HT22 (ht22-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: HT22 (ht22-il) Reynolds number: 1,000,000 Max Cl/Cd: 75.46 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ht22-il-1000000.txt Download as CSV file: xf-ht22-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HT22
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5589 0.07998 0.07852 0.0224 1.0001 0.0106
-7.750 -0.5620 0.07527 0.07382 0.0205 1.0001 0.0106
-7.500 -0.5662 0.07050 0.06907 0.0181 1.0001 0.0106
-7.250 -0.5686 0.06508 0.06366 0.0135 1.0001 0.0107
-7.000 -0.5651 0.05870 0.05725 0.0075 1.0001 0.0107
-6.750 -0.5583 0.05165 0.05014 0.0016 1.0001 0.0107
-6.500 -0.5484 0.04437 0.04277 -0.0034 1.0001 0.0107
-6.250 -0.5364 0.03671 0.03497 -0.0075 1.0001 0.0107
-6.000 -0.5294 0.02812 0.02616 -0.0113 1.0001 0.0114
-5.750 -0.5104 0.02464 0.02256 -0.0124 1.0001 0.0118
-5.500 -0.5184 0.02527 0.02211 -0.0132 1.0001 0.0114
-5.250 -0.4955 0.01733 0.01331 -0.0130 1.0001 0.0097
-5.000 -0.4691 0.01492 0.01050 -0.0128 1.0001 0.0101
-4.750 -0.4423 0.01314 0.00842 -0.0124 1.0001 0.0104
-4.500 -0.4147 0.01236 0.00749 -0.0122 1.0001 0.0107
-4.250 -0.3884 0.01036 0.00525 -0.0119 1.0001 0.0116
-4.000 -0.3609 0.00989 0.00475 -0.0117 1.0001 0.0123
-3.750 -0.3333 0.00941 0.00422 -0.0115 1.0001 0.0130
-3.500 -0.3056 0.00893 0.00370 -0.0113 1.0001 0.0138
-3.250 -0.2777 0.00868 0.00343 -0.0112 1.0001 0.0146
-3.000 -0.2502 0.00795 0.00265 -0.0110 1.0001 0.0167
-2.750 -0.2223 0.00767 0.00237 -0.0109 1.0001 0.0184
-2.500 -0.1943 0.00745 0.00213 -0.0107 1.0001 0.0199
-2.250 -0.1663 0.00703 0.00170 -0.0106 1.0001 0.0237
-2.000 -0.1384 0.00689 0.00158 -0.0104 1.0001 0.0271
-1.750 -0.1104 0.00660 0.00131 -0.0103 1.0001 0.0349
-1.500 -0.0824 0.00638 0.00115 -0.0102 1.0001 0.0464
-1.250 -0.0546 0.00616 0.00104 -0.0100 1.0001 0.0718
-1.000 -0.0223 0.00582 0.00094 -0.0111 0.9837 0.1378
-0.750 0.0072 0.00543 0.00086 -0.0113 0.8995 0.2808
-0.500 0.0292 0.00517 0.00080 -0.0099 0.7991 0.4593
-0.250 0.0537 0.00474 0.00077 -0.0094 0.7052 0.6899
0.000 0.0725 0.00427 0.00078 -0.0070 0.6294 0.9146
0.250 0.1059 0.00442 0.00073 -0.0079 0.5594 0.9999
0.500 0.1336 0.00463 0.00073 -0.0078 0.5074 0.9999
0.750 0.1614 0.00482 0.00073 -0.0077 0.4651 0.9999
1.000 0.1893 0.00498 0.00075 -0.0076 0.4299 0.9999
1.250 0.2172 0.00514 0.00077 -0.0076 0.3990 0.9999
1.500 0.2452 0.00529 0.00080 -0.0075 0.3702 0.9999
1.750 0.2731 0.00545 0.00084 -0.0074 0.3422 0.9999
2.000 0.3010 0.00561 0.00089 -0.0073 0.3141 0.9999
2.250 0.3289 0.00578 0.00095 -0.0073 0.2883 0.9999
2.500 0.3568 0.00597 0.00103 -0.0072 0.2587 0.9999
2.750 0.3846 0.00617 0.00112 -0.0072 0.2291 0.9999
3.000 0.4123 0.00641 0.00123 -0.0071 0.1970 0.9999
3.250 0.4401 0.00665 0.00135 -0.0071 0.1693 0.9999
3.500 0.4677 0.00692 0.00149 -0.0070 0.1410 0.9999
3.750 0.4954 0.00719 0.00166 -0.0070 0.1156 0.9999
4.000 0.5230 0.00747 0.00184 -0.0069 0.0947 0.9999
4.250 0.5505 0.00774 0.00203 -0.0068 0.0774 0.9999
4.500 0.5781 0.00802 0.00225 -0.0068 0.0633 0.9999
4.750 0.6056 0.00830 0.00249 -0.0067 0.0527 0.9999
5.000 0.6331 0.00860 0.00276 -0.0066 0.0447 0.9999
5.250 0.6605 0.00890 0.00303 -0.0064 0.0392 0.9999
5.500 0.6878 0.00926 0.00343 -0.0063 0.0350 0.9999
5.750 0.7153 0.00951 0.00371 -0.0062 0.0329 0.9999
6.000 0.7425 0.00984 0.00404 -0.0061 0.0301 0.9999
6.250 0.7690 0.01052 0.00480 -0.0059 0.0268 0.9999
6.500 0.7962 0.01078 0.00511 -0.0057 0.0262 0.9999
6.750 0.8232 0.01112 0.00550 -0.0055 0.0252 0.9999
7.000 0.8501 0.01147 0.00589 -0.0054 0.0238 0.9999
7.250 0.8768 0.01183 0.00627 -0.0052 0.0225 0.9999
7.500 0.9028 0.01248 0.00698 -0.0050 0.0210 0.9999
7.750 0.9272 0.01368 0.00834 -0.0047 0.0197 0.9999
8.000 0.9538 0.01399 0.00870 -0.0045 0.0193 0.9999
8.250 0.9802 0.01432 0.00909 -0.0043 0.0185 0.9999
8.500 1.0062 0.01478 0.00962 -0.0041 0.0176 0.9999
8.750 1.0320 0.01526 0.01016 -0.0039 0.0168 0.9999
9.000 1.0574 0.01579 0.01074 -0.0037 0.0160 0.9999
9.250 1.0798 0.01724 0.01232 -0.0033 0.0149 0.9999
9.500 1.1029 0.01842 0.01368 -0.0030 0.0143 0.9999
9.750 1.1285 0.01876 0.01409 -0.0028 0.0138 0.9999
10.000 1.1531 0.01937 0.01481 -0.0025 0.0132 0.9999
10.250 1.1772 0.02005 0.01558 -0.0022 0.0126 0.9999
10.500 1.2012 0.02068 0.01629 -0.0020 0.0120 0.9999
10.750 1.2249 0.02134 0.01701 -0.0017 0.0115 0.9999
11.000 1.2450 0.02281 0.01861 -0.0014 0.0109 0.9999
11.250 1.2596 0.02542 0.02153 -0.0008 0.0103 0.9999
11.500 1.2826 0.02601 0.02224 -0.0005 0.0100 0.9999
11.750 1.3036 0.02698 0.02334 -0.0002 0.0097 0.9999
12.000 1.3225 0.02827 0.02479 0.0001 0.0093 0.9999
12.250 1.3406 0.02959 0.02624 0.0005 0.0089 0.9999
12.500 1.3584 0.03085 0.02762 0.0008 0.0086 0.9999
12.750 1.3753 0.03214 0.02901 0.0010 0.0083 0.9999
13.000 1.3893 0.03378 0.03077 0.0013 0.0081 0.9999
13.250 1.3946 0.03653 0.03370 0.0013 0.0078 0.9999
13.500 1.3841 0.04089 0.03830 0.0008 0.0076 0.9999
13.750 1.3658 0.04841 0.04608 -0.0070 0.0077 0.9999
14.000 1.3418 0.06126 0.05917 -0.0198 0.0078 0.9999
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Polar data table (+)
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