HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/9 AIRFOIL (hq109-il) Reynolds number: 200,000 Max Cl/Cd: 55.28 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq109-il-200000-n5.txt Download as CSV file: xf-hq109-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/9 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.4730 0.11466 0.11130 -0.0038 1.0000 0.0110
-11.250 -0.4789 0.10827 0.10493 -0.0062 1.0000 0.0105
-10.500 -0.5767 0.10012 0.09654 -0.0080 1.0000 0.0098
-10.250 -0.5803 0.09467 0.09114 -0.0106 1.0000 0.0097
-10.000 -0.5857 0.08883 0.08534 -0.0136 1.0000 0.0096
-9.750 -0.5916 0.08309 0.07965 -0.0169 1.0000 0.0095
-9.500 -0.6016 0.07584 0.07245 -0.0215 1.0000 0.0094
-9.250 -0.6140 0.06659 0.06327 -0.0294 1.0000 0.0092
-9.000 -0.6324 0.05945 0.05606 -0.0349 1.0000 0.0091
-8.750 -0.6497 0.05440 0.05087 -0.0363 1.0000 0.0090
-8.500 -0.6608 0.04898 0.04523 -0.0371 1.0000 0.0089
-8.250 -0.6680 0.04330 0.03919 -0.0369 1.0000 0.0089
-8.000 -0.6686 0.03818 0.03362 -0.0361 1.0000 0.0089
-7.750 -0.6631 0.03371 0.02864 -0.0349 1.0000 0.0090
-7.500 -0.6523 0.03007 0.02450 -0.0336 1.0000 0.0091
-7.250 -0.6375 0.02713 0.02110 -0.0322 1.0000 0.0094
-7.000 -0.6200 0.02483 0.01834 -0.0309 1.0000 0.0097
-6.750 -0.6033 0.02227 0.01543 -0.0294 1.0000 0.0103
-6.500 -0.5840 0.02115 0.01422 -0.0285 1.0000 0.0118
-6.250 -0.5635 0.02032 0.01327 -0.0274 1.0000 0.0133
-6.000 -0.5440 0.01894 0.01170 -0.0260 1.0000 0.0143
-5.750 -0.5249 0.01773 0.01033 -0.0245 1.0000 0.0153
-5.500 -0.5060 0.01679 0.00925 -0.0229 1.0000 0.0163
-5.250 -0.4775 0.01557 0.00798 -0.0236 0.9954 0.0195
-5.000 -0.4440 0.01492 0.00719 -0.0251 0.9889 0.0247
-4.750 -0.4103 0.01399 0.00620 -0.0267 0.9831 0.0325
-4.500 -0.3772 0.01340 0.00558 -0.0281 0.9762 0.0454
-4.250 -0.3433 0.01291 0.00509 -0.0296 0.9700 0.0635
-4.000 -0.3115 0.01236 0.00462 -0.0308 0.9620 0.0889
-3.750 -0.2794 0.01178 0.00419 -0.0321 0.9544 0.1282
-3.500 -0.2477 0.01124 0.00381 -0.0333 0.9464 0.1893
-3.250 -0.2186 0.01057 0.00347 -0.0341 0.9372 0.2752
-3.000 -0.1907 0.00976 0.00323 -0.0346 0.9285 0.4241
-2.750 -0.1630 0.00933 0.00314 -0.0346 0.9193 0.5265
-2.500 -0.1350 0.00914 0.00306 -0.0345 0.9100 0.5865
-2.250 -0.1069 0.00901 0.00302 -0.0342 0.9014 0.6360
-2.000 -0.0799 0.00894 0.00300 -0.0337 0.8917 0.6744
-1.750 -0.0530 0.00889 0.00295 -0.0332 0.8817 0.6990
-1.500 -0.0261 0.00886 0.00289 -0.0327 0.8722 0.7199
-1.250 0.0006 0.00884 0.00284 -0.0322 0.8629 0.7391
-1.000 0.0268 0.00881 0.00281 -0.0316 0.8524 0.7555
-0.750 0.0529 0.00879 0.00278 -0.0309 0.8418 0.7721
-0.500 0.0789 0.00877 0.00274 -0.0302 0.8317 0.7882
-0.250 0.1052 0.00875 0.00271 -0.0297 0.8221 0.8022
0.000 0.1316 0.00874 0.00269 -0.0292 0.8117 0.8140
0.250 0.1577 0.00872 0.00266 -0.0286 0.7993 0.8259
0.500 0.1836 0.00870 0.00263 -0.0279 0.7854 0.8384
0.750 0.2096 0.00868 0.00260 -0.0272 0.7711 0.8515
1.000 0.2357 0.00866 0.00258 -0.0266 0.7564 0.8655
1.250 0.2626 0.00865 0.00258 -0.0261 0.7425 0.8808
1.500 0.2908 0.00864 0.00260 -0.0259 0.7283 0.8974
1.750 0.3210 0.00865 0.00263 -0.0262 0.7137 0.9153
2.000 0.3540 0.00868 0.00266 -0.0271 0.6978 0.9346
2.250 0.3887 0.00871 0.00270 -0.0284 0.6793 0.9563
2.500 0.4246 0.00877 0.00277 -0.0301 0.6575 0.9780
2.750 0.4589 0.00887 0.00282 -0.0315 0.6300 1.0000
3.000 0.4826 0.00903 0.00289 -0.0307 0.5978 1.0000
3.250 0.5062 0.00924 0.00297 -0.0298 0.5582 1.0000
3.500 0.5290 0.00957 0.00311 -0.0289 0.5025 1.0000
3.750 0.5510 0.01003 0.00329 -0.0278 0.4360 1.0000
4.000 0.5733 0.01054 0.00355 -0.0270 0.3732 1.0000
4.250 0.5956 0.01112 0.00386 -0.0263 0.3123 1.0000
4.750 0.6425 0.01214 0.00464 -0.0251 0.2340 1.0000
5.000 0.6664 0.01263 0.00503 -0.0247 0.2002 1.0000
5.250 0.6898 0.01318 0.00544 -0.0242 0.1597 1.0000
5.500 0.7126 0.01384 0.00592 -0.0236 0.1195 1.0000
5.750 0.7356 0.01450 0.00647 -0.0230 0.0898 1.0000
6.000 0.7588 0.01514 0.00706 -0.0224 0.0706 1.0000
6.250 0.7815 0.01585 0.00772 -0.0218 0.0541 1.0000
6.500 0.8043 0.01653 0.00846 -0.0212 0.0397 1.0000
6.750 0.8269 0.01725 0.00915 -0.0206 0.0271 1.0000
7.000 0.8483 0.01818 0.01010 -0.0197 0.0199 1.0000
7.250 0.8694 0.01915 0.01119 -0.0188 0.0153 1.0000
7.500 0.8870 0.02067 0.01283 -0.0173 0.0127 1.0000
7.750 0.9077 0.02166 0.01399 -0.0164 0.0111 1.0000
8.000 0.9286 0.02256 0.01499 -0.0156 0.0093 1.0000
8.250 0.9473 0.02378 0.01632 -0.0146 0.0084 1.0000
8.500 0.9622 0.02570 0.01847 -0.0131 0.0078 1.0000
8.750 0.9779 0.02761 0.02060 -0.0116 0.0075 1.0000
9.000 0.9936 0.02953 0.02279 -0.0103 0.0072 1.0000
9.250 1.0073 0.03173 0.02527 -0.0087 0.0070 1.0000
9.500 1.0183 0.03419 0.02807 -0.0071 0.0068 1.0000
9.750 1.0255 0.03692 0.03114 -0.0052 0.0067 1.0000
10.000 1.0275 0.03995 0.03454 -0.0030 0.0065 1.0000
10.250 1.0237 0.04294 0.03786 -0.0005 0.0065 1.0000
10.500 1.0131 0.04605 0.04127 0.0023 0.0065 1.0000
10.750 1.0000 0.04946 0.04494 0.0040 0.0065 1.0000
11.000 0.9852 0.05334 0.04906 0.0044 0.0065 1.0000
11.250 0.9677 0.05806 0.05401 0.0035 0.0065 1.0000
11.500 0.9495 0.06359 0.05974 0.0009 0.0065 1.0000
11.750 0.9319 0.06997 0.06630 -0.0031 0.0066 1.0000
12.000 0.9119 0.07820 0.07470 -0.0092 0.0067 1.0000
12.250 0.8899 0.08875 0.08539 -0.0170 0.0068 1.0000
12.500 0.8667 0.10095 0.09767 -0.0248 0.0069 1.0000
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Polar data table (+)
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