HT05 (ht05-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HT05 (ht05-il) Reynolds number: 500,000 Max Cl/Cd: 44.36 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ht05-il-500000.txt Download as CSV file: xf-ht05-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HT05
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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7414 0.13883 0.13653 0.0448 1.0000 0.0175
-11.000 -0.7385 0.13438 0.13209 0.0429 1.0000 0.0175
-7.250 -0.7415 0.03742 0.03419 -0.0096 1.0000 0.0160
-7.000 -0.7311 0.03040 0.02652 -0.0094 1.0000 0.0159
-6.750 -0.7154 0.02514 0.02062 -0.0085 1.0000 0.0159
-6.500 -0.6949 0.02159 0.01654 -0.0076 1.0000 0.0162
-6.250 -0.6715 0.01947 0.01406 -0.0068 1.0000 0.0168
-6.000 -0.6464 0.01819 0.01252 -0.0062 1.0000 0.0174
-5.750 -0.6230 0.01580 0.00978 -0.0054 1.0000 0.0182
-5.500 -0.5987 0.01426 0.00811 -0.0048 1.0000 0.0200
-5.250 -0.5726 0.01369 0.00748 -0.0044 1.0000 0.0219
-5.000 -0.5467 0.01289 0.00658 -0.0039 1.0000 0.0236
-4.750 -0.5205 0.01230 0.00590 -0.0034 1.0000 0.0251
-4.500 -0.4954 0.01114 0.00462 -0.0028 1.0000 0.0282
-4.250 -0.4691 0.01072 0.00419 -0.0024 1.0000 0.0324
-4.000 -0.4425 0.01040 0.00382 -0.0020 1.0000 0.0357
-3.750 -0.4166 0.00972 0.00313 -0.0015 1.0000 0.0445
-3.500 -0.3903 0.00929 0.00272 -0.0010 1.0000 0.0567
-3.250 -0.3641 0.00890 0.00239 -0.0006 1.0000 0.0773
-3.000 -0.3380 0.00852 0.00216 -0.0002 1.0000 0.1089
-2.750 -0.3120 0.00818 0.00199 0.0002 1.0000 0.1471
-2.500 -0.2859 0.00790 0.00184 0.0006 1.0000 0.1852
-2.250 -0.2600 0.00760 0.00171 0.0010 1.0000 0.2308
-2.000 -0.2343 0.00728 0.00160 0.0014 1.0000 0.2856
-1.750 -0.2089 0.00686 0.00150 0.0018 1.0000 0.3679
-1.500 -0.1841 0.00637 0.00147 0.0024 1.0000 0.4867
-1.250 -0.1587 0.00615 0.00153 0.0030 1.0000 0.5795
-1.000 -0.1323 0.00614 0.00153 0.0035 1.0000 0.6165
-0.750 -0.1062 0.00598 0.00144 0.0040 1.0000 0.6394
-0.500 -0.0807 0.00561 0.00133 0.0045 1.0000 0.6695
-0.250 -0.0524 0.00486 0.00151 0.0053 1.0000 0.9622
0.000 0.0003 0.00490 0.00156 -0.0001 0.9979 0.9981
0.250 0.0525 0.00486 0.00151 -0.0053 0.9622 1.0000
0.500 0.0808 0.00562 0.00133 -0.0045 0.6690 1.0000
0.750 0.1063 0.00599 0.00144 -0.0040 0.6391 1.0000
1.000 0.1325 0.00614 0.00153 -0.0035 0.6165 1.0000
1.250 0.1589 0.00615 0.00153 -0.0030 0.5795 1.0000
1.500 0.1842 0.00637 0.00147 -0.0024 0.4866 1.0000
1.750 0.2090 0.00686 0.00150 -0.0018 0.3671 1.0000
2.000 0.2344 0.00728 0.00160 -0.0014 0.2851 1.0000
2.250 0.2602 0.00760 0.00171 -0.0010 0.2308 1.0000
2.500 0.2860 0.00790 0.00184 -0.0006 0.1853 1.0000
2.750 0.3121 0.00819 0.00199 -0.0002 0.1470 1.0000
3.000 0.3381 0.00852 0.00216 0.0002 0.1088 1.0000
3.250 0.3642 0.00890 0.00240 0.0006 0.0773 1.0000
3.500 0.3904 0.00929 0.00272 0.0010 0.0566 1.0000
3.750 0.4167 0.00972 0.00313 0.0015 0.0444 1.0000
4.000 0.4427 0.01039 0.00380 0.0019 0.0358 1.0000
4.250 0.4693 0.01071 0.00418 0.0024 0.0323 1.0000
4.500 0.4955 0.01117 0.00464 0.0028 0.0280 1.0000
4.750 0.5207 0.01228 0.00588 0.0034 0.0251 1.0000
5.000 0.5469 0.01289 0.00658 0.0039 0.0236 1.0000
5.250 0.5728 0.01368 0.00747 0.0044 0.0219 1.0000
5.500 0.5988 0.01426 0.00811 0.0048 0.0201 1.0000
5.750 0.6228 0.01602 0.01002 0.0054 0.0182 1.0000
6.000 0.6466 0.01818 0.01253 0.0062 0.0174 1.0000
6.250 0.6717 0.01948 0.01407 0.0068 0.0168 1.0000
6.500 0.6951 0.02164 0.01659 0.0076 0.0162 1.0000
6.750 0.7154 0.02525 0.02074 0.0085 0.0159 1.0000
7.000 0.7314 0.03039 0.02651 0.0094 0.0159 1.0000
7.250 0.7420 0.03737 0.03414 0.0096 0.0159 1.0000
9.500 0.6071 0.09822 0.09609 -0.0206 0.0222 1.0000
9.750 0.6049 0.10276 0.10062 -0.0219 0.0213 1.0000
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Polar data table (+)
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