BOEING-VERTOL VR-12 AIRFOIL (vr12-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-12 AIRFOIL (vr12-il) Reynolds number: 100,000 Max Cl/Cd: 39.44 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr12-il-100000.txt Download as CSV file: xf-vr12-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4654 0.09620 0.09159 -0.0155 1.0000 0.0966
-8.500 -0.4846 0.09151 0.08697 -0.0247 1.0000 0.0989
-8.250 -0.5117 0.08817 0.08353 -0.0308 1.0000 0.0997
-8.000 -0.4789 0.08303 0.07858 -0.0251 1.0000 0.1038
-7.750 -0.4742 0.07984 0.07542 -0.0253 1.0000 0.1088
-7.500 -0.4929 0.07627 0.07173 -0.0298 1.0000 0.1139
-7.250 -0.5060 0.07240 0.06779 -0.0293 1.0000 0.1164
-7.000 -0.4867 0.06981 0.06540 -0.0260 1.0000 0.1240
-6.750 -0.5157 0.06890 0.06421 -0.0237 1.0000 0.1296
-6.500 -0.5124 0.06482 0.06030 -0.0209 1.0000 0.1338
-6.250 -0.5123 0.06290 0.05840 -0.0177 1.0000 0.1406
-6.000 -0.5244 0.06037 0.05571 -0.0152 1.0000 0.1483
-5.750 -0.5191 0.05815 0.05355 -0.0123 1.0000 0.1553
-5.500 -0.5218 0.05558 0.05087 -0.0098 1.0000 0.1657
-5.250 -0.5213 0.05343 0.04856 -0.0075 1.0000 0.1797
-5.000 -0.5171 0.05127 0.04633 -0.0051 1.0000 0.1951
-4.750 -0.5099 0.04900 0.04412 -0.0025 1.0000 0.2125
-4.500 -0.5020 0.04715 0.04230 0.0001 1.0000 0.2344
-4.000 -0.4512 0.03596 0.02873 0.0014 1.0000 0.1167
-3.750 -0.4267 0.03282 0.02498 0.0033 1.0000 0.0982
-3.500 -0.4041 0.03027 0.02217 0.0043 1.0000 0.0945
-3.250 -0.3793 0.02814 0.01954 0.0055 1.0000 0.0896
-3.000 -0.3558 0.02678 0.01789 0.0065 1.0000 0.0900
-2.750 -0.3325 0.02572 0.01665 0.0073 1.0000 0.0922
-2.500 -0.2915 0.02449 0.01518 0.0051 0.9956 0.0930
-2.250 -0.2490 0.02366 0.01417 0.0024 0.9898 0.0972
-2.000 -0.2053 0.02286 0.01327 -0.0005 0.9845 0.1009
-1.750 -0.1689 0.02198 0.01255 -0.0023 0.9787 0.1059
-1.500 -0.1276 0.02150 0.01211 -0.0050 0.9727 0.1163
-1.250 0.0198 0.01782 0.01170 -0.0257 0.9763 1.0000
-1.000 0.0965 0.01783 0.01132 -0.0350 0.9618 1.0000
-0.750 0.1638 0.01770 0.01096 -0.0425 0.9490 1.0000
-0.500 0.2208 0.01757 0.01068 -0.0480 0.9379 1.0000
-0.250 0.2707 0.01740 0.01041 -0.0520 0.9270 1.0000
0.000 0.3160 0.01716 0.01008 -0.0548 0.9151 1.0000
0.250 0.3472 0.01699 0.00986 -0.0548 0.9004 1.0000
0.500 0.3749 0.01679 0.00961 -0.0539 0.8852 1.0000
0.750 0.3993 0.01655 0.00934 -0.0522 0.8694 1.0000
1.000 0.4213 0.01629 0.00903 -0.0499 0.8532 1.0000
1.250 0.4410 0.01603 0.00875 -0.0472 0.8348 1.0000
1.500 0.4604 0.01574 0.00843 -0.0444 0.8149 1.0000
1.750 0.4802 0.01536 0.00801 -0.0414 0.7943 1.0000
2.000 0.4997 0.01501 0.00760 -0.0384 0.7697 1.0000
2.250 0.5193 0.01467 0.00720 -0.0355 0.7392 1.0000
2.500 0.5390 0.01437 0.00679 -0.0326 0.6960 1.0000
2.750 0.5580 0.01418 0.00635 -0.0296 0.6206 1.0000
3.000 0.5755 0.01459 0.00602 -0.0266 0.5083 1.0000
3.250 0.5947 0.01544 0.00624 -0.0248 0.4392 1.0000
3.500 0.6160 0.01618 0.00659 -0.0237 0.4005 1.0000
3.750 0.6382 0.01685 0.00699 -0.0227 0.3744 1.0000
4.000 0.6611 0.01750 0.00744 -0.0218 0.3541 1.0000
4.250 0.6842 0.01814 0.00792 -0.0210 0.3372 1.0000
4.500 0.7075 0.01880 0.00845 -0.0202 0.3230 1.0000
4.750 0.7313 0.01949 0.00899 -0.0196 0.3110 1.0000
5.000 0.7549 0.02011 0.00963 -0.0189 0.2999 1.0000
5.250 0.7790 0.02085 0.01033 -0.0182 0.2908 1.0000
5.500 0.8029 0.02155 0.01098 -0.0176 0.2820 1.0000
5.750 0.8265 0.02232 0.01177 -0.0169 0.2738 1.0000
6.000 0.8503 0.02305 0.01246 -0.0163 0.2662 1.0000
6.250 0.8737 0.02391 0.01340 -0.0157 0.2596 1.0000
6.500 0.8970 0.02472 0.01429 -0.0150 0.2534 1.0000
6.750 0.9217 0.02575 0.01522 -0.0146 0.2484 1.0000
7.000 0.9432 0.02669 0.01645 -0.0137 0.2432 1.0000
7.250 0.9656 0.02758 0.01745 -0.0129 0.2377 1.0000
7.500 0.9896 0.02867 0.01845 -0.0125 0.2329 1.0000
7.750 1.0089 0.02981 0.01993 -0.0113 0.2283 1.0000
8.000 1.0298 0.03084 0.02114 -0.0104 0.2232 1.0000
8.250 1.0532 0.03179 0.02200 -0.0099 0.2180 1.0000
8.500 1.0690 0.03298 0.02355 -0.0084 0.2120 1.0000
8.750 1.0895 0.03384 0.02449 -0.0075 0.2063 1.0000
9.000 1.1119 0.03502 0.02562 -0.0070 0.2016 1.0000
9.250 1.1234 0.03649 0.02756 -0.0050 0.1963 1.0000
9.500 1.1426 0.03756 0.02874 -0.0040 0.1914 1.0000
9.750 1.1664 0.03885 0.02992 -0.0037 0.1872 1.0000
10.000 1.1694 0.04104 0.03271 -0.0009 0.1825 1.0000
10.250 1.1825 0.04257 0.03446 0.0007 0.1775 1.0000
10.500 1.2103 0.04333 0.03501 0.0006 0.1726 1.0000
10.750 1.2017 0.04626 0.03856 0.0042 0.1677 1.0000
11.000 1.2132 0.04756 0.04001 0.0059 0.1622 1.0000
11.250 1.2392 0.04830 0.04062 0.0060 0.1572 1.0000
11.500 1.2161 0.05212 0.04503 0.0104 0.1535 1.0000
11.750 1.2127 0.05432 0.04746 0.0130 0.1490 1.0000
12.000 1.2551 0.05343 0.04628 0.0121 0.1430 1.0000
12.250 1.2188 0.05794 0.05128 0.0168 0.1413 1.0000
12.500 1.1761 0.06272 0.05634 0.0211 0.1407 1.0000
12.750 1.1231 0.06963 0.06349 0.0226 0.1411 1.0000
13.000 1.0572 0.08034 0.07434 0.0198 0.1424 1.0000
13.250 1.1773 0.06862 0.06244 0.0247 0.1305 1.0000
13.500 1.1411 0.07505 0.06903 0.0245 0.1299 1.0000
13.750 0.5643 0.14472 0.13877 -0.0084 0.1986 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-12 AIRFOIL (vr12-il)