HQ 3.5/12 AIRFOIL (hq3512-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.5/12 AIRFOIL (hq3512-il) Reynolds number: 500,000 Max Cl/Cd: 112.11 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq3512-il-500000.txt Download as CSV file: xf-hq3512-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.5/12 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3601 0.03618 0.03318 -0.1089 0.9614 0.0147
-8.250 -0.3514 0.03036 0.02697 -0.1118 0.9531 0.0141
-8.000 -0.3443 0.02494 0.02090 -0.1121 0.9429 0.0139
-7.750 -0.3294 0.02185 0.01734 -0.1114 0.9341 0.0139
-7.500 -0.3083 0.01967 0.01478 -0.1110 0.9279 0.0141
-7.250 -0.2870 0.01813 0.01296 -0.1103 0.9206 0.0143
-7.000 -0.2624 0.01718 0.01180 -0.1098 0.9146 0.0146
-6.750 -0.2412 0.01535 0.00976 -0.1091 0.9081 0.0154
-6.500 -0.2166 0.01461 0.00894 -0.1087 0.9018 0.0161
-6.250 -0.1912 0.01389 0.00810 -0.1083 0.8964 0.0167
-6.000 -0.1667 0.01323 0.00735 -0.1078 0.8895 0.0176
-5.750 -0.1407 0.01266 0.00666 -0.1074 0.8840 0.0187
-5.500 -0.1162 0.01192 0.00585 -0.1069 0.8776 0.0202
-5.250 -0.0899 0.01153 0.00544 -0.1066 0.8717 0.0228
-5.000 -0.0635 0.01107 0.00491 -0.1064 0.8663 0.0277
-4.750 -0.0369 0.01080 0.00467 -0.1062 0.8600 0.0373
-4.500 -0.0091 0.01068 0.00449 -0.1061 0.8548 0.0439
-4.250 0.0176 0.01042 0.00423 -0.1060 0.8488 0.0498
-4.000 0.0451 0.01029 0.00403 -0.1059 0.8429 0.0537
-3.750 0.0723 0.00999 0.00371 -0.1058 0.8378 0.0612
-3.500 0.0996 0.00987 0.00356 -0.1057 0.8314 0.0672
-3.250 0.1269 0.00954 0.00322 -0.1056 0.8259 0.0783
-3.000 0.1541 0.00927 0.00298 -0.1055 0.8199 0.0958
-2.750 0.1804 0.00875 0.00272 -0.1055 0.8136 0.1676
-2.500 0.2052 0.00784 0.00247 -0.1056 0.8076 0.3754
-2.250 0.2307 0.00740 0.00242 -0.1053 0.7999 0.5081
-2.000 0.2578 0.00729 0.00238 -0.1050 0.7928 0.5622
-1.750 0.2848 0.00724 0.00237 -0.1047 0.7846 0.5974
-1.500 0.3124 0.00724 0.00234 -0.1045 0.7775 0.6207
-1.250 0.3399 0.00724 0.00233 -0.1042 0.7701 0.6400
-1.000 0.3677 0.00728 0.00233 -0.1041 0.7635 0.6591
-0.750 0.3950 0.00730 0.00234 -0.1038 0.7558 0.6753
-0.500 0.4221 0.00734 0.00239 -0.1034 0.7493 0.6959
-0.250 0.4490 0.00737 0.00245 -0.1031 0.7418 0.7126
0.000 0.4766 0.00741 0.00243 -0.1029 0.7345 0.7230
0.250 0.5039 0.00742 0.00243 -0.1027 0.7259 0.7311
0.500 0.5315 0.00744 0.00242 -0.1025 0.7179 0.7379
0.750 0.5590 0.00746 0.00243 -0.1024 0.7102 0.7449
1.000 0.5866 0.00749 0.00244 -0.1023 0.7025 0.7520
1.250 0.6138 0.00752 0.00245 -0.1021 0.6942 0.7589
1.500 0.6411 0.00754 0.00248 -0.1020 0.6850 0.7665
1.750 0.6681 0.00758 0.00250 -0.1017 0.6767 0.7737
2.000 0.6951 0.00762 0.00254 -0.1015 0.6671 0.7816
2.250 0.7218 0.00765 0.00259 -0.1012 0.6569 0.7898
2.500 0.7482 0.00770 0.00263 -0.1009 0.6458 0.7983
2.750 0.7744 0.00777 0.00268 -0.1006 0.6341 0.8077
3.000 0.8001 0.00780 0.00274 -0.1001 0.6211 0.8171
3.250 0.8255 0.00787 0.00282 -0.0996 0.6060 0.8279
3.500 0.8504 0.00794 0.00289 -0.0990 0.5888 0.8402
3.750 0.8744 0.00801 0.00297 -0.0982 0.5684 0.8550
4.000 0.8965 0.00811 0.00307 -0.0970 0.5447 0.8755
4.250 0.9182 0.00819 0.00319 -0.0956 0.5160 0.9238
4.500 0.9483 0.00852 0.00337 -0.0965 0.4775 1.0000
4.750 0.9702 0.00895 0.00361 -0.0955 0.4411 1.0000
5.000 0.9913 0.00942 0.00390 -0.0945 0.4047 1.0000
5.250 1.0121 0.00991 0.00421 -0.0934 0.3700 1.0000
5.500 1.0331 0.01037 0.00452 -0.0924 0.3392 1.0000
5.750 1.0544 0.01082 0.00484 -0.0914 0.3152 1.0000
6.000 1.0758 0.01125 0.00517 -0.0905 0.2976 1.0000
6.250 1.0973 0.01165 0.00551 -0.0895 0.2835 1.0000
6.500 1.1196 0.01200 0.00583 -0.0887 0.2727 1.0000
6.750 1.1419 0.01233 0.00616 -0.0879 0.2641 1.0000
7.250 1.1856 0.01302 0.00684 -0.0861 0.2476 1.0000
7.500 1.2056 0.01343 0.00723 -0.0849 0.2402 1.0000
7.750 1.2280 0.01369 0.00753 -0.0841 0.2310 1.0000
8.000 1.2484 0.01403 0.00786 -0.0829 0.2205 1.0000
8.250 1.2668 0.01440 0.00821 -0.0815 0.2090 1.0000
8.500 1.2845 0.01478 0.00857 -0.0799 0.1966 1.0000
8.750 1.3025 0.01517 0.00894 -0.0784 0.1842 1.0000
9.000 1.3191 0.01563 0.00936 -0.0767 0.1698 1.0000
9.250 1.3347 0.01616 0.00984 -0.0749 0.1541 1.0000
9.500 1.3490 0.01677 0.01041 -0.0730 0.1380 1.0000
9.750 1.3621 0.01746 0.01104 -0.0710 0.1226 1.0000
10.000 1.3734 0.01829 0.01178 -0.0688 0.1061 1.0000
10.250 1.3834 0.01921 0.01261 -0.0666 0.0901 1.0000
10.500 1.3941 0.02012 0.01347 -0.0645 0.0762 1.0000
10.750 1.4041 0.02111 0.01441 -0.0625 0.0635 1.0000
11.000 1.4126 0.02223 0.01547 -0.0603 0.0504 1.0000
11.250 1.4140 0.02388 0.01699 -0.0576 0.0300 1.0000
11.500 1.4108 0.02597 0.01895 -0.0545 0.0146 1.0000
11.750 1.4138 0.02767 0.02068 -0.0522 0.0110 1.0000
12.000 1.4221 0.02899 0.02210 -0.0506 0.0100 1.0000
12.250 1.4278 0.03056 0.02376 -0.0488 0.0090 1.0000
12.500 1.4312 0.03239 0.02568 -0.0471 0.0084 1.0000
12.750 1.4318 0.03452 0.02792 -0.0454 0.0080 1.0000
13.000 1.4358 0.03641 0.02994 -0.0441 0.0077 1.0000
13.250 1.4384 0.03850 0.03214 -0.0429 0.0075 1.0000
13.500 1.4399 0.04077 0.03452 -0.0418 0.0072 1.0000
13.750 1.4398 0.04330 0.03718 -0.0409 0.0070 1.0000
14.000 1.4401 0.04588 0.03987 -0.0402 0.0068 1.0000
14.250 1.4397 0.04864 0.04273 -0.0397 0.0066 1.0000
14.500 1.4363 0.05189 0.04610 -0.0395 0.0064 1.0000
14.750 1.4311 0.05552 0.04986 -0.0395 0.0063 1.0000
15.000 1.4251 0.05940 0.05385 -0.0398 0.0062 1.0000
15.250 1.4163 0.06388 0.05846 -0.0405 0.0060 1.0000
15.500 1.4079 0.06852 0.06323 -0.0415 0.0060 1.0000
15.750 1.3964 0.07382 0.06867 -0.0429 0.0059 1.0000
16.000 1.3854 0.07928 0.07426 -0.0446 0.0059 1.0000
16.250 1.3747 0.08491 0.08002 -0.0466 0.0058 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.5/12 AIRFOIL (hq3512-il)