HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il) Reynolds number: 200,000 Max Cl/Cd: 55.84 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1010-il-200000-n5.txt Download as CSV file: xf-hq1010-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/10 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6979 0.05698 0.05340 -0.0397 1.0000 0.0091
-10.000 -0.7233 0.05114 0.04738 -0.0428 1.0000 0.0091
-9.750 -0.7514 0.04615 0.04211 -0.0420 1.0000 0.0088
-9.500 -0.7610 0.04231 0.03798 -0.0410 1.0000 0.0090
-9.250 -0.7668 0.03809 0.03333 -0.0395 1.0000 0.0091
-9.000 -0.7636 0.03472 0.02957 -0.0381 1.0000 0.0094
-8.750 -0.7546 0.03196 0.02644 -0.0368 1.0000 0.0099
-8.500 -0.7410 0.02987 0.02400 -0.0355 1.0000 0.0108
-8.250 -0.7254 0.02793 0.02170 -0.0343 1.0000 0.0114
-8.000 -0.7115 0.02544 0.01889 -0.0328 1.0000 0.0119
-7.750 -0.6952 0.02360 0.01684 -0.0314 1.0000 0.0126
-7.500 -0.6771 0.02229 0.01540 -0.0302 1.0000 0.0133
-7.250 -0.6586 0.02113 0.01406 -0.0288 1.0000 0.0142
-7.000 -0.6403 0.02000 0.01279 -0.0274 1.0000 0.0152
-6.750 -0.6218 0.01909 0.01175 -0.0259 1.0000 0.0167
-6.500 -0.6048 0.01811 0.01070 -0.0243 1.0000 0.0187
-6.250 -0.5823 0.01740 0.00995 -0.0238 0.9981 0.0213
-6.000 -0.5494 0.01651 0.00891 -0.0252 0.9912 0.0250
-5.750 -0.5162 0.01563 0.00801 -0.0269 0.9849 0.0313
-5.500 -0.4836 0.01500 0.00728 -0.0282 0.9778 0.0394
-5.250 -0.4489 0.01453 0.00674 -0.0299 0.9719 0.0486
-5.000 -0.4179 0.01397 0.00617 -0.0310 0.9633 0.0592
-4.750 -0.3843 0.01345 0.00565 -0.0325 0.9566 0.0729
-4.500 -0.3537 0.01295 0.00517 -0.0333 0.9473 0.0898
-4.250 -0.3226 0.01245 0.00470 -0.0343 0.9389 0.1150
-4.000 -0.2921 0.01189 0.00432 -0.0352 0.9305 0.1599
-3.500 -0.2364 0.01071 0.00367 -0.0360 0.9122 0.3060
-3.250 -0.2109 0.01006 0.00343 -0.0359 0.9032 0.4217
-3.000 -0.1856 0.00970 0.00334 -0.0353 0.8936 0.5045
-2.750 -0.1590 0.00952 0.00328 -0.0349 0.8851 0.5603
-2.500 -0.1328 0.00941 0.00325 -0.0343 0.8760 0.6053
-2.250 -0.1068 0.00935 0.00323 -0.0336 0.8670 0.6412
-2.000 -0.0802 0.00932 0.00320 -0.0331 0.8590 0.6686
-1.750 -0.0538 0.00931 0.00316 -0.0325 0.8497 0.6882
-1.500 -0.0270 0.00929 0.00311 -0.0321 0.8411 0.7028
-1.250 -0.0002 0.00929 0.00305 -0.0316 0.8326 0.7173
-1.000 0.0262 0.00928 0.00304 -0.0311 0.8232 0.7331
-0.750 0.0526 0.00927 0.00303 -0.0306 0.8147 0.7483
-0.500 0.0789 0.00927 0.00302 -0.0300 0.8059 0.7627
-0.250 0.1054 0.00926 0.00300 -0.0296 0.7963 0.7753
0.000 0.1318 0.00924 0.00297 -0.0290 0.7855 0.7848
0.250 0.1582 0.00923 0.00293 -0.0285 0.7733 0.7944
0.500 0.1846 0.00922 0.00289 -0.0280 0.7605 0.8046
0.750 0.2109 0.00920 0.00287 -0.0274 0.7477 0.8147
1.000 0.2371 0.00919 0.00285 -0.0268 0.7338 0.8249
1.250 0.2634 0.00918 0.00285 -0.0263 0.7203 0.8358
1.500 0.2898 0.00918 0.00287 -0.0258 0.7080 0.8475
1.750 0.3163 0.00918 0.00290 -0.0254 0.6954 0.8598
2.000 0.3430 0.00919 0.00294 -0.0249 0.6819 0.8730
2.500 0.3991 0.00923 0.00304 -0.0245 0.6508 0.9031
2.750 0.4300 0.00927 0.00311 -0.0250 0.6308 0.9200
3.000 0.4630 0.00935 0.00317 -0.0260 0.6047 0.9397
3.250 0.4973 0.00948 0.00326 -0.0273 0.5699 0.9609
3.500 0.5317 0.00968 0.00333 -0.0288 0.5258 0.9846
3.750 0.5601 0.01003 0.00345 -0.0291 0.4695 1.0000
4.000 0.5807 0.01050 0.00364 -0.0279 0.4092 1.0000
4.250 0.6020 0.01100 0.00392 -0.0270 0.3538 1.0000
4.500 0.6241 0.01150 0.00423 -0.0262 0.3088 1.0000
4.750 0.6467 0.01199 0.00457 -0.0255 0.2739 1.0000
5.000 0.6699 0.01245 0.00493 -0.0249 0.2455 1.0000
5.250 0.6934 0.01289 0.00530 -0.0243 0.2181 1.0000
5.500 0.7171 0.01333 0.00571 -0.0237 0.1913 1.0000
5.750 0.7402 0.01383 0.00612 -0.0232 0.1610 1.0000
6.000 0.7627 0.01442 0.00658 -0.0225 0.1298 1.0000
6.250 0.7848 0.01506 0.00711 -0.0218 0.1047 1.0000
6.500 0.8073 0.01568 0.00767 -0.0212 0.0850 1.0000
6.750 0.8295 0.01631 0.00827 -0.0205 0.0701 1.0000
7.000 0.8517 0.01695 0.00895 -0.0198 0.0584 1.0000
7.250 0.8736 0.01762 0.00967 -0.0190 0.0497 1.0000
7.500 0.8949 0.01836 0.01041 -0.0182 0.0401 1.0000
7.750 0.9161 0.01908 0.01109 -0.0174 0.0301 1.0000
8.000 0.9370 0.01983 0.01188 -0.0166 0.0235 1.0000
8.250 0.9566 0.02074 0.01288 -0.0156 0.0189 1.0000
8.500 0.9747 0.02181 0.01401 -0.0144 0.0155 1.0000
8.750 0.9907 0.02310 0.01549 -0.0129 0.0135 1.0000
9.000 1.0068 0.02433 0.01688 -0.0114 0.0121 1.0000
9.250 1.0227 0.02553 0.01820 -0.0100 0.0110 1.0000
9.500 1.0368 0.02682 0.01959 -0.0086 0.0099 1.0000
9.750 1.0474 0.02849 0.02139 -0.0068 0.0092 1.0000
10.000 1.0607 0.02982 0.02292 -0.0052 0.0086 1.0000
10.250 1.0711 0.03130 0.02460 -0.0034 0.0081 1.0000
10.500 1.0773 0.03300 0.02649 -0.0011 0.0077 1.0000
10.750 1.0814 0.03486 0.02854 0.0012 0.0075 1.0000
11.000 1.0842 0.03680 0.03068 0.0033 0.0072 1.0000
11.250 1.0849 0.03894 0.03302 0.0050 0.0070 1.0000
11.500 1.0838 0.04127 0.03555 0.0065 0.0068 1.0000
11.750 1.0800 0.04404 0.03857 0.0076 0.0067 1.0000
12.000 1.0752 0.04695 0.04167 0.0081 0.0066 1.0000
12.250 1.0670 0.05058 0.04553 0.0082 0.0066 1.0000
12.500 1.0572 0.05459 0.04976 0.0075 0.0066 1.0000
12.750 1.0462 0.05904 0.05438 0.0060 0.0064 1.0000
13.000 1.0298 0.06487 0.06047 0.0034 0.0065 1.0000
13.250 1.0130 0.07135 0.06717 -0.0002 0.0066 1.0000
13.500 0.9946 0.07881 0.07481 -0.0049 0.0065 1.0000
13.750 0.9667 0.08940 0.08567 -0.0118 0.0068 1.0000
14.000 0.9376 0.10151 0.09794 -0.0196 0.0069 1.0000
14.250 0.8952 0.11820 0.11480 -0.0294 0.0073 1.0000
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Polar data table (+)
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