HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 500,000 Max Cl/Cd: 46.12 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq07-il-500000-n5.txt Download as CSV file: xf-hq07-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 0/7 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6655 0.12429 0.12198 0.0256 1.0000 0.0081
-10.750 -0.6650 0.11949 0.11719 0.0236 1.0000 0.0071
-10.500 -0.6652 0.11433 0.11204 0.0213 1.0000 0.0063
-10.250 -0.6701 0.10759 0.10534 0.0183 1.0000 0.0054
-10.000 -0.6792 0.09916 0.09694 0.0145 1.0000 0.0045
-9.750 -0.6806 0.09374 0.09155 0.0118 1.0000 0.0043
-9.500 -0.6816 0.08872 0.08655 0.0090 1.0000 0.0043
-9.250 -0.6800 0.08425 0.08210 0.0065 1.0000 0.0041
-9.000 -0.6830 0.07879 0.07666 0.0028 1.0000 0.0041
-8.750 -0.6918 0.07131 0.06922 -0.0041 1.0000 0.0041
-8.500 -0.7084 0.06480 0.06272 -0.0089 1.0000 0.0041
-8.250 -0.7100 0.06071 0.05854 -0.0101 1.0000 0.0040
-8.000 -0.7133 0.05628 0.05397 -0.0111 1.0000 0.0039
-7.750 -0.7197 0.04997 0.04745 -0.0119 1.0000 0.0038
-7.500 -0.7199 0.04381 0.04100 -0.0118 1.0000 0.0037
-7.250 -0.7161 0.03760 0.03442 -0.0111 1.0000 0.0036
-7.000 -0.7077 0.03190 0.02827 -0.0099 1.0000 0.0035
-6.750 -0.6949 0.02697 0.02276 -0.0086 1.0000 0.0033
-6.500 -0.6778 0.02306 0.01835 -0.0073 1.0000 0.0032
-6.250 -0.6578 0.01994 0.01479 -0.0061 1.0000 0.0032
-6.000 -0.6358 0.01752 0.01199 -0.0050 1.0000 0.0031
-5.750 -0.6129 0.01560 0.00979 -0.0041 1.0000 0.0031
-5.500 -0.5896 0.01412 0.00808 -0.0032 1.0000 0.0030
-5.250 -0.5661 0.01292 0.00670 -0.0024 1.0000 0.0030
-5.000 -0.5421 0.01196 0.00558 -0.0016 1.0000 0.0030
-4.750 -0.5174 0.01122 0.00470 -0.0010 1.0000 0.0031
-4.500 -0.4921 0.01067 0.00402 -0.0005 1.0000 0.0032
-4.250 -0.4665 0.01026 0.00350 0.0000 1.0000 0.0033
-4.000 -0.4408 0.00993 0.00310 0.0004 1.0000 0.0037
-3.750 -0.4156 0.00955 0.00268 0.0009 1.0000 0.0102
-3.500 -0.3904 0.00930 0.00242 0.0014 1.0000 0.0157
-3.250 -0.3598 0.00909 0.00222 0.0006 0.9935 0.0213
-3.000 -0.3271 0.00851 0.00196 -0.0009 0.9811 0.0796
-2.750 -0.2942 0.00787 0.00171 -0.0025 0.9678 0.1746
-2.500 -0.2621 0.00726 0.00149 -0.0039 0.9510 0.2795
-2.250 -0.2336 0.00656 0.00130 -0.0045 0.9289 0.4238
-2.000 -0.2080 0.00618 0.00112 -0.0041 0.9030 0.5089
-1.500 -0.1591 0.00577 0.00103 -0.0024 0.8600 0.6496
-1.250 -0.1333 0.00569 0.00101 -0.0018 0.8438 0.6832
-1.000 -0.1067 0.00568 0.00097 -0.0014 0.8277 0.6996
-0.750 -0.0800 0.00568 0.00095 -0.0011 0.8125 0.7153
-0.500 -0.0532 0.00566 0.00093 -0.0008 0.7991 0.7322
-0.250 -0.0265 0.00564 0.00091 -0.0004 0.7847 0.7487
0.000 0.0000 0.00564 0.00091 0.0000 0.7673 0.7674
0.250 0.0265 0.00564 0.00091 0.0004 0.7488 0.7846
0.500 0.0532 0.00566 0.00093 0.0008 0.7323 0.7991
0.750 0.0800 0.00568 0.00095 0.0011 0.7153 0.8125
1.000 0.1067 0.00568 0.00097 0.0014 0.6997 0.8276
1.250 0.1333 0.00569 0.00101 0.0018 0.6834 0.8438
1.500 0.1591 0.00577 0.00103 0.0024 0.6494 0.8603
1.750 0.1840 0.00589 0.00106 0.0032 0.5976 0.8792
2.000 0.2080 0.00618 0.00112 0.0041 0.5092 0.9031
2.250 0.2337 0.00656 0.00130 0.0045 0.4224 0.9291
2.500 0.2621 0.00727 0.00150 0.0039 0.2767 0.9514
2.750 0.2944 0.00787 0.00171 0.0025 0.1744 0.9681
3.000 0.3272 0.00852 0.00196 0.0009 0.0779 0.9813
3.250 0.3599 0.00908 0.00222 -0.0006 0.0221 0.9938
3.500 0.3902 0.00930 0.00242 -0.0013 0.0157 1.0000
3.750 0.4155 0.00956 0.00268 -0.0009 0.0101 1.0000
4.000 0.4407 0.00993 0.00310 -0.0004 0.0037 1.0000
4.250 0.4664 0.01026 0.00350 0.0000 0.0033 1.0000
4.500 0.4921 0.01067 0.00402 0.0005 0.0032 1.0000
4.750 0.5173 0.01122 0.00470 0.0010 0.0031 1.0000
5.000 0.5421 0.01196 0.00558 0.0016 0.0030 1.0000
5.250 0.5662 0.01290 0.00668 0.0023 0.0030 1.0000
5.500 0.5898 0.01410 0.00806 0.0032 0.0030 1.0000
5.750 0.6130 0.01562 0.00980 0.0040 0.0031 1.0000
6.000 0.6359 0.01753 0.01201 0.0050 0.0031 1.0000
6.250 0.6579 0.02001 0.01486 0.0061 0.0032 1.0000
6.500 0.6781 0.02307 0.01836 0.0072 0.0032 1.0000
6.750 0.6951 0.02702 0.02281 0.0085 0.0033 1.0000
7.000 0.7078 0.03203 0.02841 0.0099 0.0035 1.0000
7.250 0.7161 0.03780 0.03463 0.0110 0.0036 1.0000
7.500 0.7206 0.04378 0.04098 0.0117 0.0037 1.0000
7.750 0.7202 0.04999 0.04747 0.0118 0.0038 1.0000
8.000 0.7142 0.05630 0.05400 0.0110 0.0039 1.0000
8.250 0.7114 0.06061 0.05844 0.0100 0.0041 1.0000
8.500 0.7053 0.06537 0.06330 0.0085 0.0041 1.0000
8.750 0.6913 0.07170 0.06961 0.0035 0.0041 1.0000
9.000 0.6834 0.07892 0.07680 -0.0031 0.0041 1.0000
9.250 0.6812 0.08423 0.08208 -0.0066 0.0041 1.0000
9.500 0.6805 0.08914 0.08696 -0.0095 0.0042 1.0000
9.750 0.6811 0.09387 0.09168 -0.0120 0.0043 1.0000
10.000 0.6797 0.09949 0.09727 -0.0147 0.0046 1.0000
10.250 0.6693 0.10827 0.10601 -0.0187 0.0055 1.0000
10.500 0.6674 0.11392 0.11164 -0.0213 0.0060 1.0000
10.750 0.6660 0.11952 0.11723 -0.0237 0.0070 1.0000
11.000 0.5803 0.12105 0.11896 -0.0237 0.0058 1.0000
11.250 0.5781 0.12649 0.12440 -0.0258 0.0064 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 0/7 AIRFOIL (hq07-il)