HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 200,000 Max Cl/Cd: 36.73 at α=2.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq07-il-200000-n5.txt Download as CSV file: xf-hq07-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6752 0.08429 0.08092 0.0048 1.0000 0.0068
-8.750 -0.6817 0.07815 0.07483 0.0002 1.0000 0.0067
-8.500 -0.6931 0.07133 0.06802 -0.0064 1.0000 0.0066
-8.250 -0.7045 0.06593 0.06255 -0.0092 1.0000 0.0065
-8.000 -0.7107 0.06025 0.05682 -0.0112 1.0000 0.0064
-7.750 -0.7133 0.05453 0.05090 -0.0124 1.0000 0.0063
-7.500 -0.7121 0.04883 0.04492 -0.0127 1.0000 0.0061
-7.250 -0.7072 0.04310 0.03881 -0.0124 1.0000 0.0059
-7.000 -0.6990 0.03733 0.03246 -0.0115 1.0000 0.0057
-6.750 -0.6865 0.03219 0.02676 -0.0103 1.0000 0.0055
-6.500 -0.6701 0.02795 0.02194 -0.0090 1.0000 0.0054
-6.250 -0.6503 0.02455 0.01802 -0.0078 1.0000 0.0054
-6.000 -0.6285 0.02189 0.01492 -0.0067 1.0000 0.0054
-5.750 -0.6056 0.01976 0.01247 -0.0057 1.0000 0.0056
-5.500 -0.5825 0.01803 0.01050 -0.0048 1.0000 0.0060
-5.250 -0.5595 0.01657 0.00884 -0.0039 1.0000 0.0065
-5.000 -0.5363 0.01532 0.00740 -0.0030 1.0000 0.0078
-4.750 -0.5127 0.01425 0.00614 -0.0021 1.0000 0.0107
-4.500 -0.4881 0.01360 0.00542 -0.0016 1.0000 0.0179
-4.250 -0.4633 0.01313 0.00496 -0.0012 1.0000 0.0293
-4.000 -0.4386 0.01264 0.00447 -0.0008 1.0000 0.0421
-3.750 -0.4141 0.01207 0.00387 -0.0003 1.0000 0.0613
-3.500 -0.3905 0.01137 0.00338 0.0002 1.0000 0.1095
-3.250 -0.3678 0.01062 0.00302 0.0007 1.0000 0.2025
-3.000 -0.3456 0.00988 0.00268 0.0013 1.0000 0.3004
-2.750 -0.3268 0.00888 0.00246 0.0025 1.0000 0.4774
-2.500 -0.3067 0.00838 0.00242 0.0040 1.0000 0.5906
-2.250 -0.2853 0.00815 0.00239 0.0055 1.0000 0.6588
-2.000 -0.2623 0.00798 0.00238 0.0066 0.9980 0.7132
-1.750 -0.2290 0.00788 0.00228 0.0056 0.9864 0.7532
-1.500 -0.1942 0.00783 0.00223 0.0042 0.9754 0.7730
-1.250 -0.1590 0.00779 0.00218 0.0027 0.9636 0.7922
-1.000 -0.1239 0.00775 0.00215 0.0013 0.9490 0.8112
-0.750 -0.0907 0.00771 0.00214 0.0005 0.9330 0.8320
-0.500 -0.0596 0.00768 0.00212 0.0001 0.9179 0.8520
-0.250 -0.0293 0.00766 0.00211 -0.0001 0.9027 0.8691
0.000 0.0000 0.00765 0.00211 0.0000 0.8864 0.8865
0.250 0.0294 0.00766 0.00211 0.0001 0.8692 0.9028
0.500 0.0596 0.00768 0.00212 -0.0001 0.8521 0.9179
0.750 0.0907 0.00771 0.00214 -0.0005 0.8320 0.9331
1.000 0.1239 0.00775 0.00215 -0.0013 0.8112 0.9491
1.250 0.1590 0.00779 0.00218 -0.0027 0.7923 0.9638
1.500 0.1942 0.00783 0.00222 -0.0042 0.7730 0.9757
1.750 0.2290 0.00788 0.00227 -0.0056 0.7531 0.9867
2.000 0.2623 0.00798 0.00238 -0.0066 0.7130 0.9985
2.250 0.2847 0.00814 0.00238 -0.0054 0.6592 1.0000
2.500 0.3061 0.00838 0.00241 -0.0039 0.5921 1.0000
2.750 0.3262 0.00888 0.00246 -0.0024 0.4784 1.0000
3.000 0.3452 0.00985 0.00267 -0.0012 0.3047 1.0000
3.250 0.3673 0.01062 0.00301 -0.0006 0.2027 1.0000
3.500 0.3901 0.01139 0.00339 -0.0001 0.1079 1.0000
3.750 0.4138 0.01207 0.00386 0.0004 0.0613 1.0000
4.000 0.4383 0.01264 0.00447 0.0009 0.0421 1.0000
4.250 0.4632 0.01311 0.00493 0.0012 0.0286 1.0000
4.500 0.4881 0.01361 0.00541 0.0016 0.0179 1.0000
4.750 0.5126 0.01429 0.00618 0.0021 0.0103 1.0000
5.000 0.5364 0.01531 0.00739 0.0029 0.0079 1.0000
5.250 0.5597 0.01657 0.00884 0.0038 0.0066 1.0000
5.500 0.5828 0.01803 0.01050 0.0047 0.0060 1.0000
5.750 0.6059 0.01978 0.01249 0.0057 0.0056 1.0000
6.000 0.6288 0.02191 0.01494 0.0066 0.0054 1.0000
6.250 0.6508 0.02457 0.01804 0.0077 0.0054 1.0000
6.500 0.6706 0.02799 0.02199 0.0089 0.0054 1.0000
6.750 0.6871 0.03224 0.02681 0.0101 0.0055 1.0000
7.000 0.6995 0.03742 0.03256 0.0114 0.0057 1.0000
7.250 0.7080 0.04313 0.03874 0.0122 0.0059 1.0000
7.500 0.7126 0.04900 0.04510 0.0125 0.0061 1.0000
7.750 0.7142 0.05461 0.05099 0.0121 0.0063 1.0000
8.000 0.7118 0.06029 0.05687 0.0110 0.0064 1.0000
8.250 0.7053 0.06608 0.06270 0.0089 0.0065 1.0000
8.500 0.6936 0.07156 0.06826 0.0060 0.0066 1.0000
8.750 0.6824 0.07844 0.07511 -0.0006 0.0067 1.0000
9.000 0.6763 0.08449 0.08112 -0.0052 0.0068 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 0/7 AIRFOIL (hq07-il)