AH 79-100 B AIRFOIL (ah79100b-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: AH 79-100 B AIRFOIL (ah79100b-il) Reynolds number: 100,000 Max Cl/Cd: 68.77 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ah79100b-il-100000-n5.txt Download as CSV file: xf-ah79100b-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AH 79-100 B AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.1454 0.10385 0.09901 -0.0757 0.9507 0.0455
-8.000 -0.1352 0.10107 0.09623 -0.0808 0.9432 0.0458
-7.750 -0.1300 0.09872 0.09391 -0.0848 0.9332 0.0461
-7.500 -0.1095 0.09279 0.08797 -0.0839 0.9319 0.0473
-7.250 -0.0898 0.08904 0.08420 -0.0856 0.9287 0.0487
-7.000 -0.0795 0.08622 0.08138 -0.0868 0.9211 0.0501
-6.750 -0.0657 0.08307 0.07823 -0.0897 0.9152 0.0514
-6.500 -0.0560 0.08022 0.07539 -0.0919 0.9071 0.0521
-6.250 -0.0384 0.07650 0.07165 -0.0967 0.9012 0.0535
-5.750 0.0048 0.06391 0.05893 -0.1172 0.8858 0.0352
-5.500 0.0216 0.06025 0.05524 -0.1207 0.8789 0.0347
-5.250 0.0603 0.05283 0.04764 -0.1347 0.8732 0.0351
-5.000 0.1132 0.04454 0.03887 -0.1503 0.8703 0.0356
-4.750 0.1477 0.03938 0.03333 -0.1575 0.8642 0.0355
-4.500 0.1835 0.03591 0.02968 -0.1629 0.8601 0.0373
-4.250 0.2275 0.03304 0.02647 -0.1689 0.8575 0.0408
-4.000 0.2821 0.02875 0.02144 -0.1769 0.8558 0.0419
-3.750 0.3232 0.02649 0.01849 -0.1808 0.8512 0.0453
-3.500 0.3607 0.02461 0.01611 -0.1834 0.8463 0.0465
-3.250 0.3997 0.02292 0.01418 -0.1862 0.8429 0.0482
-3.000 0.4387 0.02186 0.01292 -0.1886 0.8402 0.0520
-2.750 0.4653 0.02127 0.01217 -0.1886 0.8336 0.0553
-2.500 0.4990 0.02047 0.01119 -0.1897 0.8294 0.0576
-2.250 0.5355 0.01959 0.01027 -0.1915 0.8262 0.0611
-2.000 0.5639 0.01926 0.00994 -0.1919 0.8205 0.0677
-1.750 0.5957 0.01885 0.00947 -0.1929 0.8155 0.0761
-1.500 0.6323 0.01836 0.00893 -0.1947 0.8118 0.0913
-1.250 0.6641 0.01796 0.00864 -0.1958 0.8060 0.1314
-1.000 0.7005 0.01707 0.00869 -0.1984 0.8005 0.4243
-0.750 0.7330 0.01688 0.00870 -0.1990 0.7962 0.5605
-0.500 0.7556 0.01697 0.00894 -0.1977 0.7893 0.6378
-0.250 0.7796 0.01692 0.00901 -0.1965 0.7839 0.7086
0.000 0.8063 0.01681 0.00891 -0.1957 0.7795 0.7572
0.250 0.8260 0.01691 0.00904 -0.1940 0.7719 0.7913
0.500 0.8513 0.01671 0.00883 -0.1930 0.7666 0.8341
0.750 0.8639 0.01658 0.00877 -0.1896 0.7585 0.9720
1.000 0.8988 0.01664 0.00868 -0.1911 0.7528 1.0000
1.250 0.9287 0.01684 0.00878 -0.1917 0.7466 1.0000
1.500 0.9571 0.01705 0.00891 -0.1920 0.7400 1.0000
1.750 0.9907 0.01711 0.00886 -0.1931 0.7354 1.0000
2.000 1.0140 0.01746 0.00920 -0.1925 0.7272 1.0000
2.250 1.0455 0.01753 0.00920 -0.1932 0.7215 1.0000
2.500 1.0699 0.01781 0.00948 -0.1926 0.7133 1.0000
2.750 1.1004 0.01788 0.00949 -0.1931 0.7068 1.0000
3.000 1.1244 0.01816 0.00981 -0.1924 0.6982 1.0000
3.250 1.1543 0.01826 0.00988 -0.1928 0.6919 1.0000
3.500 1.1778 0.01858 0.01025 -0.1921 0.6835 1.0000
3.750 1.2070 0.01870 0.01036 -0.1923 0.6769 1.0000
4.000 1.2299 0.01902 0.01078 -0.1914 0.6679 1.0000
4.250 1.2597 0.01910 0.01087 -0.1917 0.6611 1.0000
4.500 1.2815 0.01943 0.01128 -0.1906 0.6508 1.0000
4.750 1.3068 0.01963 0.01152 -0.1901 0.6413 1.0000
5.000 1.3339 0.01977 0.01173 -0.1898 0.6322 1.0000
5.250 1.3558 0.02011 0.01217 -0.1888 0.6215 1.0000
5.500 1.3797 0.02035 0.01249 -0.1880 0.6108 1.0000
5.750 1.4043 0.02054 0.01277 -0.1873 0.5993 1.0000
6.000 1.4283 0.02077 0.01308 -0.1865 0.5876 1.0000
6.250 1.4499 0.02112 0.01353 -0.1854 0.5753 1.0000
6.500 1.4717 0.02144 0.01395 -0.1842 0.5621 1.0000
6.750 1.4924 0.02174 0.01433 -0.1828 0.5455 1.0000
7.000 1.5115 0.02206 0.01467 -0.1810 0.5262 1.0000
7.250 1.5271 0.02250 0.01514 -0.1787 0.5040 1.0000
7.500 1.5415 0.02302 0.01564 -0.1762 0.4798 1.0000
7.750 1.5543 0.02363 0.01625 -0.1736 0.4558 1.0000
8.000 1.5655 0.02432 0.01698 -0.1707 0.4338 1.0000
8.250 1.5756 0.02510 0.01780 -0.1677 0.4121 1.0000
8.500 1.5832 0.02602 0.01871 -0.1644 0.3882 1.0000
8.750 1.5867 0.02717 0.01981 -0.1607 0.3592 1.0000
9.000 1.5864 0.02860 0.02113 -0.1567 0.3243 1.0000
9.250 1.5837 0.03033 0.02273 -0.1526 0.2859 1.0000
9.500 1.5773 0.03250 0.02468 -0.1486 0.2454 1.0000
9.750 1.5692 0.03506 0.02702 -0.1448 0.2034 1.0000
10.000 1.5598 0.03799 0.02969 -0.1413 0.1595 1.0000
10.250 1.5501 0.04120 0.03264 -0.1383 0.1203 1.0000
10.500 1.5421 0.04446 0.03571 -0.1356 0.0931 1.0000
10.750 1.5368 0.04761 0.03877 -0.1335 0.0749 1.0000
11.000 1.5331 0.05071 0.04187 -0.1316 0.0635 1.0000
11.250 1.5297 0.05388 0.04511 -0.1299 0.0555 1.0000
11.500 1.5251 0.05730 0.04859 -0.1284 0.0487 1.0000
11.750 1.5224 0.06064 0.05209 -0.1271 0.0435 1.0000
12.000 1.5182 0.06427 0.05585 -0.1260 0.0393 1.0000
12.250 1.5125 0.06820 0.05988 -0.1250 0.0362 1.0000
12.500 1.5109 0.07171 0.06361 -0.1243 0.0328 1.0000
12.750 1.5085 0.07542 0.06744 -0.1238 0.0302 1.0000
13.000 1.5046 0.07942 0.07153 -0.1235 0.0283 1.0000
13.250 1.5042 0.08301 0.07531 -0.1231 0.0264 1.0000
13.500 1.5033 0.08670 0.07916 -0.1229 0.0250 1.0000
13.750 1.5028 0.09039 0.08297 -0.1228 0.0238 1.0000
14.000 1.5027 0.09404 0.08673 -0.1227 0.0230 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AH 79-100 B AIRFOIL (ah79100b-il)