NACA 11-H-09 AIRFOIL (n11h9-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 11-H-09 AIRFOIL (n11h9-il) Reynolds number: 100,000 Max Cl/Cd: 52.57 at α=8.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n11h9-il-100000.txt Download as CSV file: xf-n11h9-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 11-H-09 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3381 0.10597 0.10181 -0.0287 0.8899 0.0433
-8.500 -0.3399 0.10444 0.10029 -0.0317 0.8804 0.0435
-8.250 -0.3386 0.10273 0.09856 -0.0340 0.8716 0.0436
-8.000 -0.3345 0.10082 0.09662 -0.0365 0.8636 0.0437
-7.750 -0.3132 0.09105 0.08695 -0.0308 0.8593 0.0458
-7.500 -0.3081 0.08816 0.08405 -0.0310 0.8522 0.0469
-7.250 -0.3033 0.08553 0.08140 -0.0320 0.8454 0.0481
-7.000 -0.2964 0.08273 0.07859 -0.0338 0.8390 0.0495
-6.750 -0.2893 0.08011 0.07595 -0.0354 0.8326 0.0510
-6.500 -0.2806 0.07754 0.07333 -0.0373 0.8267 0.0528
-6.250 -0.2680 0.07554 0.07120 -0.0404 0.8199 0.0549
-6.000 -0.2536 0.07575 0.07107 -0.0433 0.8144 0.0562
-5.750 -0.2414 0.07054 0.06590 -0.0443 0.8086 0.0571
-5.500 -0.2325 0.06618 0.06164 -0.0431 0.8040 0.0592
-5.250 -0.2203 0.06359 0.05897 -0.0428 0.7999 0.0617
-5.000 -0.2014 0.06106 0.05630 -0.0443 0.7939 0.0660
-4.750 -0.1806 0.06006 0.05487 -0.0452 0.7889 0.0703
-4.500 -0.1708 0.05602 0.05098 -0.0441 0.7850 0.0740
-4.250 -0.1462 0.05540 0.04993 -0.0453 0.7793 0.0836
-4.000 -0.1343 0.05154 0.04623 -0.0446 0.7753 0.0895
-3.750 -0.1167 0.04973 0.04414 -0.0436 0.7720 0.0989
-3.500 -0.0942 0.04868 0.04279 -0.0441 0.7664 0.1110
-2.750 -0.0426 0.04231 0.03609 -0.0417 0.7544 0.1552
-2.500 -0.0259 0.04061 0.03433 -0.0413 0.7496 0.1852
-2.250 -0.0117 0.03868 0.03242 -0.0397 0.7457 0.2220
-2.000 -0.0008 0.01870 0.01316 -0.0400 0.7338 0.2771
-1.750 0.0114 0.01704 0.01151 -0.0377 0.7302 0.3241
-1.500 0.0246 0.01546 0.00991 -0.0349 0.7277 0.3748
-1.250 0.0396 0.01484 0.00927 -0.0347 0.7222 0.3984
-1.000 0.1179 0.03339 0.02551 -0.0351 0.7275 0.1982
-0.750 0.1591 0.03251 0.02402 -0.0358 0.7221 0.1040
-0.500 0.1888 0.03156 0.02268 -0.0351 0.7179 0.0862
-0.250 0.2177 0.03078 0.02162 -0.0341 0.7148 0.0808
0.000 0.2448 0.03036 0.02101 -0.0336 0.7114 0.0826
0.250 0.2698 0.03022 0.02085 -0.0344 0.7055 0.0826
0.500 0.2969 0.02949 0.02019 -0.0341 0.7015 0.0848
0.750 0.3243 0.02902 0.01968 -0.0331 0.6987 0.0952
1.000 0.3466 0.02957 0.02023 -0.0341 0.6920 0.1065
1.250 0.3745 0.02932 0.02005 -0.0342 0.6873 0.1542
1.500 0.5342 0.02658 0.01877 -0.0580 0.6848 1.0000
1.750 0.5548 0.02766 0.01979 -0.0584 0.6779 1.0000
2.000 0.5758 0.02821 0.02027 -0.0576 0.6725 1.0000
2.250 0.5980 0.02822 0.02018 -0.0555 0.6692 1.0000
2.500 0.6155 0.02974 0.02172 -0.0563 0.6605 1.0000
2.750 0.6377 0.02989 0.02181 -0.0547 0.6560 1.0000
3.000 0.6555 0.03095 0.02288 -0.0543 0.6488 1.0000
3.250 0.6758 0.03145 0.02336 -0.0532 0.6425 1.0000
3.500 0.7004 0.03119 0.02308 -0.0511 0.6395 1.0000
3.750 0.7138 0.03282 0.02477 -0.0512 0.6286 1.0000
4.000 0.7394 0.03243 0.02437 -0.0491 0.6252 1.0000
4.250 0.7523 0.03395 0.02595 -0.0488 0.6142 1.0000
4.500 0.7797 0.03329 0.02534 -0.0466 0.6108 1.0000
4.750 0.7927 0.03465 0.02677 -0.0460 0.5993 1.0000
5.000 0.8220 0.03369 0.02583 -0.0438 0.5961 1.0000
5.250 0.8355 0.03490 0.02713 -0.0430 0.5844 1.0000
5.500 0.8664 0.03364 0.02591 -0.0406 0.5814 1.0000
5.750 0.8814 0.03458 0.02701 -0.0396 0.5694 1.0000
6.000 0.9068 0.03401 0.02652 -0.0377 0.5630 1.0000
6.250 0.9300 0.03370 0.02631 -0.0361 0.5545 1.0000
6.500 0.9601 0.03239 0.02509 -0.0339 0.5497 1.0000
6.750 0.9825 0.03209 0.02497 -0.0323 0.5395 1.0000
7.000 1.0040 0.03186 0.02488 -0.0307 0.5286 1.0000
7.250 1.0418 0.02907 0.02217 -0.0280 0.5237 1.0000
7.500 1.0672 0.02799 0.02127 -0.0261 0.5110 1.0000
7.750 1.0942 0.02665 0.02008 -0.0241 0.4977 1.0000
8.000 1.1247 0.02441 0.01790 -0.0217 0.4808 1.0000
8.250 1.1489 0.02276 0.01633 -0.0195 0.4515 1.0000
8.500 1.1672 0.02246 0.01621 -0.0180 0.4149 1.0000
8.750 1.1834 0.02251 0.01619 -0.0162 0.3687 1.0000
9.000 1.1903 0.02354 0.01689 -0.0140 0.3102 1.0000
9.250 1.1830 0.02580 0.01876 -0.0113 0.2448 1.0000
9.500 1.1646 0.02863 0.02118 -0.0080 0.1962 1.0000
9.750 1.1426 0.03161 0.02379 -0.0044 0.1587 1.0000
10.000 1.1241 0.03488 0.02666 -0.0017 0.1275 1.0000
10.250 1.1143 0.03771 0.02924 0.0006 0.1061 1.0000
10.500 1.1105 0.04012 0.03149 0.0026 0.0918 1.0000
10.750 1.1138 0.04202 0.03328 0.0047 0.0824 1.0000
11.000 1.1257 0.04340 0.03449 0.0071 0.0747 1.0000
11.250 1.1474 0.04447 0.03570 0.0092 0.0697 1.0000
11.500 1.1685 0.04581 0.03702 0.0109 0.0646 1.0000
11.750 1.1991 0.04767 0.03898 0.0124 0.0603 1.0000
12.000 1.2205 0.05001 0.04163 0.0137 0.0580 1.0000
12.250 1.2372 0.05283 0.04479 0.0149 0.0566 1.0000
12.500 1.2451 0.05555 0.04772 0.0161 0.0547 1.0000
12.750 1.2505 0.05842 0.05073 0.0171 0.0526 1.0000
13.000 1.2588 0.06290 0.05542 0.0178 0.0512 1.0000
13.250 1.2522 0.06707 0.05987 0.0188 0.0510 1.0000
13.500 1.2402 0.07090 0.06399 0.0197 0.0511 1.0000
13.750 1.2254 0.07482 0.06817 0.0201 0.0512 1.0000
14.000 1.2078 0.07923 0.07282 0.0200 0.0512 1.0000
14.250 1.1875 0.08345 0.07729 0.0194 0.0515 1.0000
14.500 1.1609 0.08853 0.08265 0.0180 0.0522 1.0000
14.750 1.1176 0.09601 0.09047 0.0143 0.0536 1.0000
15.000 1.0728 0.10558 0.10029 0.0089 0.0550 1.0000
15.250 1.0293 0.11702 0.11190 0.0020 0.0567 1.0000
15.500 0.9943 0.12902 0.12394 -0.0046 0.0598 1.0000
15.750 0.9751 0.13845 0.13337 -0.0092 0.0613 1.0000
16.000 0.9638 0.14653 0.14145 -0.0125 0.0620 1.0000
16.250 0.7110 0.15659 0.15166 -0.0142 0.0863 1.0000
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Polar data table (+)
Polar graphs
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