AIRFOIL PROFILE12A 9.00% (rg12a-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: AIRFOIL PROFILE12A 9.00% (rg12a-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-rg12a-il-200000-n5.txt Download as CSV file: xf-rg12a-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL PROFILE12A 9.00%
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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4865 0.09534 0.09167 -0.0248 1.0000 0.0140
-9.500 -0.4881 0.09123 0.08760 -0.0263 1.0000 0.0138
-9.250 -0.4904 0.08715 0.08356 -0.0278 1.0000 0.0135
-9.000 -0.4949 0.08267 0.07914 -0.0296 1.0000 0.0132
-8.750 -0.5011 0.07806 0.07458 -0.0318 1.0000 0.0129
-8.500 -0.5119 0.07296 0.06955 -0.0347 1.0000 0.0127
-8.250 -0.5285 0.06802 0.06465 -0.0377 1.0000 0.0125
-8.000 -0.5394 0.06277 0.05936 -0.0396 1.0000 0.0123
-7.750 -0.5499 0.05715 0.05363 -0.0405 1.0000 0.0120
-7.500 -0.5577 0.05142 0.04772 -0.0403 1.0000 0.0116
-7.250 -0.5681 0.04348 0.03938 -0.0391 1.0000 0.0109
-7.000 -0.5725 0.03594 0.03113 -0.0369 1.0000 0.0103
-6.750 -0.5634 0.03233 0.02707 -0.0352 1.0000 0.0102
-6.500 -0.5513 0.02913 0.02342 -0.0336 1.0000 0.0102
-6.250 -0.5360 0.02644 0.02032 -0.0322 1.0000 0.0102
-6.000 -0.5185 0.02413 0.01764 -0.0309 1.0000 0.0104
-5.750 -0.4992 0.02226 0.01547 -0.0297 1.0000 0.0106
-5.500 -0.4723 0.02062 0.01354 -0.0300 0.9985 0.0110
-5.250 -0.4393 0.01913 0.01182 -0.0313 0.9956 0.0116
-5.000 -0.4066 0.01787 0.01038 -0.0325 0.9925 0.0125
-4.750 -0.3737 0.01700 0.00934 -0.0337 0.9887 0.0142
-4.500 -0.3407 0.01583 0.00808 -0.0351 0.9853 0.0162
-4.250 -0.3077 0.01500 0.00716 -0.0364 0.9814 0.0180
-4.000 -0.2757 0.01431 0.00633 -0.0373 0.9762 0.0208
-3.750 -0.2412 0.01364 0.00565 -0.0389 0.9721 0.0292
-3.500 -0.2096 0.01297 0.00506 -0.0398 0.9663 0.0525
-3.250 -0.1766 0.01240 0.00465 -0.0412 0.9607 0.0919
-3.000 -0.1435 0.01170 0.00430 -0.0428 0.9554 0.1713
-2.750 -0.1131 0.01102 0.00399 -0.0438 0.9481 0.2696
-2.500 -0.0793 0.01032 0.00371 -0.0454 0.9430 0.3898
-2.250 -0.0518 0.00959 0.00351 -0.0455 0.9344 0.5321
-2.000 -0.0237 0.00912 0.00362 -0.0450 0.9273 0.7013
-1.750 0.0071 0.00903 0.00358 -0.0451 0.9185 0.7576
-1.500 0.0377 0.00896 0.00350 -0.0453 0.9084 0.7860
-1.250 0.0698 0.00889 0.00340 -0.0457 0.8985 0.8085
-1.000 0.1025 0.00883 0.00329 -0.0462 0.8876 0.8279
-0.750 0.1338 0.00877 0.00319 -0.0465 0.8743 0.8445
-0.500 0.1642 0.00873 0.00310 -0.0465 0.8592 0.8595
-0.250 0.1930 0.00870 0.00302 -0.0462 0.8421 0.8737
0.000 0.2208 0.00869 0.00296 -0.0457 0.8234 0.8880
0.250 0.2482 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 0.00278 -0.0442 0.7605 0.9194
1.000 0.3307 0.00879 0.00274 -0.0439 0.7376 0.9271
1.250 0.3592 0.00886 0.00272 -0.0439 0.7145 0.9334
1.500 0.3874 0.00896 0.00272 -0.0439 0.6915 0.9406
1.750 0.4169 0.00906 0.00274 -0.0441 0.6679 0.9470
2.000 0.4464 0.00918 0.00277 -0.0445 0.6441 0.9544
2.500 0.5077 0.00947 0.00291 -0.0457 0.5939 0.9694
2.750 0.5399 0.00964 0.00300 -0.0467 0.5666 0.9775
3.000 0.5717 0.00982 0.00310 -0.0478 0.5391 0.9873
3.250 0.6028 0.01004 0.00325 -0.0487 0.5102 1.0000
3.500 0.6237 0.01027 0.00340 -0.0474 0.4832 1.0000
3.750 0.6459 0.01054 0.00357 -0.0465 0.4544 1.0000
4.000 0.6688 0.01083 0.00377 -0.0457 0.4252 1.0000
4.250 0.6921 0.01115 0.00402 -0.0450 0.3954 1.0000
4.500 0.7154 0.01150 0.00427 -0.0443 0.3639 1.0000
4.750 0.7387 0.01188 0.00455 -0.0436 0.3303 1.0000
5.000 0.7618 0.01230 0.00487 -0.0429 0.2973 1.0000
5.250 0.7849 0.01275 0.00525 -0.0422 0.2648 1.0000
5.500 0.8071 0.01330 0.00564 -0.0415 0.2243 1.0000
5.750 0.8289 0.01393 0.00608 -0.0407 0.1812 1.0000
6.000 0.8503 0.01462 0.00658 -0.0400 0.1410 1.0000
6.250 0.8719 0.01532 0.00712 -0.0392 0.1064 1.0000
6.500 0.8932 0.01605 0.00775 -0.0384 0.0785 1.0000
6.750 0.9145 0.01679 0.00841 -0.0376 0.0555 1.0000
7.000 0.9353 0.01760 0.00917 -0.0366 0.0392 1.0000
7.250 0.9555 0.01849 0.01005 -0.0356 0.0300 1.0000
7.500 0.9765 0.01926 0.01095 -0.0345 0.0262 1.0000
7.750 0.9959 0.02017 0.01194 -0.0334 0.0237 1.0000
8.000 1.0131 0.02129 0.01319 -0.0319 0.0217 1.0000
8.250 1.0324 0.02214 0.01419 -0.0307 0.0202 1.0000
8.500 1.0503 0.02314 0.01532 -0.0294 0.0189 1.0000
8.750 1.0669 0.02427 0.01658 -0.0279 0.0181 1.0000
9.000 1.0827 0.02546 0.01790 -0.0264 0.0172 1.0000
9.250 1.0973 0.02680 0.01935 -0.0248 0.0166 1.0000
9.500 1.1096 0.02843 0.02109 -0.0230 0.0157 1.0000
9.750 1.1212 0.03032 0.02314 -0.0211 0.0148 1.0000
10.000 1.1351 0.03167 0.02469 -0.0196 0.0140 1.0000
10.250 1.1456 0.03341 0.02669 -0.0176 0.0135 1.0000
10.500 1.1536 0.03529 0.02880 -0.0153 0.0130 1.0000
10.750 1.1593 0.03721 0.03095 -0.0130 0.0124 1.0000
11.000 1.1629 0.03892 0.03285 -0.0108 0.0118 1.0000
11.250 1.1649 0.04071 0.03480 -0.0087 0.0112 1.0000
11.500 1.1651 0.04257 0.03677 -0.0070 0.0107 1.0000
11.750 1.1620 0.04508 0.03944 -0.0055 0.0103 1.0000
12.000 1.1516 0.04877 0.04336 -0.0042 0.0100 1.0000
12.250 1.1397 0.05277 0.04760 -0.0037 0.0098 1.0000
12.500 1.1296 0.05660 0.05169 -0.0041 0.0097 1.0000
12.750 1.1174 0.06115 0.05649 -0.0053 0.0096 1.0000
13.000 1.1033 0.06642 0.06200 -0.0075 0.0096 1.0000
13.250 1.0878 0.07239 0.06818 -0.0107 0.0095 1.0000
13.500 1.0711 0.07915 0.07514 -0.0146 0.0096 1.0000
13.750 1.0531 0.08669 0.08286 -0.0193 0.0096 1.0000
14.000 1.0343 0.09500 0.09133 -0.0248 0.0096 1.0000
14.250 1.0142 0.10422 0.10070 -0.0307 0.0097 1.0000
14.500 0.9932 0.11433 0.11094 -0.0371 0.0098 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AIRFOIL PROFILE12A 9.00% (rg12a-il)