HQ 3.5/9 AIRFOIL (hq359-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.5/9 AIRFOIL (hq359-il) Reynolds number: 1,000,000 Max Cl/Cd: 105.79 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq359-il-1000000-n5.txt Download as CSV file: xf-hq359-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.5/9 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3440 0.10840 0.10674 -0.0337 1.0000 0.0029
-9.750 -0.3417 0.10527 0.10363 -0.0343 1.0000 0.0027
-9.500 -0.3403 0.10143 0.09981 -0.0353 0.9998 0.0024
-9.000 -0.3241 0.09222 0.09062 -0.0419 0.9901 0.0022
-8.750 -0.3145 0.08730 0.08570 -0.0461 0.9855 0.0022
-8.500 -0.3040 0.08273 0.08114 -0.0505 0.9795 0.0022
-8.250 -0.2893 0.07763 0.07604 -0.0565 0.9748 0.0022
-8.000 -0.2719 0.07214 0.07054 -0.0640 0.9681 0.0022
-7.750 -0.2517 0.06663 0.06502 -0.0726 0.9582 0.0023
-7.500 -0.2349 0.06136 0.05970 -0.0811 0.9412 0.0025
-7.250 -0.2238 0.05548 0.05374 -0.0901 0.9212 0.0029
-7.000 -0.2169 0.04950 0.04762 -0.0962 0.9053 0.0029
-6.750 -0.2076 0.04378 0.04172 -0.1002 0.8931 0.0031
-6.500 -0.1953 0.03806 0.03578 -0.1030 0.8832 0.0031
-6.250 -0.1800 0.03201 0.02942 -0.1050 0.8749 0.0032
-6.000 -0.1641 0.02318 0.02003 -0.1059 0.8669 0.0035
-5.750 -0.1440 0.01664 0.01278 -0.1059 0.8606 0.0037
-5.500 -0.1188 0.01470 0.01049 -0.1058 0.8547 0.0038
-5.250 -0.0929 0.01323 0.00871 -0.1057 0.8495 0.0038
-5.000 -0.0671 0.01149 0.00665 -0.1056 0.8437 0.0040
-4.750 -0.0409 0.01039 0.00537 -0.1055 0.8383 0.0044
-4.500 -0.0136 0.00994 0.00485 -0.1055 0.8329 0.0048
-4.250 0.0137 0.00952 0.00434 -0.1054 0.8272 0.0052
-4.000 0.0410 0.00911 0.00383 -0.1054 0.8217 0.0056
-3.750 0.0686 0.00874 0.00338 -0.1053 0.8156 0.0060
-3.500 0.0961 0.00844 0.00301 -0.1053 0.8097 0.0064
-3.250 0.1239 0.00821 0.00272 -0.1053 0.8029 0.0067
-3.000 0.1514 0.00791 0.00233 -0.1052 0.7946 0.0068
-2.750 0.1787 0.00768 0.00198 -0.1051 0.7827 0.0069
-2.500 0.2060 0.00748 0.00166 -0.1049 0.7678 0.0071
-2.250 0.2334 0.00732 0.00137 -0.1047 0.7537 0.0081
-2.000 0.2612 0.00721 0.00120 -0.1047 0.7429 0.0097
-1.750 0.2889 0.00712 0.00108 -0.1046 0.7316 0.0142
-1.500 0.3165 0.00702 0.00099 -0.1046 0.7206 0.0333
-1.250 0.3440 0.00693 0.00093 -0.1046 0.7095 0.0534
-1.000 0.3715 0.00686 0.00088 -0.1046 0.6979 0.0794
-0.750 0.3989 0.00668 0.00083 -0.1046 0.6860 0.1473
-0.500 0.4252 0.00599 0.00079 -0.1049 0.6712 0.4118
-0.250 0.4519 0.00578 0.00081 -0.1049 0.6553 0.5217
0.000 0.4788 0.00572 0.00085 -0.1047 0.6406 0.5858
0.250 0.5058 0.00571 0.00091 -0.1046 0.6262 0.6330
0.500 0.5325 0.00574 0.00097 -0.1044 0.6088 0.6735
0.750 0.5591 0.00584 0.00102 -0.1041 0.5884 0.6897
1.250 0.6124 0.00612 0.00113 -0.1037 0.5414 0.7057
1.500 0.6383 0.00632 0.00122 -0.1034 0.5098 0.7141
1.750 0.6641 0.00654 0.00132 -0.1031 0.4797 0.7231
2.000 0.6898 0.00677 0.00144 -0.1027 0.4473 0.7318
2.250 0.7153 0.00702 0.00158 -0.1024 0.4153 0.7413
2.500 0.7410 0.00725 0.00171 -0.1021 0.3873 0.7518
2.750 0.7662 0.00751 0.00187 -0.1017 0.3550 0.7632
3.000 0.7910 0.00780 0.00204 -0.1013 0.3186 0.7752
3.250 0.8159 0.00808 0.00221 -0.1008 0.2894 0.7883
3.500 0.8414 0.00828 0.00239 -0.1005 0.2718 0.8030
3.750 0.8669 0.00844 0.00255 -0.1001 0.2589 0.8206
4.000 0.8918 0.00859 0.00272 -0.0996 0.2443 0.8449
4.250 0.9130 0.00863 0.00289 -0.0983 0.2229 0.9259
4.500 0.9412 0.00902 0.00313 -0.0987 0.1874 1.0000
4.750 0.9658 0.00937 0.00336 -0.0983 0.1622 1.0000
5.000 0.9900 0.00976 0.00362 -0.0979 0.1371 1.0000
5.250 1.0127 0.01029 0.00397 -0.0972 0.1029 1.0000
5.500 1.0337 0.01100 0.00444 -0.0962 0.0609 1.0000
5.750 1.0527 0.01195 0.00512 -0.0949 0.0128 1.0000
6.000 1.0763 0.01239 0.00552 -0.0943 0.0041 1.0000
6.250 1.1009 0.01270 0.00585 -0.0938 0.0034 1.0000
6.500 1.1250 0.01305 0.00624 -0.0933 0.0029 1.0000
6.750 1.1485 0.01346 0.00670 -0.0926 0.0025 1.0000
7.000 1.1716 0.01390 0.00720 -0.0919 0.0022 1.0000
7.250 1.1935 0.01447 0.00787 -0.0910 0.0020 1.0000
7.500 1.2138 0.01522 0.00872 -0.0898 0.0018 1.0000
7.750 1.2360 0.01568 0.00923 -0.0890 0.0018 1.0000
8.000 1.2582 0.01614 0.00972 -0.0882 0.0016 1.0000
8.250 1.2790 0.01671 0.01035 -0.0872 0.0015 1.0000
8.500 1.2991 0.01733 0.01104 -0.0861 0.0015 1.0000
8.750 1.3179 0.01803 0.01182 -0.0848 0.0014 1.0000
9.000 1.3359 0.01877 0.01263 -0.0833 0.0013 1.0000
9.250 1.3527 0.01956 0.01350 -0.0817 0.0013 1.0000
9.500 1.3683 0.02038 0.01444 -0.0800 0.0012 1.0000
9.750 1.3808 0.02125 0.01540 -0.0777 0.0011 1.0000
10.000 1.3919 0.02211 0.01634 -0.0752 0.0011 1.0000
10.250 1.4023 0.02300 0.01732 -0.0727 0.0010 1.0000
10.500 1.4118 0.02396 0.01837 -0.0702 0.0010 1.0000
10.750 1.4207 0.02499 0.01950 -0.0678 0.0009 1.0000
11.000 1.4279 0.02618 0.02079 -0.0653 0.0009 1.0000
11.250 1.4357 0.02734 0.02204 -0.0631 0.0009 1.0000
11.500 1.4412 0.02874 0.02354 -0.0607 0.0008 1.0000
11.750 1.4443 0.03038 0.02529 -0.0584 0.0008 1.0000
12.000 1.4453 0.03228 0.02732 -0.0560 0.0008 1.0000
12.250 1.4374 0.03511 0.03032 -0.0533 0.0007 1.0000
12.500 1.4205 0.03909 0.03452 -0.0505 0.0006 1.0000
12.750 1.4259 0.04091 0.03645 -0.0495 0.0006 1.0000
13.000 1.4225 0.04381 0.03952 -0.0484 0.0006 1.0000
13.250 1.4221 0.04648 0.04232 -0.0476 0.0006 1.0000
13.500 1.4139 0.05026 0.04627 -0.0471 0.0006 1.0000
13.750 1.4076 0.05403 0.05020 -0.0471 0.0006 1.0000
14.000 1.4013 0.05803 0.05434 -0.0476 0.0006 1.0000
14.250 1.3916 0.06280 0.05927 -0.0486 0.0006 1.0000
14.500 1.3809 0.06806 0.06469 -0.0502 0.0006 1.0000
14.750 1.3672 0.07424 0.07103 -0.0525 0.0006 1.0000
15.000 1.3578 0.08009 0.07703 -0.0551 0.0006 1.0000
15.250 1.3412 0.08765 0.08475 -0.0588 0.0006 1.0000
15.500 1.3262 0.09535 0.09261 -0.0629 0.0006 1.0000
15.750 1.3086 0.10397 0.10138 -0.0676 0.0006 1.0000
16.000 1.2921 0.11263 0.11018 -0.0724 0.0006 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.5/9 AIRFOIL (hq359-il)