HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/8 AIRFOIL (hq108-il) Reynolds number: 200,000 Max Cl/Cd: 54.59 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq108-il-200000-n5.txt Download as CSV file: xf-hq108-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/8 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5639 0.08665 0.08321 -0.0116 1.0000 0.0107
-8.750 -0.5656 0.08196 0.07857 -0.0145 1.0000 0.0115
-8.500 -0.5720 0.07610 0.07277 -0.0186 1.0000 0.0109
-8.250 -0.5812 0.06991 0.06662 -0.0247 1.0000 0.0111
-8.000 -0.5908 0.06337 0.06001 -0.0296 1.0000 0.0105
-7.750 -0.5938 0.05773 0.05424 -0.0323 1.0000 0.0105
-7.500 -0.5959 0.05108 0.04736 -0.0341 1.0000 0.0099
-7.250 -0.5932 0.04514 0.04110 -0.0347 1.0000 0.0098
-7.000 -0.5882 0.03852 0.03401 -0.0343 1.0000 0.0094
-6.750 -0.5781 0.03334 0.02831 -0.0333 1.0000 0.0092
-6.500 -0.5639 0.02932 0.02379 -0.0321 1.0000 0.0091
-6.250 -0.5470 0.02610 0.02009 -0.0309 1.0000 0.0091
-6.000 -0.5282 0.02354 0.01713 -0.0296 1.0000 0.0094
-5.750 -0.5081 0.02142 0.01467 -0.0283 1.0000 0.0097
-5.500 -0.4874 0.01967 0.01265 -0.0271 1.0000 0.0101
-5.250 -0.4664 0.01822 0.01101 -0.0258 1.0000 0.0110
-5.000 -0.4457 0.01699 0.00961 -0.0245 1.0000 0.0121
-4.750 -0.4244 0.01634 0.00885 -0.0234 1.0000 0.0144
-4.500 -0.4051 0.01509 0.00749 -0.0219 1.0000 0.0161
-4.250 -0.3717 0.01395 0.00624 -0.0234 0.9941 0.0195
-4.000 -0.3378 0.01327 0.00536 -0.0248 0.9880 0.0252
-3.750 -0.3042 0.01249 0.00463 -0.0263 0.9821 0.0439
-3.500 -0.2704 0.01193 0.00413 -0.0279 0.9758 0.0723
-3.250 -0.2378 0.01122 0.00370 -0.0294 0.9691 0.1359
-3.000 -0.2060 0.01051 0.00334 -0.0308 0.9617 0.2356
-2.750 -0.1773 0.00955 0.00302 -0.0317 0.9534 0.4130
-2.500 -0.1469 0.00900 0.00294 -0.0324 0.9460 0.5469
-2.250 -0.1180 0.00878 0.00289 -0.0325 0.9363 0.6185
-2.000 -0.0881 0.00865 0.00284 -0.0326 0.9275 0.6676
-1.750 -0.0585 0.00854 0.00281 -0.0327 0.9190 0.7131
-1.500 -0.0306 0.00847 0.00274 -0.0323 0.9089 0.7437
-1.250 -0.0023 0.00842 0.00267 -0.0321 0.8990 0.7635
-1.000 0.0250 0.00836 0.00262 -0.0317 0.8888 0.7861
-0.500 0.0776 0.00827 0.00251 -0.0303 0.8664 0.8251
-0.250 0.1040 0.00824 0.00246 -0.0298 0.8552 0.8392
0.000 0.1305 0.00820 0.00242 -0.0292 0.8440 0.8537
0.250 0.1568 0.00817 0.00239 -0.0285 0.8306 0.8693
0.500 0.1839 0.00815 0.00235 -0.0280 0.8152 0.8866
0.750 0.2125 0.00813 0.00231 -0.0278 0.7993 0.9055
1.000 0.2444 0.00812 0.00228 -0.0284 0.7827 0.9253
1.250 0.2787 0.00812 0.00227 -0.0296 0.7654 0.9480
1.500 0.3146 0.00813 0.00230 -0.0312 0.7496 0.9716
1.750 0.3509 0.00816 0.00231 -0.0330 0.7328 0.9969
2.000 0.3771 0.00825 0.00236 -0.0327 0.7155 1.0000
2.250 0.4021 0.00835 0.00243 -0.0321 0.6953 1.0000
2.500 0.4271 0.00847 0.00250 -0.0315 0.6706 1.0000
2.750 0.4519 0.00862 0.00261 -0.0307 0.6386 1.0000
3.000 0.4762 0.00883 0.00270 -0.0299 0.5961 1.0000
3.250 0.4995 0.00915 0.00279 -0.0289 0.5335 1.0000
3.500 0.5220 0.00962 0.00296 -0.0279 0.4598 1.0000
3.750 0.5446 0.01017 0.00322 -0.0271 0.3864 1.0000
4.000 0.5671 0.01080 0.00360 -0.0264 0.3113 1.0000
4.250 0.5903 0.01141 0.00398 -0.0259 0.2562 1.0000
4.500 0.6145 0.01193 0.00439 -0.0254 0.2152 1.0000
4.750 0.6382 0.01253 0.00481 -0.0250 0.1625 1.0000
5.250 0.6843 0.01400 0.00589 -0.0239 0.0738 1.0000
5.500 0.7077 0.01473 0.00653 -0.0234 0.0484 1.0000
5.750 0.7308 0.01553 0.00730 -0.0228 0.0275 1.0000
6.000 0.7532 0.01653 0.00835 -0.0220 0.0179 1.0000
6.250 0.7757 0.01752 0.00946 -0.0211 0.0132 1.0000
6.500 0.7969 0.01868 0.01073 -0.0202 0.0100 1.0000
6.750 0.8184 0.01988 0.01210 -0.0192 0.0089 1.0000
7.000 0.8392 0.02128 0.01368 -0.0182 0.0081 1.0000
7.250 0.8599 0.02282 0.01539 -0.0171 0.0074 1.0000
7.500 0.8801 0.02452 0.01731 -0.0161 0.0070 1.0000
7.750 0.8997 0.02649 0.01953 -0.0150 0.0068 1.0000
8.000 0.9177 0.02875 0.02210 -0.0138 0.0066 1.0000
8.250 0.9335 0.03135 0.02505 -0.0125 0.0065 1.0000
8.500 0.9463 0.03435 0.02854 -0.0109 0.0064 1.0000
8.750 0.9550 0.03768 0.03230 -0.0093 0.0064 1.0000
9.000 0.9589 0.04139 0.03644 -0.0074 0.0064 1.0000
9.250 0.9571 0.04540 0.04087 -0.0054 0.0064 1.0000
9.500 0.9489 0.04958 0.04542 -0.0035 0.0065 1.0000
9.750 0.9330 0.05348 0.04959 -0.0012 0.0065 1.0000
10.000 0.9132 0.05785 0.05419 -0.0005 0.0066 1.0000
10.250 0.8921 0.06320 0.05974 -0.0023 0.0067 1.0000
10.500 0.8690 0.07046 0.06717 -0.0070 0.0068 1.0000
10.750 0.8471 0.08051 0.07729 -0.0152 0.0069 1.0000
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