BOEING-VERTOL VR-12 AIRFOIL (vr12-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-12 AIRFOIL (vr12-il) Reynolds number: 50,000 Max Cl/Cd: 28.89 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr12-il-50000-n5.txt Download as CSV file: xf-vr12-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-12 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4804 0.13327 0.12623 0.0001 1.0000 0.1082
-10.750 -0.4830 0.13031 0.12332 -0.0026 1.0000 0.1129
-10.500 -0.4978 0.12806 0.12117 -0.0079 1.0000 0.1148
-10.250 -0.4856 0.12286 0.11602 -0.0075 1.0000 0.1165
-10.000 -0.4714 0.11858 0.11172 -0.0067 1.0000 0.1192
-9.750 -0.4658 0.11486 0.10804 -0.0077 1.0000 0.1218
-9.500 -0.4638 0.11117 0.10439 -0.0094 1.0000 0.1243
-9.250 -0.3884 0.09214 0.08572 -0.0306 1.0000 0.0712
-9.000 -0.3817 0.08806 0.08167 -0.0302 1.0000 0.0704
-8.750 -0.3796 0.08385 0.07750 -0.0307 1.0000 0.0692
-8.500 -0.5012 0.08467 0.07789 -0.0325 1.0000 0.0586
-8.250 -0.5005 0.08063 0.07385 -0.0330 1.0000 0.0583
-8.000 -0.5022 0.07662 0.06979 -0.0334 1.0000 0.0581
-7.750 -0.5044 0.07269 0.06578 -0.0334 1.0000 0.0580
-7.500 -0.5060 0.06895 0.06195 -0.0328 1.0000 0.0578
-7.250 -0.5071 0.06544 0.05833 -0.0315 1.0000 0.0576
-7.000 -0.5083 0.06219 0.05497 -0.0295 1.0000 0.0573
-6.750 -0.5096 0.05912 0.05173 -0.0271 1.0000 0.0568
-6.500 -0.5102 0.05614 0.04856 -0.0244 1.0000 0.0565
-6.250 -0.5098 0.05325 0.04542 -0.0215 1.0000 0.0565
-6.000 -0.5082 0.05045 0.04226 -0.0185 1.0000 0.0572
-5.750 -0.5047 0.04780 0.03912 -0.0154 1.0000 0.0582
-5.500 -0.4977 0.04526 0.03620 -0.0126 1.0000 0.0590
-5.250 -0.4874 0.04271 0.03341 -0.0104 1.0000 0.0596
-5.000 -0.4749 0.04042 0.03085 -0.0083 1.0000 0.0601
-4.750 -0.4606 0.03851 0.02876 -0.0065 1.0000 0.0616
-4.500 -0.4450 0.03689 0.02690 -0.0048 1.0000 0.0641
-4.250 -0.4276 0.03520 0.02488 -0.0032 1.0000 0.0663
-4.000 -0.4080 0.03351 0.02283 -0.0017 1.0000 0.0676
-3.750 -0.3871 0.03205 0.02100 -0.0003 1.0000 0.0693
-3.500 -0.3551 0.03058 0.01922 -0.0012 0.9962 0.0728
-3.250 -0.3208 0.02927 0.01780 -0.0025 0.9918 0.0761
-3.000 -0.2853 0.02817 0.01654 -0.0038 0.9872 0.0793
-2.750 -0.2480 0.02741 0.01552 -0.0054 0.9829 0.0850
-2.500 -0.2130 0.02652 0.01464 -0.0069 0.9781 0.0906
-2.250 -0.1760 0.02588 0.01385 -0.0087 0.9733 0.0969
-2.000 -0.1436 0.02530 0.01324 -0.0098 0.9676 0.1070
-1.750 -0.1101 0.02474 0.01270 -0.0112 0.9616 0.1259
-1.500 0.0038 0.02150 0.01279 -0.0258 0.9747 1.0000
-1.250 0.0404 0.02168 0.01258 -0.0280 0.9661 1.0000
-1.000 0.0820 0.02186 0.01246 -0.0311 0.9578 1.0000
-0.750 0.1213 0.02201 0.01238 -0.0337 0.9469 1.0000
-0.500 0.1707 0.02204 0.01219 -0.0379 0.9330 1.0000
-0.250 0.2259 0.02183 0.01180 -0.0426 0.9139 1.0000
0.000 0.2649 0.02161 0.01144 -0.0440 0.8915 1.0000
0.250 0.2991 0.02143 0.01115 -0.0445 0.8730 1.0000
0.500 0.3278 0.02133 0.01097 -0.0440 0.8564 1.0000
0.750 0.3549 0.02121 0.01079 -0.0432 0.8392 1.0000
1.000 0.3807 0.02106 0.01058 -0.0420 0.8213 1.0000
1.250 0.4025 0.02097 0.01044 -0.0402 0.8005 1.0000
1.500 0.4262 0.02079 0.01022 -0.0384 0.7794 1.0000
1.750 0.4488 0.02062 0.01001 -0.0365 0.7560 1.0000
2.000 0.4707 0.02047 0.00981 -0.0344 0.7290 1.0000
2.250 0.4912 0.02037 0.00966 -0.0321 0.6959 1.0000
2.500 0.5115 0.02028 0.00952 -0.0299 0.6552 1.0000
2.750 0.5325 0.02020 0.00928 -0.0275 0.6034 1.0000
3.000 0.5537 0.02022 0.00894 -0.0251 0.5404 1.0000
3.250 0.5729 0.02060 0.00883 -0.0227 0.4800 1.0000
3.500 0.5917 0.02123 0.00904 -0.0209 0.4334 1.0000
3.750 0.6111 0.02191 0.00945 -0.0194 0.3982 1.0000
4.250 0.6516 0.02328 0.01038 -0.0169 0.3522 1.0000
4.500 0.6729 0.02396 0.01092 -0.0158 0.3350 1.0000
4.750 0.6948 0.02462 0.01150 -0.0148 0.3207 1.0000
5.000 0.7173 0.02530 0.01208 -0.0140 0.3081 1.0000
5.250 0.7402 0.02597 0.01269 -0.0132 0.2965 1.0000
5.500 0.7636 0.02665 0.01341 -0.0125 0.2859 1.0000
5.750 0.7873 0.02737 0.01407 -0.0118 0.2773 1.0000
6.000 0.8105 0.02809 0.01486 -0.0111 0.2682 1.0000
6.250 0.8340 0.02887 0.01562 -0.0104 0.2603 1.0000
6.500 0.8569 0.02966 0.01652 -0.0097 0.2526 1.0000
6.750 0.8810 0.03053 0.01736 -0.0092 0.2469 1.0000
7.000 0.9032 0.03148 0.01852 -0.0085 0.2404 1.0000
7.250 0.9262 0.03239 0.01947 -0.0078 0.2348 1.0000
7.500 0.9476 0.03343 0.02065 -0.0070 0.2289 1.0000
7.750 0.9681 0.03448 0.02187 -0.0062 0.2229 1.0000
8.000 0.9907 0.03549 0.02290 -0.0056 0.2184 1.0000
8.250 1.0100 0.03686 0.02454 -0.0046 0.2140 1.0000
8.500 1.0280 0.03823 0.02619 -0.0036 0.2090 1.0000
8.750 1.0484 0.03936 0.02739 -0.0028 0.2044 1.0000
9.000 1.0657 0.04077 0.02897 -0.0018 0.1998 1.0000
9.250 1.0784 0.04249 0.03107 -0.0003 0.1948 1.0000
9.500 1.0954 0.04380 0.03253 0.0008 0.1902 1.0000
9.750 1.1120 0.04512 0.03394 0.0019 0.1856 1.0000
10.000 1.1164 0.04719 0.03642 0.0039 0.1800 1.0000
10.250 1.1305 0.04846 0.03782 0.0052 0.1752 1.0000
10.500 1.1432 0.04991 0.03941 0.0065 0.1708 1.0000
10.750 1.1373 0.05270 0.04262 0.0091 0.1662 1.0000
11.000 1.1414 0.05461 0.04473 0.0109 0.1619 1.0000
11.250 1.1652 0.05508 0.04511 0.0115 0.1577 1.0000
11.500 1.1366 0.05932 0.04982 0.0149 0.1548 1.0000
11.750 1.1039 0.06367 0.05444 0.0180 0.1527 1.0000
12.000 1.0623 0.06971 0.06069 0.0189 0.1511 1.0000
12.250 0.9522 0.08754 0.07862 0.0111 0.1513 1.0000
12.500 0.9027 0.10109 0.09216 0.0035 0.1471 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-12 AIRFOIL (vr12-il)