NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 200,000 Max Cl/Cd: 54.51 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64110-il-200000.txt Download as CSV file: xf-n64110-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-110 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6359 0.08386 0.08039 -0.0245 1.0000 0.0535
-9.500 -0.6542 0.07846 0.07495 -0.0288 1.0000 0.0535
-9.250 -0.6758 0.07463 0.07103 -0.0303 1.0000 0.0536
-9.000 -0.6966 0.07190 0.06806 -0.0306 1.0000 0.0541
-8.750 -0.7055 0.06921 0.06509 -0.0304 1.0000 0.0543
-8.500 -0.7063 0.06021 0.05619 -0.0316 1.0000 0.0560
-8.250 -0.6899 0.05746 0.05357 -0.0314 1.0000 0.0584
-8.000 -0.6831 0.05466 0.05066 -0.0311 1.0000 0.0620
-7.500 -0.6360 0.03501 0.03093 -0.0310 1.0000 0.0716
-7.250 -0.6254 0.03187 0.02765 -0.0304 1.0000 0.0758
-7.000 -0.6209 0.02783 0.02319 -0.0295 1.0000 0.0835
-6.750 -0.6054 0.02513 0.02049 -0.0289 1.0000 0.0871
-6.500 -0.5961 0.02232 0.01736 -0.0277 1.0000 0.0976
-6.250 -0.5967 0.02803 0.02096 -0.0217 1.0000 0.0425
-6.000 -0.5778 0.02578 0.01848 -0.0202 1.0000 0.0421
-5.750 -0.5596 0.02328 0.01577 -0.0186 1.0000 0.0419
-5.500 -0.5418 0.02184 0.01413 -0.0168 1.0000 0.0422
-5.250 -0.5247 0.01976 0.01191 -0.0149 1.0000 0.0430
-5.000 -0.5089 0.01818 0.01032 -0.0131 1.0000 0.0451
-4.750 -0.4933 0.01743 0.00958 -0.0113 1.0000 0.0482
-4.500 -0.4728 0.01656 0.00868 -0.0103 0.9991 0.0505
-4.250 -0.4331 0.01563 0.00768 -0.0129 0.9936 0.0541
-4.000 -0.3965 0.01438 0.00650 -0.0154 0.9872 0.0625
-3.750 -0.3570 0.01346 0.00557 -0.0183 0.9815 0.0765
-3.500 -0.3255 0.01116 0.00442 -0.0208 0.9740 0.2966
-3.250 -0.2937 0.00998 0.00450 -0.0221 0.9677 0.6084
-3.000 -0.2574 0.00997 0.00455 -0.0235 0.9603 0.6596
-2.750 -0.2196 0.01000 0.00458 -0.0252 0.9541 0.6901
-2.500 -0.1879 0.01016 0.00477 -0.0253 0.9452 0.7270
-2.250 -0.1601 0.01046 0.00516 -0.0241 0.9363 0.7610
-2.000 -0.1294 0.01053 0.00521 -0.0240 0.9283 0.7771
-1.750 -0.1021 0.01051 0.00514 -0.0237 0.9175 0.7877
-1.500 -0.0747 0.01049 0.00507 -0.0234 0.9076 0.7986
-1.250 -0.0479 0.01046 0.00500 -0.0227 0.8990 0.8067
-1.000 -0.0227 0.01043 0.00494 -0.0220 0.8879 0.8162
-0.750 0.0026 0.01041 0.00490 -0.0213 0.8779 0.8258
-0.500 0.0278 0.01038 0.00484 -0.0203 0.8694 0.8346
-0.250 0.0528 0.01037 0.00481 -0.0196 0.8590 0.8447
0.000 0.0775 0.01037 0.00481 -0.0187 0.8495 0.8548
0.250 0.1020 0.01035 0.00479 -0.0176 0.8412 0.8644
0.500 0.1266 0.01036 0.00481 -0.0168 0.8311 0.8750
0.750 0.1513 0.01037 0.00483 -0.0160 0.8227 0.8864
1.000 0.1755 0.01037 0.00485 -0.0148 0.8140 0.8971
1.250 0.2001 0.01038 0.00490 -0.0139 0.8048 0.9084
1.500 0.2258 0.01037 0.00490 -0.0131 0.7976 0.9200
1.750 0.2524 0.01038 0.00496 -0.0128 0.7877 0.9321
2.000 0.2814 0.01037 0.00498 -0.0128 0.7784 0.9442
2.250 0.3150 0.01035 0.00500 -0.0137 0.7692 0.9546
2.500 0.3524 0.01033 0.00504 -0.0156 0.7575 0.9633
2.750 0.3875 0.01002 0.00468 -0.0162 0.7295 0.9725
3.000 0.4253 0.00974 0.00432 -0.0176 0.6902 0.9806
3.250 0.4667 0.00967 0.00426 -0.0202 0.6639 0.9879
3.500 0.5064 0.00963 0.00418 -0.0226 0.6232 0.9970
3.750 0.5287 0.00970 0.00414 -0.0217 0.5678 1.0000
4.000 0.5357 0.01037 0.00412 -0.0181 0.3956 1.0000
4.250 0.5313 0.01306 0.00512 -0.0140 0.0989 1.0000
4.500 0.5497 0.01401 0.00592 -0.0128 0.0721 1.0000
4.750 0.5711 0.01482 0.00672 -0.0119 0.0611 1.0000
5.000 0.5909 0.01599 0.00784 -0.0109 0.0551 1.0000
5.250 0.6140 0.01684 0.00875 -0.0101 0.0505 1.0000
5.500 0.6364 0.01792 0.00976 -0.0095 0.0459 1.0000
5.750 0.6595 0.01967 0.01153 -0.0087 0.0435 1.0000
6.000 0.6852 0.02093 0.01290 -0.0082 0.0422 1.0000
6.250 0.7113 0.02253 0.01465 -0.0077 0.0413 1.0000
6.500 0.7371 0.02428 0.01661 -0.0072 0.0402 1.0000
6.750 0.7617 0.02584 0.01836 -0.0067 0.0382 1.0000
7.000 0.7855 0.02833 0.02118 -0.0059 0.0383 1.0000
7.250 0.8068 0.03189 0.02521 -0.0046 0.0400 1.0000
7.500 0.8232 0.03699 0.03065 -0.0036 0.0432 1.0000
8.250 0.8688 0.05525 0.05020 0.0008 0.0697 1.0000
8.500 0.8067 0.04766 0.04374 0.0057 0.0670 1.0000
8.750 0.7961 0.05150 0.04790 0.0073 0.0636 1.0000
9.000 0.7849 0.05567 0.05223 0.0083 0.0620 1.0000
9.250 0.7643 0.05962 0.05629 0.0097 0.0616 1.0000
9.500 0.7347 0.06456 0.06135 0.0091 0.0619 1.0000
9.750 0.6982 0.07183 0.06872 0.0055 0.0636 1.0000
10.000 0.6682 0.08065 0.07759 0.0001 0.0653 1.0000
10.250 0.6462 0.08995 0.08688 -0.0057 0.0659 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 64-110 AIRFOIL (n64110-il)