NACA 1-H-15 AIRFOIL (n1h15-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 1-H-15 AIRFOIL (n1h15-il) Reynolds number: 200,000 Max Cl/Cd: 31.12 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n1h15-il-200000-n5.txt Download as CSV file: xf-n1h15-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 1-H-15 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.2141 0.10828 0.10392 -0.0526 0.7207 0.0367
-10.250 -0.2185 0.10484 0.10048 -0.0558 0.7201 0.0388
-10.000 -0.2242 0.10054 0.09627 -0.0630 0.7171 0.0393
-9.750 -0.2213 0.09676 0.09253 -0.0669 0.7149 0.0394
-9.500 -0.2163 0.09285 0.08865 -0.0688 0.7133 0.0396
-9.250 -0.2032 0.09031 0.08610 -0.0684 0.7120 0.0401
-9.000 -0.1949 0.08769 0.08348 -0.0691 0.7108 0.0406
-8.750 -0.1907 0.08480 0.08059 -0.0706 0.7095 0.0410
-8.500 -0.1874 0.08196 0.07775 -0.0720 0.7084 0.0418
-8.250 -0.1894 0.07867 0.07446 -0.0743 0.7074 0.0424
-8.000 -0.1939 0.07565 0.07142 -0.0759 0.7064 0.0435
-7.750 -0.2180 0.07106 0.06667 -0.0798 0.7053 0.0466
-7.500 -0.2202 0.06686 0.06224 -0.0840 0.7010 0.0469
-7.250 -0.2139 0.06362 0.05884 -0.0854 0.6975 0.0470
-7.000 -0.2011 0.05923 0.05456 -0.0850 0.6960 0.0386
-6.750 -0.1929 0.05675 0.05200 -0.0845 0.6946 0.0376
-6.500 -0.1866 0.05352 0.04860 -0.0837 0.6934 0.0365
-6.250 -0.1803 0.05002 0.04485 -0.0824 0.6924 0.0359
-6.000 -0.1709 0.04712 0.04171 -0.0809 0.6916 0.0363
-5.750 -0.1609 0.04417 0.03848 -0.0791 0.6909 0.0364
-5.500 -0.1485 0.04255 0.03662 -0.0799 0.6842 0.0357
-5.250 -0.1383 0.04001 0.03373 -0.0780 0.6815 0.0353
-5.000 -0.1258 0.03776 0.03112 -0.0758 0.6798 0.0351
-4.750 -0.1109 0.03586 0.02889 -0.0738 0.6786 0.0351
-4.500 -0.0936 0.03433 0.02708 -0.0721 0.6777 0.0353
-4.250 -0.0741 0.03317 0.02569 -0.0708 0.6769 0.0361
-4.000 -0.0538 0.03191 0.02415 -0.0694 0.6763 0.0367
-3.750 -0.0479 0.03262 0.02474 -0.0683 0.6687 0.0368
-3.500 -0.0291 0.03179 0.02368 -0.0670 0.6666 0.0368
-3.250 -0.0082 0.03098 0.02265 -0.0659 0.6652 0.0370
-3.000 0.0144 0.03025 0.02174 -0.0650 0.6641 0.0373
-2.750 0.0382 0.02958 0.02090 -0.0643 0.6633 0.0376
-2.500 0.0629 0.02896 0.02015 -0.0638 0.6625 0.0379
-2.250 0.0880 0.02842 0.01950 -0.0633 0.6619 0.0384
-2.000 0.0786 0.02987 0.02093 -0.0593 0.6533 0.0386
-1.750 0.0983 0.02976 0.02073 -0.0582 0.6516 0.0392
-1.500 0.1200 0.02938 0.02034 -0.0574 0.6504 0.0402
-1.250 0.1423 0.02907 0.02004 -0.0566 0.6493 0.0410
-1.000 0.1653 0.02878 0.01973 -0.0558 0.6485 0.0416
-0.750 0.1890 0.02847 0.01942 -0.0551 0.6478 0.0423
-0.500 0.1624 0.03006 0.02100 -0.0476 0.6388 0.0424
-0.250 0.1807 0.03007 0.02100 -0.0462 0.6371 0.0431
0.000 0.2018 0.02998 0.02090 -0.0452 0.6358 0.0439
0.250 0.2242 0.02986 0.02076 -0.0443 0.6347 0.0449
0.500 0.2480 0.02972 0.02058 -0.0436 0.6339 0.0460
0.750 0.2730 0.02952 0.02037 -0.0431 0.6332 0.0478
1.000 0.2981 0.02936 0.02022 -0.0426 0.6326 0.0514
1.500 0.5953 0.02921 0.02291 -0.0945 0.6349 1.0000
1.750 0.6147 0.02929 0.02293 -0.0931 0.6339 1.0000
2.000 0.6350 0.02935 0.02292 -0.0919 0.6332 1.0000
2.250 0.6569 0.02934 0.02287 -0.0908 0.6325 1.0000
2.750 0.6271 0.03192 0.02541 -0.0789 0.6200 1.0000
3.250 0.6260 0.03407 0.02752 -0.0718 0.6100 1.0000
3.500 0.6436 0.03437 0.02780 -0.0704 0.6080 1.0000
3.750 0.6648 0.03448 0.02789 -0.0695 0.6065 1.0000
4.000 0.6874 0.03453 0.02792 -0.0687 0.6054 1.0000
4.250 0.7108 0.03456 0.02794 -0.0680 0.6045 1.0000
4.500 0.7346 0.03458 0.02797 -0.0674 0.6038 1.0000
5.000 0.7358 0.03713 0.03054 -0.0614 0.5918 1.0000
5.250 0.7593 0.03717 0.03058 -0.0608 0.5907 1.0000
5.500 0.7838 0.03714 0.03057 -0.0603 0.5898 1.0000
5.750 0.8089 0.03708 0.03055 -0.0599 0.5891 1.0000
6.250 0.8158 0.03955 0.03308 -0.0550 0.5767 1.0000
6.500 0.8472 0.03878 0.03234 -0.0549 0.5754 1.0000
6.750 0.8744 0.03825 0.03186 -0.0545 0.5733 1.0000
8.250 0.9824 0.03773 0.03160 -0.0468 0.5301 1.0000
8.750 1.0109 0.03847 0.03241 -0.0439 0.4988 1.0000
9.000 1.0499 0.03654 0.03039 -0.0440 0.4652 1.0000
9.250 1.0757 0.03457 0.02732 -0.0417 0.3325 1.0000
9.500 1.0435 0.03849 0.03060 -0.0364 0.2434 1.0000
9.750 1.0248 0.04173 0.03338 -0.0328 0.1749 1.0000
10.000 1.0192 0.04411 0.03550 -0.0304 0.1304 1.0000
10.250 1.0183 0.04617 0.03734 -0.0284 0.0971 1.0000
10.500 1.0219 0.04794 0.03900 -0.0268 0.0771 1.0000
10.750 1.0284 0.04951 0.04052 -0.0254 0.0669 1.0000
11.000 1.0362 0.05099 0.04197 -0.0241 0.0612 1.0000
11.250 1.0445 0.05245 0.04347 -0.0230 0.0571 1.0000
11.500 1.0532 0.05391 0.04497 -0.0219 0.0542 1.0000
11.750 1.0608 0.05548 0.04654 -0.0208 0.0516 1.0000
12.000 1.0678 0.05713 0.04821 -0.0197 0.0498 1.0000
12.250 1.0769 0.05859 0.04975 -0.0187 0.0484 1.0000
12.500 1.0860 0.06008 0.05132 -0.0178 0.0468 1.0000
12.750 1.0946 0.06161 0.05290 -0.0169 0.0453 1.0000
13.000 1.1029 0.06319 0.05451 -0.0160 0.0439 1.0000
13.250 1.1098 0.06489 0.05623 -0.0151 0.0426 1.0000
13.500 1.1163 0.06659 0.05794 -0.0140 0.0415 1.0000
13.750 1.1269 0.06795 0.05941 -0.0133 0.0403 1.0000
14.000 1.1371 0.06935 0.06089 -0.0125 0.0393 1.0000
14.250 1.1469 0.07078 0.06237 -0.0118 0.0381 1.0000
14.500 1.1565 0.07224 0.06388 -0.0111 0.0370 1.0000
14.750 1.1658 0.07374 0.06542 -0.0104 0.0359 1.0000
15.000 1.1747 0.07519 0.06686 -0.0096 0.0347 1.0000
15.250 1.1853 0.07646 0.06819 -0.0089 0.0338 1.0000
15.500 1.1959 0.07784 0.06970 -0.0082 0.0329 1.0000
15.750 1.2066 0.07919 0.07114 -0.0076 0.0320 1.0000
16.000 1.2156 0.08079 0.07283 -0.0071 0.0307 1.0000
16.250 1.2250 0.08232 0.07443 -0.0066 0.0299 1.0000
16.500 1.2330 0.08402 0.07616 -0.0061 0.0290 1.0000
16.750 1.2432 0.08530 0.07743 -0.0054 0.0281 1.0000
17.000 1.2507 0.08716 0.07948 -0.0050 0.0274 1.0000
17.250 1.2575 0.08911 0.08159 -0.0046 0.0265 1.0000
17.500 1.2633 0.09119 0.08382 -0.0043 0.0257 1.0000
17.750 1.2684 0.09338 0.08613 -0.0041 0.0249 1.0000
18.000 1.2725 0.09572 0.08855 -0.0041 0.0242 1.0000
18.250 1.2771 0.09794 0.09082 -0.0040 0.0236 1.0000
18.500 1.2825 0.09999 0.09290 -0.0039 0.0231 1.0000
18.750 1.2823 0.10305 0.09621 -0.0039 0.0225 1.0000
19.000 1.2819 0.10616 0.09955 -0.0040 0.0220 1.0000
19.250 1.2798 0.10955 0.10315 -0.0043 0.0213 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 1-H-15 AIRFOIL (n1h15-il)