HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 500,000 Max Cl/Cd: 50.25 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq07-il-500000.txt Download as CSV file: xf-hq07-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6514 0.08050 0.07837 -0.0005 1.0000 0.0122
-8.500 -0.6579 0.07352 0.07136 -0.0072 1.0000 0.0122
-8.250 -0.6704 0.06894 0.06671 -0.0096 1.0000 0.0122
-8.000 -0.6780 0.06449 0.06215 -0.0107 1.0000 0.0122
-7.750 -0.6994 0.05600 0.05345 -0.0123 1.0000 0.0125
-7.500 -0.7080 0.04914 0.04633 -0.0128 1.0000 0.0130
-7.250 -0.7047 0.04418 0.04112 -0.0127 1.0000 0.0133
-7.000 -0.6952 0.04003 0.03672 -0.0122 1.0000 0.0136
-6.750 -0.6826 0.03607 0.03247 -0.0115 1.0000 0.0139
-6.500 -0.6649 0.03216 0.02814 -0.0101 1.0000 0.0131
-6.250 -0.6515 0.02634 0.02191 -0.0087 1.0000 0.0091
-6.000 -0.6352 0.01882 0.01342 -0.0058 1.0000 0.0070
-5.750 -0.6122 0.01643 0.01070 -0.0047 1.0000 0.0069
-5.500 -0.5894 0.01450 0.00853 -0.0036 1.0000 0.0071
-5.250 -0.5659 0.01319 0.00708 -0.0027 1.0000 0.0079
-5.000 -0.5412 0.01236 0.00617 -0.0021 1.0000 0.0100
-4.750 -0.5176 0.01123 0.00490 -0.0012 1.0000 0.0162
-4.500 -0.4899 0.01146 0.00516 -0.0012 1.0000 0.0220
-4.250 -0.4662 0.01065 0.00432 -0.0007 1.0000 0.0290
-4.000 -0.4410 0.01017 0.00376 -0.0001 1.0000 0.0318
-3.750 -0.4163 0.00962 0.00313 0.0005 1.0000 0.0367
-3.500 -0.3919 0.00902 0.00258 0.0012 1.0000 0.0555
-3.250 -0.3697 0.00813 0.00216 0.0018 1.0000 0.1551
-3.000 -0.3471 0.00748 0.00187 0.0025 1.0000 0.2580
-2.750 -0.3260 0.00671 0.00164 0.0033 1.0000 0.3947
-2.500 -0.3055 0.00609 0.00150 0.0044 1.0000 0.5197
-2.250 -0.2824 0.00568 0.00150 0.0051 0.9985 0.6331
-2.000 -0.2444 0.00548 0.00149 0.0029 0.9897 0.6967
-1.750 -0.2058 0.00537 0.00144 0.0006 0.9817 0.7343
-1.500 -0.1699 0.00528 0.00142 -0.0010 0.9709 0.7650
-1.250 -0.1351 0.00521 0.00141 -0.0023 0.9583 0.7904
-1.000 -0.1030 0.00518 0.00138 -0.0030 0.9423 0.8060
-0.750 -0.0753 0.00516 0.00135 -0.0028 0.9235 0.8200
-0.500 -0.0494 0.00514 0.00132 -0.0020 0.9053 0.8350
-0.250 -0.0243 0.00512 0.00131 -0.0011 0.8883 0.8508
0.000 0.0000 0.00512 0.00130 0.0000 0.8692 0.8694
0.250 0.0244 0.00512 0.00131 0.0011 0.8507 0.8883
0.500 0.0495 0.00514 0.00132 0.0020 0.8350 0.9053
0.750 0.0753 0.00516 0.00135 0.0027 0.8200 0.9235
1.000 0.1031 0.00518 0.00138 0.0030 0.8060 0.9424
1.250 0.1352 0.00521 0.00141 0.0023 0.7904 0.9582
1.500 0.1699 0.00528 0.00142 0.0010 0.7650 0.9710
1.750 0.2058 0.00537 0.00144 -0.0006 0.7342 0.9818
2.000 0.2445 0.00548 0.00149 -0.0029 0.6971 0.9897
2.250 0.2823 0.00569 0.00150 -0.0051 0.6306 0.9986
2.500 0.3055 0.00608 0.00150 -0.0044 0.5216 1.0000
2.750 0.3259 0.00672 0.00164 -0.0033 0.3934 1.0000
3.000 0.3471 0.00748 0.00187 -0.0025 0.2581 1.0000
3.250 0.3696 0.00813 0.00216 -0.0018 0.1551 1.0000
3.500 0.3918 0.00902 0.00259 -0.0012 0.0552 1.0000
3.750 0.4162 0.00963 0.00314 -0.0005 0.0365 1.0000
4.000 0.4410 0.01016 0.00376 0.0001 0.0318 1.0000
4.250 0.4661 0.01065 0.00432 0.0007 0.0290 1.0000
4.500 0.4899 0.01147 0.00517 0.0012 0.0220 1.0000
4.750 0.5176 0.01124 0.00491 0.0012 0.0162 1.0000
5.000 0.5411 0.01237 0.00618 0.0021 0.0101 1.0000
5.250 0.5658 0.01320 0.00710 0.0027 0.0080 1.0000
5.500 0.5893 0.01451 0.00855 0.0036 0.0071 1.0000
5.750 0.6122 0.01643 0.01070 0.0047 0.0069 1.0000
6.000 0.6352 0.01878 0.01338 0.0058 0.0070 1.0000
6.250 0.6515 0.02639 0.02196 0.0087 0.0091 1.0000
6.500 0.6465 0.02006 0.01635 0.0104 0.0136 1.0000
6.750 0.6591 0.02490 0.02162 0.0116 0.0140 1.0000
7.000 0.6693 0.02927 0.02627 0.0122 0.0138 1.0000
7.250 0.6762 0.03381 0.03104 0.0125 0.0135 1.0000
7.500 0.6786 0.03883 0.03627 0.0127 0.0133 1.0000
7.750 0.6743 0.04435 0.04198 0.0126 0.0130 1.0000
8.000 0.6646 0.04986 0.04765 0.0122 0.0129 1.0000
8.250 0.6499 0.05431 0.05222 0.0120 0.0128 1.0000
8.500 0.6321 0.05881 0.05681 0.0098 0.0129 1.0000
8.750 0.6172 0.06716 0.06523 0.0026 0.0132 1.0000
9.000 0.6067 0.07484 0.07289 -0.0023 0.0133 1.0000
9.250 0.5992 0.08136 0.07940 -0.0056 0.0133 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 0/7 AIRFOIL (hq07-il)