NACA 642-015A AIRFOIL (n64015a-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 642-015A AIRFOIL (n64015a-il) Reynolds number: 50,000 Max Cl/Cd: 27.88 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64015a-il-50000-n5.txt Download as CSV file: xf-n64015a-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 642-015A 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
------ -------- --------- --------- -------- -------- --------
-13.250 -0.7529 0.09641 0.08854 -0.0323 1.0000 0.0566
-13.000 -0.7820 0.08859 0.08058 -0.0365 1.0000 0.0562
-12.750 -0.8089 0.08199 0.07381 -0.0394 1.0000 0.0559
-12.500 -0.8327 0.07641 0.06801 -0.0411 1.0000 0.0557
-12.250 -0.8531 0.07163 0.06301 -0.0418 1.0000 0.0557
-12.000 -0.8702 0.06745 0.05857 -0.0416 1.0000 0.0558
-11.750 -0.8837 0.06374 0.05457 -0.0406 1.0000 0.0561
-11.500 -0.8942 0.06043 0.05095 -0.0391 1.0000 0.0567
-11.250 -0.9022 0.05744 0.04762 -0.0371 1.0000 0.0575
-11.000 -0.9080 0.05475 0.04453 -0.0346 1.0000 0.0584
-10.750 -0.8946 0.05235 0.04214 -0.0339 1.0000 0.0602
-10.500 -0.8831 0.05022 0.03988 -0.0328 1.0000 0.0621
-10.250 -0.8691 0.04798 0.03742 -0.0318 1.0000 0.0641
-10.000 -0.8492 0.04568 0.03483 -0.0312 1.0000 0.0664
-9.750 -0.8273 0.04363 0.03244 -0.0307 1.0000 0.0698
-9.500 -0.8005 0.04194 0.03083 -0.0310 1.0000 0.0740
-9.250 -0.7701 0.04038 0.02907 -0.0311 1.0000 0.0792
-9.000 -0.7417 0.03901 0.02769 -0.0309 1.0000 0.0853
-8.750 -0.7191 0.03781 0.02637 -0.0301 1.0000 0.0928
-8.500 -0.7022 0.03652 0.02515 -0.0289 1.0000 0.1001
-8.250 -0.6886 0.03526 0.02386 -0.0272 1.0000 0.1096
-8.000 -0.6799 0.03394 0.02260 -0.0252 1.0000 0.1200
-7.750 -0.6752 0.03259 0.02136 -0.0226 1.0000 0.1324
-7.500 -0.6738 0.03126 0.02017 -0.0194 1.0000 0.1477
-7.250 -0.6745 0.02987 0.01899 -0.0159 1.0000 0.1688
-7.000 -0.6788 0.02844 0.01787 -0.0119 1.0000 0.1985
-6.750 -0.6868 0.02700 0.01686 -0.0071 1.0000 0.2424
-6.500 -0.6967 0.02566 0.01615 -0.0018 1.0000 0.3100
-6.250 -0.6995 0.02510 0.01641 0.0038 1.0000 0.4143
-6.000 -0.6972 0.02508 0.01655 0.0085 1.0000 0.4847
-5.750 -0.6937 0.02515 0.01661 0.0130 1.0000 0.5279
-5.500 -0.6873 0.02534 0.01675 0.0170 1.0000 0.5615
-5.250 -0.6798 0.02557 0.01691 0.0208 1.0000 0.5901
-5.000 -0.6627 0.02607 0.01732 0.0232 0.9969 0.6202
-4.750 -0.6263 0.02720 0.01833 0.0233 0.9881 0.6530
-4.500 -0.5870 0.02830 0.01929 0.0229 0.9803 0.6776
-4.250 -0.5481 0.02876 0.01954 0.0215 0.9724 0.6950
-4.000 -0.5107 0.02897 0.01955 0.0199 0.9635 0.7072
-3.750 -0.4765 0.02886 0.01925 0.0183 0.9543 0.7187
-3.500 -0.4376 0.02878 0.01897 0.0158 0.9465 0.7286
-3.250 -0.4032 0.02867 0.01870 0.0142 0.9372 0.7366
-3.000 -0.3644 0.02844 0.01830 0.0115 0.9301 0.7452
-2.750 -0.3328 0.02827 0.01800 0.0102 0.9203 0.7523
-2.500 -0.2934 0.02809 0.01769 0.0076 0.9135 0.7593
-2.250 -0.2682 0.02782 0.01731 0.0073 0.9030 0.7672
-2.000 -0.2276 0.02771 0.01711 0.0047 0.8968 0.7729
-1.750 -0.2078 0.02741 0.01672 0.0053 0.8857 0.7819
-1.500 -0.1663 0.02735 0.01658 0.0026 0.8800 0.7868
-1.250 -0.1452 0.02720 0.01638 0.0032 0.8694 0.7951
-1.000 -0.1098 0.02707 0.01619 0.0015 0.8632 0.8012
-0.750 -0.0849 0.02705 0.01615 0.0015 0.8535 0.8083
-0.500 -0.0538 0.02688 0.01593 0.0005 0.8473 0.8161
-0.250 -0.0279 0.02696 0.01602 0.0004 0.8378 0.8227
0.000 0.0000 0.02681 0.01584 0.0000 0.8314 0.8314
0.250 0.0279 0.02696 0.01602 -0.0004 0.8227 0.8378
0.500 0.0538 0.02688 0.01593 -0.0005 0.8161 0.8473
0.750 0.0849 0.02705 0.01615 -0.0015 0.8083 0.8535
1.000 0.1097 0.02707 0.01619 -0.0015 0.8012 0.8632
1.250 0.1453 0.02720 0.01638 -0.0032 0.7951 0.8694
1.500 0.1662 0.02735 0.01657 -0.0026 0.7868 0.8801
1.750 0.2078 0.02741 0.01672 -0.0053 0.7819 0.8857
2.000 0.2276 0.02771 0.01710 -0.0047 0.7730 0.8968
2.250 0.2682 0.02782 0.01731 -0.0073 0.7672 0.9030
2.500 0.2934 0.02809 0.01769 -0.0076 0.7593 0.9135
2.750 0.3328 0.02827 0.01800 -0.0102 0.7523 0.9204
3.000 0.3645 0.02844 0.01829 -0.0115 0.7453 0.9301
3.250 0.4032 0.02866 0.01870 -0.0142 0.7366 0.9373
3.500 0.4377 0.02878 0.01897 -0.0158 0.7286 0.9465
3.750 0.4765 0.02885 0.01925 -0.0183 0.7188 0.9544
4.000 0.5108 0.02896 0.01955 -0.0199 0.7072 0.9636
4.250 0.5481 0.02876 0.01953 -0.0215 0.6951 0.9725
4.500 0.5871 0.02829 0.01928 -0.0229 0.6776 0.9804
4.750 0.6264 0.02719 0.01832 -0.0233 0.6530 0.9881
5.000 0.6628 0.02606 0.01731 -0.0233 0.6202 0.9969
5.250 0.6796 0.02557 0.01691 -0.0208 0.5902 1.0000
5.500 0.6872 0.02533 0.01675 -0.0170 0.5616 1.0000
5.750 0.6936 0.02514 0.01660 -0.0129 0.5280 1.0000
6.000 0.6971 0.02508 0.01655 -0.0085 0.4849 1.0000
6.250 0.6994 0.02509 0.01641 -0.0038 0.4146 1.0000
6.500 0.6966 0.02565 0.01614 0.0018 0.3101 1.0000
6.750 0.6868 0.02700 0.01685 0.0071 0.2424 1.0000
7.000 0.6789 0.02844 0.01786 0.0119 0.1984 1.0000
7.250 0.6746 0.02987 0.01899 0.0159 0.1688 1.0000
7.500 0.6740 0.03126 0.02016 0.0194 0.1476 1.0000
7.750 0.6755 0.03259 0.02135 0.0225 0.1323 1.0000
8.000 0.6802 0.03394 0.02260 0.0251 0.1199 1.0000
8.250 0.6890 0.03526 0.02386 0.0272 0.1096 1.0000
8.500 0.7026 0.03652 0.02514 0.0288 0.1000 1.0000
8.750 0.7196 0.03781 0.02637 0.0300 0.0927 1.0000
9.000 0.7423 0.03901 0.02768 0.0308 0.0853 1.0000
9.250 0.7707 0.04038 0.02907 0.0310 0.0792 1.0000
9.500 0.8011 0.04195 0.03083 0.0309 0.0740 1.0000
9.750 0.8280 0.04364 0.03244 0.0306 0.0697 1.0000
10.000 0.8498 0.04569 0.03484 0.0311 0.0664 1.0000
10.250 0.8696 0.04799 0.03743 0.0317 0.0641 1.0000
10.500 0.8837 0.05023 0.03990 0.0327 0.0621 1.0000
10.750 0.8952 0.05237 0.04216 0.0338 0.0602 1.0000
11.000 0.9085 0.05477 0.04456 0.0345 0.0584 1.0000
11.250 0.9027 0.05747 0.04765 0.0370 0.0575 1.0000
11.500 0.8948 0.06046 0.05099 0.0390 0.0567 1.0000
11.750 0.8843 0.06378 0.05461 0.0405 0.0561 1.0000
12.000 0.8708 0.06749 0.05862 0.0414 0.0558 1.0000
12.250 0.8538 0.07168 0.06307 0.0416 0.0557 1.0000
12.500 0.8334 0.07648 0.06808 0.0409 0.0557 1.0000
12.750 0.8096 0.08207 0.07389 0.0392 0.0559 1.0000
13.000 0.7827 0.08869 0.08068 0.0363 0.0562 1.0000
13.250 0.7536 0.09653 0.08867 0.0320 0.0566 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 642-015A AIRFOIL (n64015a-il)