HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 0/9 AIRFOIL (hq09-il) Reynolds number: 200,000 Max Cl/Cd: 38.62 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq09-il-200000-n5.txt Download as CSV file: xf-hq09-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 0/9 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.7192 0.08439 0.08111 0.0026 1.0000 0.0057
-10.000 -0.7327 0.07470 0.07143 -0.0036 1.0000 0.0056
-9.750 -0.7502 0.06419 0.06085 -0.0121 1.0000 0.0055
-9.500 -0.7707 0.05790 0.05442 -0.0164 1.0000 0.0054
-9.250 -0.7898 0.05351 0.04988 -0.0168 1.0000 0.0054
-9.000 -0.8048 0.04874 0.04485 -0.0162 1.0000 0.0054
-8.750 -0.8137 0.04363 0.03936 -0.0152 1.0000 0.0053
-8.500 -0.8155 0.03880 0.03408 -0.0140 1.0000 0.0053
-8.250 -0.8108 0.03446 0.02923 -0.0127 1.0000 0.0053
-8.000 -0.8009 0.03055 0.02478 -0.0112 1.0000 0.0054
-7.750 -0.7861 0.02739 0.02114 -0.0100 1.0000 0.0055
-7.500 -0.7693 0.02452 0.01786 -0.0087 1.0000 0.0057
-7.250 -0.7505 0.02232 0.01536 -0.0077 1.0000 0.0063
-7.000 -0.7285 0.02107 0.01394 -0.0070 1.0000 0.0071
-6.750 -0.7064 0.01974 0.01243 -0.0062 1.0000 0.0085
-6.500 -0.6836 0.01858 0.01097 -0.0053 1.0000 0.0110
-6.250 -0.6600 0.01791 0.01029 -0.0050 1.0000 0.0163
-6.000 -0.6349 0.01762 0.00992 -0.0048 1.0000 0.0224
-5.750 -0.6108 0.01705 0.00929 -0.0044 1.0000 0.0266
-5.500 -0.5872 0.01624 0.00836 -0.0037 1.0000 0.0292
-5.250 -0.5640 0.01541 0.00741 -0.0029 1.0000 0.0306
-5.000 -0.5412 0.01454 0.00645 -0.0021 1.0000 0.0333
-4.750 -0.5177 0.01387 0.00573 -0.0015 1.0000 0.0380
-4.500 -0.4936 0.01335 0.00507 -0.0008 1.0000 0.0447
-4.250 -0.4700 0.01279 0.00455 -0.0002 1.0000 0.0587
-4.000 -0.4472 0.01218 0.00410 0.0005 1.0000 0.0902
-3.750 -0.4259 0.01140 0.00368 0.0012 1.0000 0.1570
-3.500 -0.4043 0.01082 0.00336 0.0020 1.0000 0.2297
-3.250 -0.3847 0.01006 0.00306 0.0030 1.0000 0.3248
-3.000 -0.3588 0.00930 0.00285 0.0028 0.9928 0.4565
-2.750 -0.3267 0.00883 0.00274 0.0016 0.9817 0.5593
-2.500 -0.2939 0.00862 0.00273 0.0006 0.9705 0.6297
-2.250 -0.2613 0.00849 0.00273 -0.0002 0.9593 0.6821
-2.000 -0.2283 0.00843 0.00269 -0.0011 0.9479 0.7123
-1.750 -0.1951 0.00838 0.00261 -0.0021 0.9353 0.7279
-1.500 -0.1633 0.00834 0.00251 -0.0027 0.9200 0.7431
-1.250 -0.1334 0.00830 0.00246 -0.0028 0.9037 0.7587
-1.000 -0.1051 0.00828 0.00242 -0.0026 0.8887 0.7743
-0.750 -0.0781 0.00826 0.00241 -0.0021 0.8757 0.7912
-0.500 -0.0519 0.00826 0.00239 -0.0014 0.8624 0.8072
-0.250 -0.0258 0.00825 0.00238 -0.0007 0.8488 0.8212
0.000 0.0000 0.00825 0.00237 0.0000 0.8349 0.8349
0.250 0.0259 0.00825 0.00238 0.0007 0.8212 0.8488
0.500 0.0519 0.00826 0.00239 0.0014 0.8072 0.8623
0.750 0.0781 0.00826 0.00241 0.0021 0.7912 0.8756
1.000 0.1051 0.00828 0.00242 0.0026 0.7743 0.8888
1.250 0.1334 0.00830 0.00246 0.0028 0.7587 0.9037
1.500 0.1633 0.00834 0.00251 0.0027 0.7431 0.9200
1.750 0.1951 0.00838 0.00261 0.0021 0.7279 0.9353
2.000 0.2282 0.00843 0.00269 0.0011 0.7123 0.9479
2.250 0.2613 0.00849 0.00273 0.0002 0.6819 0.9594
2.500 0.2938 0.00862 0.00272 -0.0006 0.6294 0.9706
2.750 0.3266 0.00883 0.00274 -0.0016 0.5597 0.9818
3.000 0.3588 0.00929 0.00285 -0.0028 0.4572 0.9929
3.250 0.3845 0.01006 0.00306 -0.0030 0.3246 1.0000
3.500 0.4042 0.01081 0.00336 -0.0020 0.2304 1.0000
3.750 0.4258 0.01140 0.00368 -0.0012 0.1577 1.0000
4.000 0.4471 0.01217 0.00410 -0.0004 0.0904 1.0000
4.250 0.4699 0.01279 0.00455 0.0002 0.0584 1.0000
4.500 0.4935 0.01335 0.00507 0.0008 0.0446 1.0000
4.750 0.5176 0.01387 0.00573 0.0015 0.0380 1.0000
5.000 0.5411 0.01454 0.00645 0.0021 0.0332 1.0000
5.250 0.5639 0.01540 0.00740 0.0029 0.0307 1.0000
5.500 0.5872 0.01624 0.00836 0.0037 0.0291 1.0000
5.750 0.6108 0.01705 0.00929 0.0044 0.0266 1.0000
6.000 0.6350 0.01761 0.00991 0.0048 0.0222 1.0000
6.250 0.6600 0.01794 0.01033 0.0050 0.0164 1.0000
6.500 0.6838 0.01856 0.01095 0.0053 0.0110 1.0000
6.750 0.7066 0.01975 0.01244 0.0062 0.0084 1.0000
7.000 0.7287 0.02107 0.01395 0.0070 0.0071 1.0000
7.250 0.7507 0.02233 0.01537 0.0077 0.0063 1.0000
7.500 0.7695 0.02453 0.01787 0.0087 0.0057 1.0000
7.750 0.7864 0.02739 0.02114 0.0099 0.0055 1.0000
8.000 0.8012 0.03057 0.02480 0.0111 0.0054 1.0000
8.250 0.8114 0.03441 0.02917 0.0125 0.0054 1.0000
8.500 0.8154 0.03896 0.03426 0.0140 0.0053 1.0000
8.750 0.8139 0.04373 0.03948 0.0151 0.0053 1.0000
9.000 0.8046 0.04893 0.04506 0.0160 0.0053 1.0000
9.250 0.7908 0.05349 0.04986 0.0166 0.0054 1.0000
9.500 0.7703 0.05812 0.05465 0.0161 0.0054 1.0000
9.750 0.7514 0.06411 0.06077 0.0120 0.0055 1.0000
10.000 0.7328 0.07508 0.07181 0.0032 0.0056 1.0000
10.250 0.7202 0.08446 0.08118 -0.0028 0.0057 1.0000
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Polar data table (+)
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