NACA 11-H-09 AIRFOIL (n11h9-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 11-H-09 AIRFOIL (n11h9-il) Reynolds number: 200,000 Max Cl/Cd: 79.63 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n11h9-il-200000.txt Download as CSV file: xf-n11h9-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3168 0.08931 0.08604 -0.0085 0.7590 0.0233
-7.750 -0.3198 0.08636 0.08309 -0.0109 0.7548 0.0233
-7.500 -0.3223 0.08341 0.08013 -0.0124 0.7496 0.0234
-7.250 -0.3209 0.08010 0.07678 -0.0141 0.7456 0.0234
-7.000 -0.3171 0.07663 0.07321 -0.0154 0.7425 0.0234
-6.750 -0.3122 0.06981 0.06646 -0.0147 0.7381 0.0239
-6.500 -0.3049 0.06525 0.06190 -0.0140 0.7339 0.0245
-6.250 -0.2986 0.06142 0.05802 -0.0143 0.7303 0.0249
-6.000 -0.2918 0.05773 0.05426 -0.0148 0.7275 0.0255
-5.750 -0.2815 0.05398 0.05049 -0.0162 0.7223 0.0261
-5.500 -0.2706 0.05027 0.04671 -0.0173 0.7182 0.0268
-5.250 -0.2584 0.04667 0.04300 -0.0181 0.7148 0.0277
-5.000 -0.2439 0.04314 0.03935 -0.0189 0.7113 0.0288
-4.750 -0.2252 0.03985 0.03596 -0.0201 0.7066 0.0302
-4.500 -0.1974 0.03824 0.03401 -0.0211 0.7027 0.0318
-4.250 -0.1766 0.03588 0.03136 -0.0206 0.6996 0.0321
-4.000 -0.1651 0.03038 0.02577 -0.0206 0.6971 0.0328
-3.750 -0.1511 0.02665 0.02208 -0.0207 0.6926 0.0339
-3.500 -0.1338 0.02394 0.01929 -0.0204 0.6887 0.0355
-3.250 -0.1140 0.02156 0.01673 -0.0198 0.6856 0.0378
-3.000 -0.0896 0.01984 0.01471 -0.0188 0.6831 0.0419
-2.750 -0.0649 0.01766 0.01215 -0.0180 0.6786 0.0446
-2.500 -0.0476 0.01528 0.00983 -0.0180 0.6751 0.0480
-2.250 -0.0182 0.01562 0.00964 -0.0163 0.6716 0.0563
-2.000 -0.0023 0.01210 0.00612 -0.0160 0.6696 0.0595
-1.750 0.0226 0.01137 0.00516 -0.0153 0.6659 0.0682
-1.500 0.0445 0.00981 0.00351 -0.0151 0.6621 0.0756
-1.250 0.0680 0.00889 0.00244 -0.0145 0.6589 0.0899
-1.000 0.0908 0.00805 0.00139 -0.0137 0.6562 0.1130
-0.750 0.1208 0.02327 0.01601 -0.0132 0.6582 0.1407
-0.500 0.1437 0.02179 0.01451 -0.0126 0.6556 0.1739
-0.250 0.1870 0.02028 0.01210 -0.0106 0.6534 0.0720
0.000 0.2193 0.01901 0.01071 -0.0104 0.6490 0.0525
0.250 0.2485 0.01803 0.00962 -0.0102 0.6456 0.0488
0.500 0.2757 0.01734 0.00888 -0.0097 0.6426 0.0503
0.750 0.3017 0.01685 0.00832 -0.0090 0.6401 0.0545
1.000 0.3265 0.01645 0.00792 -0.0083 0.6368 0.0572
1.250 0.3502 0.01612 0.00764 -0.0076 0.6324 0.0672
1.500 0.5050 0.01359 0.00714 -0.0343 0.6282 1.0000
1.750 0.5299 0.01362 0.00703 -0.0335 0.6252 1.0000
2.000 0.5554 0.01384 0.00723 -0.0333 0.6206 1.0000
2.250 0.5808 0.01400 0.00737 -0.0330 0.6161 1.0000
2.500 0.6059 0.01408 0.00738 -0.0324 0.6127 1.0000
2.750 0.6308 0.01413 0.00734 -0.0317 0.6100 1.0000
3.000 0.6563 0.01443 0.00771 -0.0317 0.6047 1.0000
3.250 0.6815 0.01457 0.00786 -0.0313 0.6003 1.0000
3.500 0.7066 0.01460 0.00785 -0.0307 0.5970 1.0000
3.750 0.7317 0.01476 0.00801 -0.0302 0.5929 1.0000
4.000 0.7568 0.01496 0.00828 -0.0300 0.5871 1.0000
4.250 0.7819 0.01491 0.00822 -0.0292 0.5828 1.0000
4.500 0.8069 0.01500 0.00833 -0.0287 0.5778 1.0000
4.750 0.8318 0.01508 0.00848 -0.0283 0.5713 1.0000
5.000 0.8571 0.01494 0.00830 -0.0274 0.5668 1.0000
5.250 0.8818 0.01511 0.00862 -0.0271 0.5592 1.0000
5.500 0.9071 0.01500 0.00851 -0.0264 0.5536 1.0000
5.750 0.9318 0.01509 0.00868 -0.0259 0.5465 1.0000
6.000 0.9570 0.01499 0.00863 -0.0252 0.5396 1.0000
6.250 0.9817 0.01497 0.00872 -0.0246 0.5308 1.0000
6.500 1.0070 0.01466 0.00839 -0.0237 0.5218 1.0000
6.750 1.0315 0.01454 0.00838 -0.0230 0.5098 1.0000
7.000 1.0560 0.01435 0.00828 -0.0223 0.4956 1.0000
7.250 1.0799 0.01411 0.00814 -0.0215 0.4742 1.0000
7.500 1.1036 0.01401 0.00809 -0.0208 0.4469 1.0000
7.750 1.1268 0.01415 0.00829 -0.0202 0.4172 1.0000
8.000 1.1478 0.01455 0.00858 -0.0195 0.3726 1.0000
8.250 1.1640 0.01557 0.00931 -0.0186 0.3142 1.0000
8.500 1.1761 0.01706 0.01049 -0.0176 0.2525 1.0000
8.750 1.1819 0.01905 0.01210 -0.0163 0.1839 1.0000
9.000 1.1760 0.02181 0.01434 -0.0143 0.1118 1.0000
9.250 1.1633 0.02455 0.01676 -0.0114 0.0755 1.0000
9.500 1.1516 0.02659 0.01873 -0.0077 0.0643 1.0000
9.750 1.1460 0.02864 0.02080 -0.0052 0.0569 1.0000
10.000 1.1407 0.03089 0.02304 -0.0031 0.0522 1.0000
10.250 1.1387 0.03301 0.02521 -0.0014 0.0497 1.0000
10.500 1.1418 0.03478 0.02710 0.0001 0.0467 1.0000
10.750 1.1446 0.03663 0.02900 0.0014 0.0441 1.0000
11.000 1.1454 0.03867 0.03104 0.0029 0.0418 1.0000
11.250 1.1486 0.04055 0.03289 0.0051 0.0400 1.0000
11.500 1.1589 0.04191 0.03437 0.0065 0.0390 1.0000
11.750 1.1710 0.04321 0.03577 0.0080 0.0379 1.0000
12.000 1.1834 0.04459 0.03727 0.0095 0.0364 1.0000
12.250 1.1942 0.04612 0.03892 0.0107 0.0348 1.0000
12.500 1.2046 0.04775 0.04064 0.0119 0.0334 1.0000
12.750 1.2194 0.04942 0.04239 0.0134 0.0322 1.0000
13.000 1.2394 0.05258 0.04575 0.0155 0.0304 1.0000
13.250 1.2364 0.05483 0.04822 0.0160 0.0297 1.0000
13.500 1.2368 0.05767 0.05132 0.0168 0.0292 1.0000
13.750 1.2343 0.06089 0.05479 0.0174 0.0288 1.0000
14.000 1.2278 0.06449 0.05865 0.0178 0.0283 1.0000
14.250 1.2188 0.06847 0.06288 0.0179 0.0281 1.0000
14.500 1.2068 0.07276 0.06741 0.0176 0.0279 1.0000
14.750 1.1941 0.07697 0.07182 0.0168 0.0272 1.0000
15.000 1.1760 0.08254 0.07764 0.0157 0.0276 1.0000
15.250 1.1588 0.08799 0.08331 0.0140 0.0274 1.0000
15.500 1.1412 0.09373 0.08923 0.0118 0.0270 1.0000
15.750 1.1171 0.10124 0.09695 0.0087 0.0279 1.0000
16.000 1.0966 0.10831 0.10420 0.0052 0.0277 1.0000
16.250 1.0727 0.11665 0.11271 0.0008 0.0278 1.0000
16.500 1.0452 0.12653 0.12274 -0.0045 0.0288 1.0000
16.750 1.0222 0.13596 0.13225 -0.0095 0.0297 1.0000
17.000 0.9987 0.14621 0.14254 -0.0143 0.0305 1.0000
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Polar data table (+)
Polar graphs
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