HQ 3.0/12 AIRFOIL (hq3012-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.0/12 AIRFOIL (hq3012-il) Reynolds number: 500,000 Max Cl/Cd: 88.26 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq3012-il-500000-n5.txt Download as CSV file: xf-hq3012-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.0/12 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7029 0.04537 0.04250 -0.0769 1.0000 0.0062
-11.750 -0.7259 0.04027 0.03724 -0.0794 1.0000 0.0062
-11.500 -0.7461 0.03753 0.03443 -0.0784 1.0000 0.0063
-11.250 -0.7409 0.03331 0.02991 -0.0836 0.9873 0.0063
-11.000 -0.7206 0.03112 0.02755 -0.0864 0.9810 0.0065
-10.750 -0.6995 0.02881 0.02502 -0.0890 0.9745 0.0067
-10.500 -0.6766 0.02686 0.02283 -0.0910 0.9684 0.0069
-10.250 -0.6541 0.02500 0.02074 -0.0924 0.9609 0.0072
-10.000 -0.6311 0.02329 0.01878 -0.0934 0.9534 0.0073
-9.750 -0.6079 0.02183 0.01709 -0.0940 0.9450 0.0076
-9.500 -0.5854 0.02061 0.01567 -0.0941 0.9359 0.0078
-9.250 -0.5620 0.01948 0.01434 -0.0942 0.9277 0.0080
-9.000 -0.5395 0.01853 0.01322 -0.0939 0.9187 0.0081
-8.750 -0.5163 0.01766 0.01217 -0.0935 0.9107 0.0083
-8.500 -0.4930 0.01688 0.01125 -0.0931 0.9027 0.0084
-8.250 -0.4711 0.01590 0.01013 -0.0926 0.8950 0.0088
-8.000 -0.4473 0.01524 0.00937 -0.0923 0.8874 0.0092
-7.750 -0.4229 0.01468 0.00872 -0.0919 0.8802 0.0097
-7.500 -0.3980 0.01419 0.00814 -0.0916 0.8732 0.0102
-7.250 -0.3730 0.01368 0.00751 -0.0913 0.8667 0.0106
-7.000 -0.3478 0.01319 0.00692 -0.0909 0.8598 0.0111
-6.750 -0.3221 0.01277 0.00639 -0.0906 0.8536 0.0115
-6.500 -0.2966 0.01229 0.00584 -0.0903 0.8470 0.0122
-6.000 -0.2443 0.01157 0.00499 -0.0899 0.8347 0.0149
-5.750 -0.2180 0.01124 0.00464 -0.0897 0.8286 0.0180
-5.500 -0.1913 0.01096 0.00436 -0.0896 0.8228 0.0230
-5.250 -0.1642 0.01077 0.00413 -0.0895 0.8165 0.0280
-5.000 -0.1372 0.01055 0.00389 -0.0894 0.8110 0.0330
-4.750 -0.1098 0.01038 0.00367 -0.0893 0.8048 0.0362
-4.500 -0.0827 0.01017 0.00343 -0.0892 0.7989 0.0401
-4.250 -0.0553 0.01001 0.00326 -0.0892 0.7932 0.0454
-4.000 -0.0279 0.00984 0.00305 -0.0891 0.7871 0.0499
-3.750 -0.0005 0.00966 0.00285 -0.0891 0.7816 0.0566
-3.500 0.0270 0.00949 0.00268 -0.0890 0.7755 0.0647
-3.250 0.0545 0.00933 0.00251 -0.0890 0.7696 0.0761
-3.000 0.0820 0.00914 0.00235 -0.0890 0.7636 0.0929
-2.750 0.1093 0.00894 0.00220 -0.0889 0.7569 0.1181
-2.500 0.1362 0.00867 0.00206 -0.0889 0.7499 0.1647
-2.250 0.1631 0.00838 0.00193 -0.0889 0.7422 0.2270
-2.000 0.1899 0.00806 0.00179 -0.0889 0.7355 0.2965
-1.750 0.2157 0.00752 0.00169 -0.0889 0.7289 0.4377
-1.500 0.2426 0.00734 0.00168 -0.0888 0.7230 0.5095
-1.250 0.2702 0.00727 0.00168 -0.0887 0.7166 0.5459
-1.000 0.2978 0.00725 0.00168 -0.0886 0.7101 0.5735
-0.750 0.3251 0.00722 0.00172 -0.0884 0.7018 0.6070
-0.500 0.3524 0.00724 0.00173 -0.0882 0.6922 0.6283
-0.250 0.3799 0.00727 0.00173 -0.0880 0.6823 0.6386
0.000 0.4076 0.00730 0.00174 -0.0879 0.6730 0.6501
0.250 0.4351 0.00734 0.00176 -0.0878 0.6652 0.6613
0.500 0.4629 0.00736 0.00178 -0.0877 0.6567 0.6693
0.750 0.4905 0.00741 0.00180 -0.0876 0.6480 0.6759
1.000 0.5180 0.00746 0.00183 -0.0875 0.6382 0.6816
1.250 0.5452 0.00752 0.00187 -0.0873 0.6262 0.6880
1.500 0.5721 0.00760 0.00190 -0.0871 0.6108 0.6946
1.750 0.5986 0.00769 0.00195 -0.0868 0.5945 0.7010
2.000 0.6254 0.00778 0.00201 -0.0866 0.5799 0.7080
2.250 0.6517 0.00789 0.00208 -0.0862 0.5635 0.7147
2.500 0.6777 0.00802 0.00217 -0.0859 0.5438 0.7221
2.750 0.7032 0.00818 0.00227 -0.0854 0.5224 0.7292
3.000 0.7282 0.00838 0.00239 -0.0849 0.4977 0.7372
3.250 0.7525 0.00861 0.00254 -0.0843 0.4706 0.7448
3.750 0.8003 0.00915 0.00291 -0.0830 0.4174 0.7617
4.000 0.8243 0.00942 0.00311 -0.0823 0.3929 0.7709
4.250 0.8479 0.00971 0.00333 -0.0816 0.3670 0.7806
4.500 0.8713 0.00998 0.00355 -0.0809 0.3439 0.7909
4.750 0.8943 0.01027 0.00378 -0.0801 0.3197 0.8025
5.000 0.9170 0.01057 0.00403 -0.0793 0.2983 0.8156
5.250 0.9398 0.01082 0.00429 -0.0784 0.2815 0.8312
5.500 0.9619 0.01106 0.00454 -0.0774 0.2655 0.8523
5.750 0.9824 0.01122 0.00478 -0.0760 0.2503 0.8927
6.000 1.0141 0.01149 0.00508 -0.0772 0.2330 1.0000
6.250 1.0368 0.01186 0.00538 -0.0765 0.2163 1.0000
6.750 1.0806 0.01267 0.00603 -0.0748 0.1812 1.0000
7.000 1.1009 0.01316 0.00640 -0.0737 0.1611 1.0000
7.250 1.1218 0.01359 0.00679 -0.0727 0.1468 1.0000
7.500 1.1427 0.01401 0.00716 -0.0717 0.1342 1.0000
7.750 1.1613 0.01455 0.00760 -0.0704 0.1165 1.0000
8.000 1.1799 0.01507 0.00804 -0.0691 0.1011 1.0000
8.250 1.1978 0.01558 0.00849 -0.0676 0.0887 1.0000
8.500 1.2140 0.01610 0.00898 -0.0659 0.0776 1.0000
8.750 1.2286 0.01669 0.00951 -0.0639 0.0653 1.0000
9.000 1.2424 0.01734 0.01010 -0.0618 0.0538 1.0000
9.250 1.2565 0.01798 0.01071 -0.0599 0.0451 1.0000
9.500 1.2705 0.01864 0.01135 -0.0580 0.0376 1.0000
9.750 1.2837 0.01937 0.01206 -0.0561 0.0305 1.0000
10.000 1.2964 0.02014 0.01282 -0.0542 0.0250 1.0000
10.250 1.3082 0.02099 0.01365 -0.0523 0.0190 1.0000
10.500 1.3206 0.02182 0.01450 -0.0505 0.0151 1.0000
10.750 1.3280 0.02301 0.01565 -0.0483 0.0068 1.0000
11.000 1.3365 0.02416 0.01683 -0.0462 0.0048 1.0000
11.250 1.3476 0.02516 0.01789 -0.0446 0.0045 1.0000
11.500 1.3577 0.02626 0.01907 -0.0430 0.0041 1.0000
11.750 1.3674 0.02742 0.02031 -0.0414 0.0039 1.0000
12.000 1.3764 0.02867 0.02166 -0.0399 0.0037 1.0000
12.250 1.3852 0.02997 0.02305 -0.0385 0.0036 1.0000
12.500 1.3936 0.03134 0.02451 -0.0371 0.0035 1.0000
12.750 1.4005 0.03288 0.02617 -0.0358 0.0033 1.0000
13.000 1.4077 0.03443 0.02782 -0.0347 0.0033 1.0000
13.250 1.4130 0.03620 0.02970 -0.0335 0.0032 1.0000
13.500 1.4175 0.03812 0.03172 -0.0325 0.0032 1.0000
13.750 1.4206 0.04022 0.03395 -0.0316 0.0031 1.0000
14.000 1.4227 0.04250 0.03634 -0.0308 0.0030 1.0000
14.250 1.4236 0.04499 0.03895 -0.0302 0.0030 1.0000
14.500 1.4233 0.04771 0.04181 -0.0298 0.0029 1.0000
14.750 1.4218 0.05068 0.04490 -0.0296 0.0029 1.0000
15.000 1.4196 0.05388 0.04823 -0.0296 0.0028 1.0000
15.250 1.4141 0.05765 0.05214 -0.0300 0.0027 1.0000
15.500 1.4100 0.06140 0.05601 -0.0306 0.0027 1.0000
15.750 1.4024 0.06583 0.06058 -0.0316 0.0027 1.0000
16.000 1.3933 0.07071 0.06560 -0.0330 0.0026 1.0000
16.250 1.3845 0.07575 0.07079 -0.0347 0.0027 1.0000
16.500 1.3746 0.08118 0.07636 -0.0368 0.0026 1.0000
16.750 1.3619 0.08736 0.08268 -0.0394 0.0026 1.0000
17.000 1.3456 0.09440 0.08988 -0.0426 0.0025 1.0000
17.250 1.3324 0.10109 0.09671 -0.0458 0.0025 1.0000
17.500 1.3186 0.10808 0.10385 -0.0493 0.0025 1.0000
17.750 1.3035 0.11548 0.11139 -0.0531 0.0025 1.0000
18.000 1.2896 0.12277 0.11882 -0.0569 0.0025 1.0000
18.250 1.2751 0.13033 0.12651 -0.0610 0.0025 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.0/12 AIRFOIL (hq3012-il)