BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il) Reynolds number: 200,000 Max Cl/Cd: 54.23 at α=9.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-v23010-il-200000.txt Download as CSV file: xf-v23010-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL V23010-1.58 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5227 0.08654 0.08283 -0.0187 1.0000 0.0518
-8.750 -0.5489 0.08082 0.07705 -0.0252 1.0000 0.0521
-8.500 -0.5716 0.07755 0.07362 -0.0261 1.0000 0.0523
-8.250 -0.5875 0.07484 0.07065 -0.0261 1.0000 0.0525
-8.000 -0.5832 0.06872 0.06454 -0.0264 1.0000 0.0531
-7.750 -0.5678 0.06458 0.06055 -0.0262 1.0000 0.0538
-7.500 -0.5582 0.06155 0.05754 -0.0256 1.0000 0.0545
-7.250 -0.5503 0.05871 0.05468 -0.0250 1.0000 0.0555
-7.000 -0.5432 0.05588 0.05177 -0.0242 1.0000 0.0568
-6.750 -0.5356 0.05310 0.04885 -0.0233 1.0000 0.0586
-6.500 -0.5338 0.05408 0.04903 -0.0201 1.0000 0.0626
-6.250 -0.5265 0.04797 0.04294 -0.0195 1.0000 0.0638
-6.000 -0.5125 0.04460 0.03968 -0.0187 1.0000 0.0649
-5.750 -0.4988 0.04230 0.03741 -0.0174 1.0000 0.0668
-5.500 -0.4854 0.04043 0.03544 -0.0158 1.0000 0.0697
-5.250 -0.4767 0.04001 0.03437 -0.0124 1.0000 0.0758
-5.000 -0.4631 0.03641 0.03092 -0.0114 1.0000 0.0775
-4.750 -0.4495 0.03443 0.02898 -0.0097 1.0000 0.0798
-4.500 -0.4363 0.03312 0.02756 -0.0075 1.0000 0.0838
-4.250 -0.4270 0.03185 0.02595 -0.0047 1.0000 0.0910
-4.000 -0.4149 0.02990 0.02407 -0.0029 1.0000 0.0938
-3.750 -0.3779 0.02814 0.02206 -0.0056 0.9942 0.1070
-3.500 -0.3347 0.02593 0.01975 -0.0096 0.9855 0.1236
-3.250 -0.2948 0.02396 0.01775 -0.0128 0.9764 0.1430
-3.000 -0.2545 0.02243 0.01602 -0.0156 0.9674 0.1690
-2.750 -0.2152 0.02029 0.01393 -0.0182 0.9589 0.1876
-2.500 -0.1517 0.01906 0.01142 -0.0171 0.9527 0.0638
-2.250 -0.1115 0.01755 0.00993 -0.0190 0.9425 0.0604
-2.000 -0.0749 0.01629 0.00864 -0.0201 0.9285 0.0586
-1.750 -0.0392 0.01530 0.00762 -0.0209 0.9121 0.0579
-1.500 -0.0057 0.01458 0.00688 -0.0214 0.8933 0.0594
-1.250 0.0233 0.01401 0.00628 -0.0210 0.8710 0.0610
-1.000 0.0497 0.01346 0.00569 -0.0200 0.8474 0.0617
-0.750 0.0742 0.01304 0.00522 -0.0188 0.8220 0.0626
-0.500 0.0967 0.01243 0.00457 -0.0172 0.7961 0.0649
-0.250 0.1203 0.01210 0.00414 -0.0159 0.7695 0.0686
0.000 0.1449 0.01196 0.00382 -0.0147 0.7427 0.0748
0.250 0.1686 0.01169 0.00347 -0.0133 0.7136 0.0938
0.500 0.1686 0.00902 0.00355 -0.0069 0.6874 0.8431
0.750 0.2441 0.00949 0.00382 -0.0141 0.6165 0.9638
1.000 0.3349 0.01026 0.00384 -0.0261 0.4937 0.9970
1.250 0.3612 0.01057 0.00373 -0.0262 0.4346 1.0000
1.500 0.3796 0.01082 0.00371 -0.0247 0.4035 1.0000
1.750 0.3989 0.01106 0.00374 -0.0233 0.3827 1.0000
2.000 0.4189 0.01129 0.00380 -0.0220 0.3667 1.0000
2.250 0.4396 0.01150 0.00389 -0.0207 0.3534 1.0000
2.500 0.4609 0.01170 0.00400 -0.0195 0.3417 1.0000
2.750 0.4824 0.01196 0.00413 -0.0184 0.3320 1.0000
3.000 0.5046 0.01216 0.00429 -0.0173 0.3238 1.0000
3.250 0.5270 0.01247 0.00448 -0.0163 0.3172 1.0000
3.500 0.5500 0.01268 0.00468 -0.0153 0.3101 1.0000
3.750 0.5728 0.01299 0.00490 -0.0143 0.3037 1.0000
4.000 0.5962 0.01325 0.00515 -0.0134 0.2970 1.0000
4.250 0.6195 0.01351 0.00538 -0.0125 0.2904 1.0000
4.500 0.6430 0.01390 0.00569 -0.0117 0.2846 1.0000
4.750 0.6667 0.01415 0.00599 -0.0109 0.2782 1.0000
5.000 0.6904 0.01451 0.00628 -0.0101 0.2722 1.0000
5.250 0.7141 0.01488 0.00668 -0.0093 0.2660 1.0000
5.500 0.7378 0.01519 0.00701 -0.0084 0.2593 1.0000
5.750 0.7615 0.01571 0.00744 -0.0077 0.2529 1.0000
6.000 0.7849 0.01595 0.00780 -0.0068 0.2453 1.0000
6.250 0.8086 0.01640 0.00817 -0.0061 0.2384 1.0000
6.500 0.8318 0.01675 0.00865 -0.0052 0.2310 1.0000
6.750 0.8555 0.01713 0.00900 -0.0045 0.2243 1.0000
7.000 0.8790 0.01765 0.00958 -0.0037 0.2177 1.0000
7.250 0.9024 0.01798 0.00997 -0.0029 0.2107 1.0000
7.500 0.9259 0.01847 0.01042 -0.0022 0.2045 1.0000
7.750 0.9487 0.01865 0.01073 -0.0013 0.1971 1.0000
8.000 0.9716 0.01891 0.01092 -0.0006 0.1902 1.0000
8.250 0.9939 0.01897 0.01114 0.0003 0.1823 1.0000
8.500 1.0168 0.01929 0.01139 0.0011 0.1764 1.0000
8.750 1.0393 0.01955 0.01185 0.0019 0.1700 1.0000
9.000 1.0616 0.01978 0.01210 0.0028 0.1641 1.0000
9.250 1.0837 0.02015 0.01259 0.0036 0.1578 1.0000
9.500 1.1052 0.02039 0.01286 0.0045 0.1511 1.0000
9.750 1.1264 0.02077 0.01337 0.0054 0.1438 1.0000
10.000 1.1461 0.02118 0.01374 0.0064 0.1360 1.0000
10.250 1.1660 0.02154 0.01428 0.0075 0.1265 1.0000
10.500 1.1839 0.02211 0.01492 0.0087 0.1151 1.0000
10.750 1.1992 0.02286 0.01571 0.0101 0.1008 1.0000
11.000 1.2113 0.02391 0.01669 0.0119 0.0860 1.0000
11.250 1.2237 0.02496 0.01776 0.0136 0.0745 1.0000
11.500 1.2341 0.02611 0.01893 0.0156 0.0672 1.0000
11.750 1.2412 0.02742 0.02031 0.0179 0.0626 1.0000
12.000 1.2447 0.02868 0.02161 0.0206 0.0591 1.0000
12.250 1.2429 0.03035 0.02326 0.0234 0.0565 1.0000
12.500 1.2463 0.03183 0.02491 0.0255 0.0542 1.0000
12.750 1.2474 0.03357 0.02674 0.0272 0.0522 1.0000
13.000 1.2472 0.03556 0.02879 0.0285 0.0507 1.0000
13.250 1.2445 0.03801 0.03123 0.0295 0.0491 1.0000
13.500 1.2442 0.04044 0.03379 0.0301 0.0479 1.0000
13.750 1.2433 0.04304 0.03656 0.0302 0.0467 1.0000
14.000 1.2416 0.04586 0.03952 0.0301 0.0455 1.0000
14.250 1.2400 0.04879 0.04254 0.0296 0.0444 1.0000
14.500 1.2388 0.05175 0.04556 0.0290 0.0433 1.0000
14.750 1.2394 0.05462 0.04842 0.0288 0.0422 1.0000
15.000 1.2398 0.05773 0.05159 0.0288 0.0414 1.0000
15.250 1.2322 0.06184 0.05591 0.0271 0.0410 1.0000
15.500 1.2226 0.06642 0.06071 0.0250 0.0405 1.0000
15.750 1.2124 0.07130 0.06577 0.0226 0.0401 1.0000
16.000 1.1997 0.07672 0.07138 0.0197 0.0397 1.0000
16.250 1.1858 0.08256 0.07740 0.0164 0.0393 1.0000
16.500 1.1699 0.08902 0.08404 0.0126 0.0390 1.0000
16.750 1.1506 0.09639 0.09160 0.0082 0.0390 1.0000
17.000 1.1256 0.10519 0.10060 0.0026 0.0390 1.0000
17.250 1.0913 0.11644 0.11208 -0.0046 0.0394 1.0000
17.500 1.0348 0.13341 0.12930 -0.0155 0.0405 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il)