NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 200,000 Max Cl/Cd: 71.51 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n6h10-il-200000.txt Download as CSV file: xf-n6h10-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 6-H-10 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4335 0.09422 0.09052 0.0020 0.7044 0.0243
-7.500 -0.4339 0.09132 0.08764 0.0001 0.7015 0.0246
-7.250 -0.4332 0.08872 0.08503 -0.0018 0.6989 0.0250
-7.000 -0.4268 0.08621 0.08245 -0.0044 0.6966 0.0255
-6.750 -0.4168 0.08359 0.07972 -0.0064 0.6946 0.0258
-6.500 -0.4054 0.08075 0.07676 -0.0076 0.6928 0.0259
-6.250 -0.3926 0.07775 0.07360 -0.0083 0.6912 0.0260
-6.000 -0.3848 0.07226 0.06810 -0.0093 0.6887 0.0264
-5.750 -0.3784 0.06742 0.06330 -0.0091 0.6863 0.0273
-5.500 -0.3665 0.06421 0.06006 -0.0090 0.6840 0.0281
-5.250 -0.3527 0.06124 0.05700 -0.0089 0.6818 0.0293
-5.000 -0.3371 0.05835 0.05399 -0.0088 0.6798 0.0307
-4.750 -0.3198 0.05552 0.05099 -0.0086 0.6782 0.0325
-4.500 -0.2916 0.05378 0.04894 -0.0086 0.6762 0.0361
-4.250 -0.2641 0.05305 0.04783 -0.0080 0.6735 0.0370
-4.000 -0.2574 0.04660 0.04145 -0.0080 0.6714 0.0390
-3.750 -0.2398 0.04439 0.03921 -0.0077 0.6693 0.0422
-3.500 -0.2080 0.04514 0.03939 -0.0061 0.6674 0.0500
-3.250 -0.1949 0.03989 0.03407 -0.0054 0.6660 0.0520
-3.000 -0.1758 0.03765 0.03179 -0.0049 0.6645 0.0550
-2.750 -0.1463 0.03916 0.03266 -0.0028 0.6632 0.0642
-2.500 -0.1288 0.03408 0.02777 -0.0037 0.6606 0.0681
-2.250 -0.1034 0.03302 0.02641 -0.0031 0.6578 0.0799
-2.000 -0.0798 0.03130 0.02470 -0.0031 0.6556 0.0871
-1.750 -0.0558 0.02995 0.02320 -0.0026 0.6539 0.1009
-1.500 -0.0323 0.02883 0.02185 -0.0019 0.6524 0.1238
-1.250 -0.0084 0.02743 0.02040 -0.0014 0.6511 0.1443
-1.000 0.0385 0.02442 0.01624 0.0011 0.6501 0.0507
-0.750 0.0707 0.02240 0.01408 0.0008 0.6491 0.0452
-0.500 0.1039 0.02156 0.01316 -0.0004 0.6467 0.0430
-0.250 0.1341 0.02107 0.01269 -0.0013 0.6437 0.0434
0.000 0.1606 0.02070 0.01233 -0.0012 0.6410 0.0474
0.250 0.1846 0.02029 0.01189 -0.0004 0.6390 0.0514
0.500 0.2068 0.01987 0.01144 0.0009 0.6373 0.0589
0.750 0.2265 0.01925 0.01115 0.0028 0.6359 0.1530
1.000 0.4728 0.01860 0.01216 -0.0373 0.6354 1.0000
1.250 0.4994 0.01896 0.01251 -0.0378 0.6332 1.0000
1.500 0.5285 0.01956 0.01316 -0.0393 0.6293 1.0000
1.750 0.5545 0.01985 0.01344 -0.0396 0.6265 1.0000
2.000 0.5794 0.02005 0.01364 -0.0394 0.6245 1.0000
2.250 0.6036 0.02017 0.01375 -0.0389 0.6229 1.0000
2.500 0.6269 0.02019 0.01373 -0.0380 0.6214 1.0000
2.750 0.6524 0.02060 0.01418 -0.0382 0.6186 1.0000
3.000 0.6803 0.02139 0.01510 -0.0398 0.6137 1.0000
3.250 0.7050 0.02173 0.01550 -0.0398 0.6112 1.0000
3.500 0.7285 0.02181 0.01561 -0.0391 0.6092 1.0000
3.750 0.7491 0.02106 0.01478 -0.0365 0.6070 1.0000
4.000 0.7735 0.02106 0.01488 -0.0363 0.5978 1.0000
4.250 0.7931 0.01956 0.01328 -0.0326 0.5935 1.0000
4.500 0.8170 0.01930 0.01310 -0.0319 0.5854 1.0000
4.750 0.8390 0.01837 0.01213 -0.0296 0.5804 1.0000
5.000 0.8624 0.01795 0.01176 -0.0285 0.5736 1.0000
5.250 0.8857 0.01733 0.01116 -0.0271 0.5671 1.0000
5.500 0.9091 0.01678 0.01070 -0.0258 0.5589 1.0000
5.750 0.9319 0.01579 0.00966 -0.0238 0.5502 1.0000
6.000 0.9555 0.01506 0.00899 -0.0225 0.5353 1.0000
6.250 0.9799 0.01447 0.00855 -0.0216 0.5118 1.0000
6.500 1.0025 0.01402 0.00768 -0.0204 0.3828 1.0000
6.750 1.0081 0.01904 0.01097 -0.0229 0.1170 1.0000
7.000 1.0213 0.02059 0.01244 -0.0224 0.0994 1.0000
7.250 1.0323 0.02197 0.01389 -0.0213 0.0930 1.0000
7.500 1.0402 0.02332 0.01533 -0.0199 0.0887 1.0000
7.750 1.0415 0.02495 0.01701 -0.0180 0.0858 1.0000
8.000 1.0309 0.02664 0.01873 -0.0142 0.0835 1.0000
8.250 1.0139 0.02779 0.01989 -0.0084 0.0823 1.0000
8.500 1.0066 0.02868 0.02085 -0.0038 0.0809 1.0000
8.750 1.0006 0.02956 0.02178 0.0005 0.0796 1.0000
9.000 0.9963 0.03039 0.02265 0.0048 0.0780 1.0000
9.250 0.9954 0.03117 0.02346 0.0086 0.0765 1.0000
9.500 0.9980 0.03194 0.02424 0.0120 0.0742 1.0000
9.750 1.0122 0.03267 0.02483 0.0150 0.0691 1.0000
10.000 1.0161 0.03353 0.02582 0.0172 0.0667 1.0000
10.250 1.0236 0.03445 0.02683 0.0192 0.0629 1.0000
10.500 1.0337 0.03541 0.02780 0.0210 0.0586 1.0000
10.750 1.0492 0.03645 0.02895 0.0228 0.0538 1.0000
11.000 1.0527 0.03770 0.03030 0.0243 0.0491 1.0000
11.250 1.0673 0.03912 0.03173 0.0260 0.0431 1.0000
11.500 1.0708 0.04053 0.03332 0.0274 0.0388 1.0000
11.750 1.0789 0.04192 0.03472 0.0286 0.0355 1.0000
12.000 1.0893 0.04379 0.03675 0.0302 0.0316 1.0000
12.250 1.0963 0.04548 0.03867 0.0315 0.0287 1.0000
12.500 1.1029 0.04722 0.04051 0.0326 0.0269 1.0000
12.750 1.1122 0.04923 0.04258 0.0337 0.0253 1.0000
13.000 1.1162 0.05275 0.04641 0.0352 0.0240 1.0000
13.250 1.1131 0.05520 0.04912 0.0361 0.0229 1.0000
13.500 1.1097 0.05854 0.05275 0.0370 0.0223 1.0000
13.750 1.1024 0.06213 0.05661 0.0377 0.0218 1.0000
14.000 1.0931 0.06571 0.06041 0.0379 0.0209 1.0000
14.250 1.0775 0.07069 0.06567 0.0380 0.0214 1.0000
14.500 1.0619 0.07531 0.07051 0.0376 0.0210 1.0000
14.750 1.0418 0.08087 0.07630 0.0366 0.0211 1.0000
15.000 1.0211 0.08680 0.08244 0.0350 0.0210 1.0000
15.250 0.9969 0.09377 0.08962 0.0325 0.0215 1.0000
15.500 0.9736 0.10105 0.09707 0.0294 0.0217 1.0000
15.750 0.9494 0.10916 0.10533 0.0255 0.0222 1.0000
16.000 0.9255 0.11792 0.11423 0.0208 0.0226 1.0000
16.250 0.9012 0.12780 0.12420 0.0153 0.0232 1.0000
16.500 0.8777 0.13862 0.13508 0.0092 0.0237 1.0000
16.750 0.8629 0.14725 0.14368 0.0054 0.0245 1.0000
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Polar data table (+)
Polar graphs
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