NACA 5-H-10 AIRFOIL (n5h10-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 5-H-10 AIRFOIL (n5h10-il) Reynolds number: 100,000 Max Cl/Cd: 48.44 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n5h10-il-100000.txt Download as CSV file: xf-n5h10-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 5-H-10 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4097 0.10541 0.10134 -0.0111 1.0000 0.0853
-9.750 -0.4216 0.10184 0.09782 -0.0155 1.0000 0.0869
-9.500 -0.4359 0.09800 0.09404 -0.0206 1.0000 0.0875
-9.250 -0.4048 0.09238 0.08839 -0.0166 1.0000 0.0925
-9.000 -0.4023 0.08832 0.08435 -0.0183 1.0000 0.0965
-8.750 -0.4114 0.08397 0.08005 -0.0224 1.0000 0.0999
-8.500 -0.4311 0.07935 0.07550 -0.0290 1.0000 0.1012
-8.250 -0.5021 0.08375 0.07965 -0.0226 1.0000 0.0940
-8.000 -0.5068 0.08021 0.07612 -0.0249 1.0000 0.0976
-7.750 -0.5232 0.07731 0.07310 -0.0296 1.0000 0.1013
-7.500 -0.5284 0.07320 0.06895 -0.0305 1.0000 0.1032
-7.250 -0.5135 0.06974 0.06562 -0.0283 1.0000 0.1073
-7.000 -0.5294 0.06848 0.06438 -0.0245 1.0000 0.1093
-6.750 -0.5273 0.06472 0.06027 -0.0284 0.9873 0.1178
-6.500 -0.4997 0.06055 0.05612 -0.0306 0.9806 0.1271
-6.250 -0.4770 0.05624 0.05167 -0.0336 0.9725 0.1386
-6.000 -0.4596 0.05336 0.04834 -0.0365 0.9627 0.1609
-5.750 -0.4313 0.04926 0.04438 -0.0382 0.9571 0.1799
-5.500 -0.4114 0.04663 0.04166 -0.0386 0.9490 0.2092
-5.250 -0.3462 0.02735 0.02282 -0.0418 0.9269 0.2574
-5.000 -0.3758 0.04186 0.03686 -0.0373 0.9337 0.2819
-4.750 -0.3246 0.03288 0.02538 -0.0393 0.9265 0.0986
-4.500 -0.3027 0.03039 0.02247 -0.0380 0.9182 0.0946
-4.250 -0.2714 0.02867 0.01992 -0.0370 0.9123 0.0872
-4.000 -0.2467 0.02674 0.01771 -0.0362 0.9051 0.0865
-3.750 -0.2177 0.02499 0.01572 -0.0363 0.8996 0.0905
-3.500 -0.1899 0.02377 0.01436 -0.0362 0.8933 0.0930
-3.250 -0.1598 0.02272 0.01319 -0.0363 0.8874 0.0959
-3.000 -0.1320 0.02214 0.01245 -0.0361 0.8818 0.1025
-2.750 -0.1059 0.02118 0.01164 -0.0360 0.8759 0.1085
-2.500 0.1400 0.01674 0.00992 -0.0705 0.8897 1.0000
-2.250 0.1623 0.01696 0.00992 -0.0698 0.8833 1.0000
-2.000 0.1856 0.01715 0.00996 -0.0697 0.8762 1.0000
-1.750 0.2068 0.01741 0.01007 -0.0687 0.8702 1.0000
-1.500 0.2308 0.01764 0.01018 -0.0688 0.8636 1.0000
-1.250 0.2533 0.01791 0.01034 -0.0682 0.8582 1.0000
-1.000 0.2766 0.01822 0.01057 -0.0681 0.8529 1.0000
-0.750 0.2996 0.01852 0.01081 -0.0678 0.8470 1.0000
-0.500 0.3211 0.01885 0.01106 -0.0669 0.8425 1.0000
-0.250 0.3456 0.01923 0.01142 -0.0673 0.8371 1.0000
0.000 0.3683 0.01960 0.01177 -0.0671 0.8323 1.0000
0.250 0.3899 0.01998 0.01211 -0.0661 0.8281 1.0000
0.500 0.4127 0.02044 0.01260 -0.0663 0.8219 1.0000
0.750 0.4341 0.02083 0.01298 -0.0655 0.8167 1.0000
1.000 0.4549 0.02128 0.01346 -0.0648 0.8103 1.0000
1.250 0.4752 0.02164 0.01383 -0.0637 0.8034 1.0000
1.500 0.4954 0.02210 0.01434 -0.0628 0.7971 1.0000
1.750 0.5155 0.02251 0.01480 -0.0618 0.7901 1.0000
2.000 0.5351 0.02291 0.01524 -0.0605 0.7834 1.0000
2.250 0.5546 0.02327 0.01566 -0.0592 0.7755 1.0000
2.500 0.5740 0.02371 0.01620 -0.0580 0.7689 1.0000
2.750 0.5931 0.02422 0.01680 -0.0570 0.7620 1.0000
3.000 0.6132 0.02456 0.01722 -0.0555 0.7561 1.0000
3.250 0.6316 0.02475 0.01751 -0.0537 0.7461 1.0000
3.500 0.6497 0.02449 0.01733 -0.0507 0.7329 1.0000
3.750 0.6685 0.02356 0.01647 -0.0466 0.7175 1.0000
4.000 0.6873 0.02215 0.01508 -0.0418 0.7015 1.0000
4.250 0.7052 0.02105 0.01405 -0.0378 0.6831 1.0000
4.500 0.7258 0.02024 0.01331 -0.0348 0.6710 1.0000
4.750 0.7466 0.01952 0.01273 -0.0320 0.6582 1.0000
5.000 0.7656 0.01864 0.01198 -0.0288 0.6355 1.0000
5.250 0.7850 0.01748 0.01087 -0.0252 0.6061 1.0000
5.500 0.8031 0.01658 0.01000 -0.0217 0.5569 1.0000
5.750 0.8040 0.01735 0.00946 -0.0161 0.3048 1.0000
6.000 0.7933 0.02078 0.01152 -0.0116 0.1354 1.0000
6.250 0.7994 0.02224 0.01284 -0.0084 0.1095 1.0000
6.500 0.8075 0.02340 0.01394 -0.0053 0.0973 1.0000
6.750 0.8164 0.02463 0.01510 -0.0022 0.0902 1.0000
7.000 0.8288 0.02568 0.01612 0.0004 0.0826 1.0000
7.250 0.8450 0.02708 0.01746 0.0025 0.0763 1.0000
7.500 0.8658 0.02849 0.01895 0.0042 0.0721 1.0000
7.750 0.8877 0.03036 0.02074 0.0052 0.0673 1.0000
8.000 0.9075 0.03210 0.02273 0.0069 0.0636 1.0000
8.250 0.9285 0.03432 0.02522 0.0085 0.0620 1.0000
8.500 0.9461 0.03684 0.02807 0.0104 0.0609 1.0000
8.750 0.9599 0.03918 0.03069 0.0126 0.0592 1.0000
9.000 0.9745 0.04136 0.03295 0.0141 0.0567 1.0000
9.250 0.9834 0.04446 0.03646 0.0167 0.0568 1.0000
9.500 0.9810 0.04848 0.04119 0.0205 0.0594 1.0000
9.750 0.9795 0.05272 0.04586 0.0234 0.0619 1.0000
10.000 0.9798 0.05720 0.05058 0.0255 0.0644 1.0000
10.250 0.9574 0.06173 0.05575 0.0298 0.0720 1.0000
10.500 0.9303 0.06733 0.06175 0.0330 0.0845 1.0000
10.750 0.8458 0.05812 0.05272 0.0382 0.0717 1.0000
11.000 0.8150 0.06275 0.05752 0.0391 0.0730 1.0000
11.250 0.7849 0.06799 0.06290 0.0388 0.0739 1.0000
11.500 0.7542 0.07399 0.06901 0.0374 0.0746 1.0000
11.750 0.7235 0.08090 0.07600 0.0349 0.0756 1.0000
12.000 0.6545 0.09620 0.09140 0.0253 0.0854 1.0000
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Polar data table (+)
Polar graphs
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