RG 12A-1.8/9.0 AIRFOIL (rg12a189-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RG 12A-1.8/9.0 AIRFOIL (rg12a189-il) Reynolds number: 200,000 Max Cl/Cd: 62.21 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg12a189-il-200000-n5.txt Download as CSV file: xf-rg12a189-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 12A-1.8/9.0 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3926 0.09361 0.09012 -0.0274 1.0000 0.0142
-10.000 -0.3944 0.08947 0.08601 -0.0282 1.0000 0.0139
-9.750 -0.4872 0.09500 0.09134 -0.0252 1.0000 0.0141
-9.500 -0.4886 0.09108 0.08745 -0.0264 1.0000 0.0138
-9.250 -0.4910 0.08697 0.08338 -0.0279 1.0000 0.0135
-9.000 -0.4947 0.08271 0.07917 -0.0296 1.0000 0.0132
-8.750 -0.5020 0.07784 0.07436 -0.0319 1.0000 0.0130
-8.500 -0.5112 0.07312 0.06970 -0.0346 1.0000 0.0127
-8.250 -0.5283 0.06806 0.06469 -0.0377 1.0000 0.0126
-8.000 -0.5404 0.06255 0.05913 -0.0397 1.0000 0.0123
-7.750 -0.5507 0.05696 0.05344 -0.0405 1.0000 0.0119
-7.500 -0.5583 0.05127 0.04756 -0.0403 1.0000 0.0115
-7.250 -0.5673 0.04381 0.03973 -0.0391 1.0000 0.0109
-7.000 -0.5726 0.03599 0.03120 -0.0369 1.0000 0.0103
-6.750 -0.5637 0.03229 0.02703 -0.0352 1.0000 0.0102
-6.500 -0.5514 0.02912 0.02341 -0.0336 1.0000 0.0102
-6.250 -0.5361 0.02642 0.02028 -0.0322 1.0000 0.0102
-6.000 -0.5185 0.02415 0.01766 -0.0309 1.0000 0.0103
-5.750 -0.4993 0.02225 0.01546 -0.0297 1.0000 0.0106
-5.500 -0.4727 0.02060 0.01353 -0.0299 0.9986 0.0110
-5.250 -0.4398 0.01913 0.01182 -0.0312 0.9957 0.0115
-5.000 -0.4072 0.01787 0.01038 -0.0324 0.9927 0.0124
-4.750 -0.3742 0.01701 0.00936 -0.0336 0.9888 0.0142
-4.500 -0.3411 0.01585 0.00810 -0.0350 0.9853 0.0161
-4.250 -0.3079 0.01500 0.00715 -0.0363 0.9814 0.0180
-4.000 -0.2757 0.01430 0.00632 -0.0373 0.9761 0.0210
-3.750 -0.2410 0.01364 0.00564 -0.0389 0.9720 0.0293
-3.500 -0.2096 0.01297 0.00506 -0.0398 0.9662 0.0526
-3.250 -0.1766 0.01240 0.00465 -0.0412 0.9606 0.0922
-3.000 -0.1437 0.01170 0.00430 -0.0427 0.9554 0.1720
-2.750 -0.1134 0.01103 0.00400 -0.0437 0.9482 0.2693
-2.500 -0.0793 0.01032 0.00371 -0.0454 0.9433 0.3898
-2.250 -0.0520 0.00959 0.00351 -0.0455 0.9345 0.5327
-2.000 -0.0238 0.00912 0.00363 -0.0450 0.9275 0.7024
-1.750 0.0069 0.00903 0.00358 -0.0451 0.9186 0.7571
-1.500 0.0375 0.00896 0.00350 -0.0452 0.9085 0.7862
-1.250 0.0698 0.00889 0.00340 -0.0457 0.8987 0.8084
-1.000 0.1026 0.00883 0.00329 -0.0462 0.8879 0.8278
-0.750 0.1337 0.00878 0.00319 -0.0465 0.8745 0.8446
-0.500 0.1642 0.00873 0.00311 -0.0465 0.8595 0.8595
-0.250 0.1928 0.00870 0.00302 -0.0462 0.8422 0.8738
0.000 0.2206 0.00868 0.00296 -0.0457 0.8235 0.8880
0.250 0.2481 0.00869 0.00289 -0.0451 0.8038 0.9014
0.500 0.2751 0.00870 0.00284 -0.0445 0.7823 0.9116
0.750 0.3032 0.00874 0.00277 -0.0442 0.7605 0.9194
1.000 0.3307 0.00879 0.00274 -0.0439 0.7374 0.9271
1.250 0.3592 0.00886 0.00271 -0.0439 0.7145 0.9334
1.500 0.3875 0.00896 0.00272 -0.0439 0.6914 0.9406
1.750 0.4169 0.00906 0.00273 -0.0441 0.6678 0.9470
2.000 0.4465 0.00918 0.00277 -0.0445 0.6439 0.9544
2.250 0.4771 0.00932 0.00284 -0.0451 0.6192 0.9614
2.500 0.5077 0.00947 0.00291 -0.0457 0.5934 0.9694
2.750 0.5399 0.00964 0.00299 -0.0468 0.5666 0.9775
3.000 0.5718 0.00982 0.00310 -0.0478 0.5390 0.9874
3.250 0.6026 0.01004 0.00325 -0.0486 0.5099 1.0000
3.500 0.6236 0.01027 0.00340 -0.0474 0.4830 1.0000
3.750 0.6459 0.01054 0.00357 -0.0465 0.4547 1.0000
4.000 0.6689 0.01083 0.00377 -0.0457 0.4252 1.0000
4.250 0.6921 0.01115 0.00402 -0.0450 0.3952 1.0000
4.500 0.7154 0.01150 0.00427 -0.0443 0.3638 1.0000
4.750 0.7388 0.01188 0.00455 -0.0436 0.3303 1.0000
5.000 0.7619 0.01230 0.00487 -0.0429 0.2971 1.0000
5.250 0.7851 0.01274 0.00525 -0.0423 0.2651 1.0000
5.500 0.8072 0.01330 0.00564 -0.0415 0.2240 1.0000
5.750 0.8290 0.01394 0.00608 -0.0408 0.1803 1.0000
6.000 0.8505 0.01462 0.00658 -0.0400 0.1409 1.0000
6.250 0.8721 0.01532 0.00712 -0.0393 0.1065 1.0000
6.500 0.8935 0.01604 0.00775 -0.0385 0.0789 1.0000
6.750 0.9148 0.01680 0.00841 -0.0376 0.0555 1.0000
7.000 0.9356 0.01760 0.00917 -0.0367 0.0393 1.0000
7.250 0.9559 0.01848 0.01005 -0.0356 0.0302 1.0000
7.500 0.9768 0.01926 0.01096 -0.0346 0.0262 1.0000
7.750 0.9963 0.02017 0.01195 -0.0334 0.0237 1.0000
8.000 1.0135 0.02128 0.01319 -0.0320 0.0217 1.0000
8.250 1.0328 0.02215 0.01420 -0.0308 0.0202 1.0000
8.500 1.0506 0.02315 0.01533 -0.0295 0.0190 1.0000
8.750 1.0673 0.02428 0.01659 -0.0280 0.0180 1.0000
9.000 1.0831 0.02547 0.01791 -0.0265 0.0173 1.0000
9.250 1.0977 0.02681 0.01936 -0.0249 0.0165 1.0000
9.500 1.1100 0.02846 0.02112 -0.0231 0.0157 1.0000
9.750 1.1216 0.03038 0.02320 -0.0212 0.0148 1.0000
10.000 1.1356 0.03170 0.02473 -0.0197 0.0141 1.0000
10.250 1.1462 0.03343 0.02671 -0.0177 0.0135 1.0000
10.500 1.1542 0.03533 0.02884 -0.0155 0.0130 1.0000
10.750 1.1598 0.03719 0.03093 -0.0132 0.0124 1.0000
11.000 1.1633 0.03901 0.03294 -0.0109 0.0118 1.0000
11.250 1.1655 0.04069 0.03478 -0.0089 0.0112 1.0000
11.500 1.1655 0.04267 0.03689 -0.0071 0.0108 1.0000
11.750 1.1621 0.04518 0.03954 -0.0056 0.0103 1.0000
12.000 1.1512 0.04892 0.04350 -0.0043 0.0100 1.0000
12.250 1.1400 0.05280 0.04764 -0.0039 0.0098 1.0000
12.500 1.1298 0.05669 0.05178 -0.0042 0.0097 1.0000
12.750 1.1177 0.06122 0.05656 -0.0055 0.0096 1.0000
13.000 1.1033 0.06652 0.06210 -0.0077 0.0096 1.0000
13.250 1.0877 0.07254 0.06833 -0.0108 0.0096 1.0000
13.500 1.0711 0.07925 0.07524 -0.0148 0.0096 1.0000
13.750 1.0532 0.08681 0.08299 -0.0196 0.0096 1.0000
14.000 1.0341 0.09521 0.09156 -0.0250 0.0096 1.0000
14.250 1.0147 0.10422 0.10070 -0.0308 0.0097 1.0000
14.500 0.9943 0.11415 0.11076 -0.0371 0.0098 1.0000
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Polar data table (+)
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