AH 79-100 C AIRFOIL (ah79100c-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: AH 79-100 C AIRFOIL (ah79100c-il) Reynolds number: 100,000 Max Cl/Cd: 61.47 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ah79100c-il-100000-n5.txt Download as CSV file: xf-ah79100c-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AH 79-100 C 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.1481 0.10415 0.09920 -0.0794 0.9356 0.0473
-7.500 -0.1434 0.10137 0.09644 -0.0827 0.9298 0.0475
-7.250 -0.1437 0.09908 0.09418 -0.0827 0.9218 0.0475
-7.000 -0.1339 0.09559 0.09070 -0.0860 0.9166 0.0476
-6.250 -0.1030 0.08351 0.07863 -0.0900 0.9018 0.0397
-6.000 -0.1013 0.08113 0.07625 -0.0888 0.8932 0.0366
-5.750 -0.0844 0.07700 0.07210 -0.0934 0.8888 0.0360
-5.500 -0.0768 0.07364 0.06874 -0.0961 0.8811 0.0359
-5.250 -0.0585 0.06912 0.06418 -0.1015 0.8758 0.0350
-5.000 -0.0269 0.06290 0.05786 -0.1112 0.8729 0.0339
-4.750 -0.0104 0.05819 0.05306 -0.1164 0.8644 0.0336
-4.500 0.0326 0.05140 0.04604 -0.1279 0.8612 0.0359
-4.250 0.1043 0.03909 0.03280 -0.1480 0.8604 0.0394
-4.000 0.1401 0.03738 0.03095 -0.1513 0.8587 0.0425
-3.750 0.1932 0.03251 0.02520 -0.1593 0.8580 0.0456
-3.500 0.2236 0.03027 0.02223 -0.1613 0.8522 0.0475
-3.250 0.2588 0.02843 0.02001 -0.1636 0.8491 0.0489
-3.000 0.2939 0.02729 0.01865 -0.1655 0.8468 0.0514
-2.750 0.3306 0.02625 0.01734 -0.1674 0.8450 0.0554
-2.500 0.3685 0.02511 0.01587 -0.1693 0.8435 0.0583
-2.250 0.3905 0.02470 0.01540 -0.1687 0.8372 0.0609
-2.000 0.4209 0.02417 0.01484 -0.1694 0.8336 0.0659
-1.750 0.4553 0.02354 0.01410 -0.1707 0.8312 0.0742
-1.500 0.4916 0.02294 0.01343 -0.1723 0.8294 0.0859
-1.250 0.5288 0.02231 0.01286 -0.1742 0.8279 0.1054
-1.000 0.5443 0.02255 0.01316 -0.1723 0.8193 0.1241
-0.750 0.5786 0.02209 0.01282 -0.1736 0.8166 0.1690
-0.500 0.6155 0.02158 0.01248 -0.1755 0.8146 0.2442
-0.250 0.6543 0.02103 0.01217 -0.1776 0.8130 0.3466
0.000 0.6693 0.02131 0.01273 -0.1757 0.8043 0.4446
0.250 0.6979 0.02052 0.01267 -0.1753 0.8014 0.7008
0.500 0.7240 0.01977 0.01210 -0.1740 0.7986 1.0000
1.000 0.7744 0.02022 0.01227 -0.1732 0.7865 1.0000
1.250 0.8127 0.02001 0.01192 -0.1750 0.7840 1.0000
1.750 0.8640 0.02037 0.01211 -0.1743 0.7715 1.0000
2.000 0.9034 0.02010 0.01176 -0.1763 0.7686 1.0000
2.250 0.9203 0.02055 0.01219 -0.1744 0.7602 1.0000
2.500 0.9543 0.02045 0.01203 -0.1754 0.7556 1.0000
2.750 0.9958 0.02014 0.01167 -0.1778 0.7525 1.0000
3.000 1.0087 0.02065 0.01221 -0.1751 0.7426 1.0000
3.250 1.0483 0.02036 0.01189 -0.1771 0.7383 1.0000
3.500 1.0652 0.02075 0.01230 -0.1751 0.7291 1.0000
3.750 1.1011 0.02058 0.01212 -0.1764 0.7237 1.0000
4.000 1.1208 0.02088 0.01247 -0.1749 0.7148 1.0000
4.250 1.1566 0.02066 0.01227 -0.1762 0.7080 1.0000
4.500 1.1737 0.02099 0.01264 -0.1741 0.6976 1.0000
4.750 1.2137 0.02068 0.01238 -0.1761 0.6911 1.0000
5.000 1.2265 0.02112 0.01289 -0.1733 0.6796 1.0000
5.250 1.2483 0.02131 0.01313 -0.1721 0.6690 1.0000
5.500 1.2790 0.02122 0.01309 -0.1724 0.6591 1.0000
5.750 1.2989 0.02144 0.01339 -0.1708 0.6468 1.0000
6.000 1.3167 0.02173 0.01376 -0.1689 0.6336 1.0000
6.250 1.3366 0.02195 0.01405 -0.1673 0.6195 1.0000
6.500 1.3570 0.02216 0.01433 -0.1659 0.6045 1.0000
6.750 1.3771 0.02242 0.01467 -0.1644 0.5892 1.0000
7.000 1.3965 0.02272 0.01503 -0.1628 0.5733 1.0000
7.250 1.4161 0.02304 0.01540 -0.1612 0.5566 1.0000
7.500 1.4347 0.02338 0.01578 -0.1595 0.5377 1.0000
7.750 1.4495 0.02389 0.01635 -0.1571 0.5165 1.0000
8.000 1.4659 0.02439 0.01682 -0.1551 0.4941 1.0000
8.250 1.4793 0.02505 0.01750 -0.1527 0.4719 1.0000
8.500 1.4922 0.02579 0.01821 -0.1502 0.4494 1.0000
8.750 1.5028 0.02665 0.01910 -0.1475 0.4264 1.0000
9.000 1.5093 0.02772 0.02009 -0.1443 0.3992 1.0000
9.250 1.5108 0.02907 0.02131 -0.1405 0.3670 1.0000
9.500 1.5086 0.03071 0.02278 -0.1365 0.3308 1.0000
9.750 1.5037 0.03266 0.02449 -0.1324 0.2919 1.0000
10.000 1.4992 0.03476 0.02637 -0.1288 0.2535 1.0000
10.250 1.4964 0.03693 0.02836 -0.1256 0.2187 1.0000
10.500 1.4958 0.03909 0.03041 -0.1229 0.1878 1.0000
10.750 1.4945 0.04141 0.03260 -0.1203 0.1562 1.0000
11.000 1.4926 0.04394 0.03499 -0.1179 0.1232 1.0000
11.250 1.4870 0.04695 0.03777 -0.1154 0.0875 1.0000
11.500 1.4804 0.05023 0.04083 -0.1130 0.0585 1.0000
11.750 1.4753 0.05348 0.04398 -0.1110 0.0425 1.0000
12.000 1.4730 0.05655 0.04709 -0.1092 0.0348 1.0000
12.250 1.4712 0.05964 0.05028 -0.1077 0.0311 1.0000
12.500 1.4704 0.06271 0.05354 -0.1064 0.0287 1.0000
12.750 1.4681 0.06604 0.05700 -0.1053 0.0270 1.0000
13.000 1.4635 0.06973 0.06081 -0.1044 0.0259 1.0000
13.250 1.4611 0.07327 0.06456 -0.1037 0.0250 1.0000
13.500 1.4581 0.07697 0.06846 -0.1031 0.0242 1.0000
13.750 1.4547 0.08082 0.07249 -0.1027 0.0236 1.0000
14.000 1.4512 0.08474 0.07658 -0.1026 0.0230 1.0000
14.250 1.4480 0.08869 0.08070 -0.1025 0.0224 1.0000
14.500 1.4453 0.09264 0.08479 -0.1027 0.0219 1.0000
14.750 1.4427 0.09659 0.08886 -0.1029 0.0213 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AH 79-100 C AIRFOIL (ah79100c-il)