NACA 6-H-15 AIRFOIL (n6h15-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 6-H-15 AIRFOIL (n6h15-il) Reynolds number: 200,000 Max Cl/Cd: 64.8 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n6h15-il-200000.txt Download as CSV file: xf-n6h15-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 6-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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.2623 0.10538 0.10147 -0.0255 0.6824 0.0485
-10.250 -0.2787 0.10070 0.09685 -0.0309 0.6814 0.0496
-10.000 -0.2943 0.09568 0.09186 -0.0364 0.6803 0.0500
-9.750 -0.2805 0.09091 0.08711 -0.0347 0.6780 0.0511
-9.500 -0.2671 0.08815 0.08434 -0.0340 0.6758 0.0524
-9.250 -0.2642 0.08460 0.08078 -0.0350 0.6741 0.0537
-9.000 -0.2671 0.08049 0.07668 -0.0370 0.6727 0.0549
-8.750 -0.4082 0.07971 0.07550 -0.0426 0.6769 0.0506
-8.500 -0.3773 0.07650 0.07240 -0.0431 0.6745 0.0522
-8.250 -0.3810 0.07413 0.06999 -0.0415 0.6727 0.0530
-8.000 -0.3820 0.07159 0.06739 -0.0404 0.6711 0.0544
-7.750 -0.3845 0.06880 0.06449 -0.0391 0.6696 0.0561
-7.500 -0.4006 0.06728 0.06249 -0.0357 0.6683 0.0601
-7.250 -0.4063 0.06280 0.05781 -0.0337 0.6662 0.0612
-7.000 -0.3903 0.05910 0.05426 -0.0341 0.6634 0.0624
-6.750 -0.3773 0.05682 0.05197 -0.0334 0.6609 0.0639
-6.500 -0.3661 0.05469 0.04974 -0.0320 0.6589 0.0665
-6.250 -0.3690 0.05352 0.04785 -0.0270 0.6572 0.0732
-6.000 -0.3515 0.04957 0.04407 -0.0273 0.6556 0.0751
-5.750 -0.3350 0.04777 0.04226 -0.0265 0.6542 0.0784
-5.500 -0.3284 0.04662 0.04050 -0.0225 0.6529 0.0874
-5.250 -0.3100 0.04374 0.03770 -0.0223 0.6514 0.0895
-5.000 -0.2905 0.04220 0.03616 -0.0219 0.6487 0.0936
-4.750 -0.2768 0.04099 0.03456 -0.0197 0.6463 0.1032
-4.500 -0.2562 0.03899 0.03262 -0.0194 0.6443 0.1071
-4.250 -0.2311 0.03229 0.02434 -0.0138 0.6427 0.0608
-4.000 -0.2044 0.03020 0.02203 -0.0135 0.6411 0.0600
-3.750 -0.1776 0.02861 0.02013 -0.0131 0.6397 0.0610
-3.500 -0.1465 0.02697 0.01828 -0.0136 0.6384 0.0614
-3.250 -0.1135 0.02561 0.01681 -0.0145 0.6373 0.0626
-3.000 -0.0841 0.02503 0.01619 -0.0150 0.6363 0.0654
-2.750 -0.0526 0.02499 0.01621 -0.0168 0.6333 0.0674
-2.500 -0.0218 0.02489 0.01612 -0.0183 0.6308 0.0693
-2.250 0.0069 0.02484 0.01606 -0.0192 0.6287 0.0716
-2.000 0.0340 0.02455 0.01584 -0.0197 0.6271 0.0756
-1.750 0.0580 0.02450 0.01584 -0.0195 0.6256 0.0796
-1.500 0.0807 0.02448 0.01581 -0.0189 0.6244 0.0847
-1.250 0.1013 0.02426 0.01563 -0.0176 0.6233 0.0924
-1.000 0.4497 0.02322 0.01700 -0.0720 0.6240 1.0000
-0.750 0.4730 0.02330 0.01698 -0.0716 0.6229 1.0000
-0.500 0.4958 0.02332 0.01689 -0.0709 0.6218 1.0000
-0.250 0.5184 0.02337 0.01685 -0.0701 0.6206 1.0000
0.000 0.5456 0.02486 0.01848 -0.0727 0.6163 1.0000
0.250 0.5698 0.02570 0.01934 -0.0736 0.6135 1.0000
0.500 0.5931 0.02631 0.01994 -0.0739 0.6115 1.0000
0.750 0.6159 0.02688 0.02050 -0.0740 0.6100 1.0000
1.000 0.6383 0.02717 0.02077 -0.0736 0.6087 1.0000
1.250 0.6608 0.02724 0.02080 -0.0729 0.6075 1.0000
1.500 0.6833 0.02721 0.02072 -0.0719 0.6065 1.0000
1.750 0.7059 0.02722 0.02069 -0.0710 0.6057 1.0000
2.000 0.7161 0.03229 0.02605 -0.0754 0.5982 1.0000
2.250 0.7347 0.03300 0.02678 -0.0749 0.5953 1.0000
2.500 0.7548 0.03332 0.02710 -0.0740 0.5937 1.0000
2.750 0.7762 0.03337 0.02715 -0.0729 0.5924 1.0000
3.000 0.7992 0.03309 0.02685 -0.0717 0.5913 1.0000
3.250 0.8230 0.03272 0.02646 -0.0704 0.5905 1.0000
3.500 0.7955 0.03949 0.03344 -0.0704 0.5786 1.0000
3.750 0.8228 0.03875 0.03272 -0.0693 0.5773 1.0000
4.000 0.8607 0.03647 0.03040 -0.0680 0.5764 1.0000
4.250 0.9019 0.03333 0.02719 -0.0664 0.5758 1.0000
4.500 0.9374 0.03092 0.02473 -0.0649 0.5751 1.0000
4.750 0.9657 0.02978 0.02356 -0.0637 0.5742 1.0000
5.000 0.9492 0.03483 0.02886 -0.0628 0.5619 1.0000
5.250 0.9962 0.03107 0.02505 -0.0618 0.5616 1.0000
5.500 1.0345 0.02835 0.02226 -0.0607 0.5606 1.0000
5.750 1.0678 0.02641 0.02027 -0.0596 0.5594 1.0000
6.000 1.0665 0.02933 0.02345 -0.0585 0.5494 1.0000
6.250 1.1015 0.02734 0.02146 -0.0578 0.5478 1.0000
6.500 1.1387 0.02492 0.01897 -0.0570 0.5458 1.0000
6.750 1.1592 0.02486 0.01902 -0.0560 0.5406 1.0000
7.000 1.1809 0.02452 0.01879 -0.0551 0.5347 1.0000
7.250 1.2174 0.02224 0.01643 -0.0544 0.5304 1.0000
7.500 1.2359 0.02212 0.01647 -0.0533 0.5211 1.0000
7.750 1.2660 0.02061 0.01495 -0.0525 0.5126 1.0000
8.000 1.2885 0.01991 0.01432 -0.0513 0.4989 1.0000
8.250 1.3023 0.02013 0.01473 -0.0498 0.4771 1.0000
8.500 1.3147 0.02029 0.01491 -0.0479 0.4368 1.0000
8.750 1.2972 0.02176 0.01576 -0.0422 0.3533 1.0000
9.000 1.2621 0.02399 0.01799 -0.0349 0.3428 1.0000
9.250 1.2284 0.02619 0.02014 -0.0278 0.3322 1.0000
9.500 1.1952 0.02865 0.02253 -0.0212 0.3152 1.0000
10.000 1.1338 0.03381 0.02738 -0.0097 0.2450 1.0000
10.250 1.0996 0.03680 0.03012 -0.0041 0.2168 1.0000
10.500 1.0687 0.03973 0.03283 0.0010 0.1916 1.0000
10.750 1.0466 0.04224 0.03520 0.0052 0.1656 1.0000
11.000 1.0266 0.04465 0.03745 0.0092 0.1413 1.0000
11.250 1.0122 0.04664 0.03933 0.0129 0.1285 1.0000
11.500 0.9992 0.04847 0.04110 0.0165 0.1205 1.0000
11.750 0.9910 0.05003 0.04263 0.0197 0.1148 1.0000
12.000 0.9850 0.05169 0.04426 0.0223 0.1108 1.0000
12.250 0.9827 0.05335 0.04589 0.0245 0.1076 1.0000
12.500 0.9857 0.05475 0.04732 0.0261 0.1046 1.0000
12.750 0.9900 0.05614 0.04873 0.0275 0.1016 1.0000
13.000 0.9945 0.05757 0.05013 0.0288 0.0986 1.0000
13.250 1.0013 0.05869 0.05119 0.0302 0.0948 1.0000
13.500 1.0094 0.06000 0.05258 0.0310 0.0911 1.0000
13.750 1.0173 0.06129 0.05387 0.0319 0.0873 1.0000
14.000 1.0296 0.06182 0.05429 0.0333 0.0833 1.0000
14.250 1.0414 0.06277 0.05531 0.0341 0.0809 1.0000
14.500 1.0508 0.06409 0.05675 0.0347 0.0778 1.0000
14.750 1.0607 0.06529 0.05796 0.0354 0.0744 1.0000
15.000 1.0801 0.06513 0.05770 0.0369 0.0695 1.0000
15.250 1.0854 0.06705 0.05979 0.0372 0.0669 1.0000
15.500 1.0901 0.06902 0.06185 0.0375 0.0629 1.0000
15.750 1.1050 0.06956 0.06232 0.0387 0.0577 1.0000
16.000 1.1049 0.07222 0.06517 0.0387 0.0542 1.0000
16.250 1.1083 0.07438 0.06735 0.0388 0.0505 1.0000
16.500 1.1148 0.07612 0.06913 0.0393 0.0466 1.0000
16.750 1.1166 0.07865 0.07182 0.0393 0.0436 1.0000
17.000 1.1205 0.08086 0.07404 0.0392 0.0412 1.0000
17.250 1.1309 0.08216 0.07534 0.0399 0.0389 1.0000
17.500 1.1293 0.08524 0.07864 0.0396 0.0369 1.0000
17.750 1.1333 0.08755 0.08109 0.0396 0.0353 1.0000
18.000 1.1336 0.09037 0.08391 0.0388 0.0336 1.0000
18.250 1.1418 0.09197 0.08549 0.0391 0.0322 1.0000
18.500 1.1387 0.09546 0.08925 0.0387 0.0310 1.0000
18.750 1.1384 0.09852 0.09251 0.0384 0.0300 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 6-H-15 AIRFOIL (n6h15-il)