.06755 0.06152 -0.0004 0.0348 1.0000
16.750 1.2415 0.06993 0.06392 -0.0005 0.0321 1.0000
17.000 1.2466 0.07223 0.06627 -0.0007 0.0302 1.0000
17.250 1.2512 0.07461 0.06869 -0.0009 0.0279 1.0000
17.500 1.2541 0.07723 0.07136 -0.0013 0.0259 1.0000
17.750 1.2585 0.07970 0.07388 -0.0016 0.0241 1.0000
18.000 1.2600 0.08259 0.07681 -0.0021 0.0221 1.0000
18.250 1.2638 0.08522 0.07951 -0.0026 0.0206 1.0000
18.500 1.2641 0.08832 0.08265 -0.0032 0.0186 1.0000
18.750 1.2665 0.09118 0.08558 -0.0039 0.0173 1.0000
19.000 1.2664 0.09444 0.08889 -0.0047 0.0159 1.0000
19.250 1.2673 0.09760 0.09212 -0.0056 0.0148 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL (r1082-il)