AH 79-100 C AIRFOIL (ah79100c-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: AH 79-100 C AIRFOIL (ah79100c-il) Reynolds number: 500,000 Max Cl/Cd: 147.56 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ah79100c-il-500000.txt Download as CSV file: xf-ah79100c-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.0099 0.08652 0.08406 -0.1127 0.9537 0.0197
-8.000 -0.0128 0.08229 0.07986 -0.1168 0.9444 0.0204
-7.750 0.0013 0.07764 0.07521 -0.1212 0.9424 0.0206
-7.500 0.0218 0.07518 0.07274 -0.1224 0.9416 0.0212
-7.250 0.0229 0.07284 0.07042 -0.1220 0.9331 0.0212
-7.000 0.0421 0.07059 0.06816 -0.1244 0.9309 0.0226
-6.750 0.0615 0.06643 0.06399 -0.1300 0.9289 0.0239
-6.500 0.0726 0.06071 0.05828 -0.1387 0.9187 0.0252
-6.250 0.1117 0.02077 0.01688 -0.1889 0.9067 0.0164
-6.000 0.1511 0.01797 0.01374 -0.1937 0.9051 0.0177
-5.750 0.1927 0.01608 0.01138 -0.1976 0.9038 0.0203
-5.500 0.2334 0.01526 0.01047 -0.2010 0.9024 0.0233
-5.250 0.2760 0.01411 0.00903 -0.2048 0.9011 0.0264
-5.000 0.2979 0.01399 0.00890 -0.2037 0.8938 0.0280
-4.750 0.3345 0.01391 0.00869 -0.2056 0.8906 0.0308
-4.500 0.3737 0.01324 0.00796 -0.2086 0.8880 0.0339
-4.250 0.4029 0.01291 0.00754 -0.2091 0.8823 0.0359
-4.000 0.4345 0.01213 0.00659 -0.2100 0.8771 0.0367
-3.750 0.4716 0.01157 0.00588 -0.2121 0.8735 0.0379
-3.500 0.4998 0.01115 0.00535 -0.2122 0.8674 0.0384
-3.250 0.5302 0.01029 0.00440 -0.2130 0.8618 0.0400
-3.000 0.5662 0.00987 0.00392 -0.2149 0.8578 0.0421
-2.750 0.5910 0.00965 0.00368 -0.2143 0.8509 0.0446
-2.500 0.6219 0.00947 0.00341 -0.2150 0.8456 0.0478
-2.250 0.6530 0.00916 0.00309 -0.2158 0.8403 0.0580
-2.000 0.6796 0.00897 0.00290 -0.2156 0.8338 0.0711
-1.750 0.7117 0.00879 0.00272 -0.2166 0.8288 0.0907
-1.500 0.7378 0.00864 0.00267 -0.2164 0.8223 0.1263
-1.250 0.7662 0.00849 0.00264 -0.2167 0.8163 0.1852
-1.000 0.7960 0.00837 0.00265 -0.2173 0.8109 0.2480
-0.750 0.8214 0.00826 0.00269 -0.2169 0.8042 0.3142
-0.500 0.8516 0.00817 0.00270 -0.2175 0.7986 0.3851
-0.250 0.8770 0.00805 0.00277 -0.2172 0.7921 0.4664
0.000 0.9040 0.00784 0.00288 -0.2172 0.7862 0.6041
0.250 0.9284 0.00757 0.00304 -0.2164 0.7808 0.7921
0.500 0.9449 0.00729 0.00302 -0.2135 0.7739 1.0000
0.750 0.9742 0.00739 0.00301 -0.2138 0.7681 1.0000
1.000 0.9992 0.00748 0.00306 -0.2132 0.7616 1.0000
1.250 1.0261 0.00757 0.00310 -0.2131 0.7556 1.0000
1.500 1.0533 0.00768 0.00315 -0.2130 0.7494 1.0000
1.750 1.0780 0.00777 0.00321 -0.2123 0.7423 1.0000
2.000 1.1060 0.00789 0.00326 -0.2124 0.7360 1.0000
2.250 1.1295 0.00798 0.00336 -0.2115 0.7289 1.0000
2.500 1.1568 0.00810 0.00342 -0.2115 0.7226 1.0000
2.750 1.1801 0.00820 0.00353 -0.2105 0.7148 1.0000
3.000 1.2058 0.00833 0.00360 -0.2101 0.7069 1.0000
3.250 1.2280 0.00843 0.00372 -0.2090 0.6983 1.0000
3.500 1.2528 0.00857 0.00381 -0.2083 0.6899 1.0000
3.750 1.2747 0.00868 0.00394 -0.2071 0.6804 1.0000
4.000 1.2981 0.00881 0.00407 -0.2062 0.6719 1.0000
4.250 1.3209 0.00896 0.00421 -0.2052 0.6630 1.0000
4.500 1.3428 0.00910 0.00436 -0.2040 0.6534 1.0000
4.750 1.3640 0.00926 0.00451 -0.2027 0.6427 1.0000
5.000 1.3814 0.00944 0.00467 -0.2005 0.6287 1.0000
5.250 1.3966 0.00965 0.00485 -0.1978 0.6115 1.0000
5.500 1.4112 0.00991 0.00506 -0.1951 0.5918 1.0000
5.750 1.4251 0.01024 0.00532 -0.1922 0.5704 1.0000
6.000 1.4357 0.01067 0.00564 -0.1888 0.5412 1.0000
6.250 1.4443 0.01124 0.00605 -0.1850 0.5085 1.0000
6.500 1.4525 0.01189 0.00654 -0.1813 0.4727 1.0000
6.750 1.4613 0.01261 0.00708 -0.1778 0.4383 1.0000
7.000 1.4708 0.01333 0.00767 -0.1745 0.4071 1.0000
7.250 1.4818 0.01405 0.00827 -0.1716 0.3789 1.0000
7.500 1.4912 0.01487 0.00894 -0.1684 0.3478 1.0000
7.750 1.4972 0.01590 0.00977 -0.1648 0.3099 1.0000
8.000 1.5063 0.01682 0.01054 -0.1619 0.2796 1.0000
8.250 1.5137 0.01788 0.01144 -0.1587 0.2459 1.0000
8.500 1.5200 0.01904 0.01243 -0.1555 0.2112 1.0000
8.750 1.5233 0.02046 0.01361 -0.1520 0.1710 1.0000
9.000 1.5214 0.02230 0.01512 -0.1478 0.1207 1.0000
9.250 1.5118 0.02480 0.01719 -0.1429 0.0609 1.0000
9.500 1.5040 0.02734 0.01943 -0.1384 0.0222 1.0000
9.750 1.5127 0.02869 0.02083 -0.1361 0.0183 1.0000
10.000 1.5207 0.03012 0.02232 -0.1338 0.0165 1.0000
10.250 1.5290 0.03155 0.02385 -0.1317 0.0156 1.0000
10.500 1.5387 0.03290 0.02530 -0.1298 0.0149 1.0000
10.750 1.5472 0.03438 0.02686 -0.1278 0.0144 1.0000
11.000 1.5547 0.03599 0.02856 -0.1259 0.0139 1.0000
11.250 1.5606 0.03778 0.03044 -0.1239 0.0134 1.0000
11.500 1.5658 0.03970 0.03246 -0.1220 0.0132 1.0000
11.750 1.5682 0.04195 0.03479 -0.1200 0.0128 1.0000
12.000 1.5680 0.04453 0.03747 -0.1180 0.0126 1.0000
12.250 1.5656 0.04744 0.04049 -0.1159 0.0124 1.0000
12.500 1.5629 0.05049 0.04365 -0.1141 0.0122 1.0000
12.750 1.5667 0.05291 0.04617 -0.1128 0.0121 1.0000
13.000 1.5705 0.05535 0.04872 -0.1116 0.0119 1.0000
13.250 1.5732 0.05802 0.05149 -0.1105 0.0117 1.0000
13.500 1.5758 0.06075 0.05434 -0.1095 0.0115 1.0000
13.750 1.5775 0.06366 0.05735 -0.1085 0.0113 1.0000
14.000 1.5790 0.06664 0.06044 -0.1076 0.0112 1.0000
14.250 1.5806 0.06968 0.06358 -0.1069 0.0111 1.0000
14.500 1.5827 0.07270 0.06670 -0.1062 0.0110 1.0000
14.750 1.5851 0.07572 0.06983 -0.1056 0.0109 1.0000
15.000 1.5880 0.07873 0.07295 -0.1051 0.0108 1.0000
15.250 1.5912 0.08171 0.07603 -0.1046 0.0107 1.0000
15.500 1.5945 0.08471 0.07915 -0.1042 0.0106 1.0000
15.750 1.5981 0.08770 0.08226 -0.1039 0.0105 1.0000
16.000 1.6016 0.09075 0.08542 -0.1037 0.0104 1.0000
16.250 1.6048 0.09387 0.08867 -0.1035 0.0104 1.0000
16.500 1.6072 0.09711 0.09205 -0.1035 0.0103 1.0000
16.750 1.6087 0.10053 0.09561 -0.1036 0.0102 1.0000
17.000 1.6089 0.10416 0.09939 -0.1039 0.0101 1.0000
17.250 1.6077 0.10802 0.10337 -0.1046 0.0100 1.0000
17.500 1.6058 0.11200 0.10750 -0.1055 0.0099 1.0000
17.750 1.6033 0.11609 0.11172 -0.1065 0.0098 1.0000
18.000 1.5996 0.12045 0.11623 -0.1078 0.0097 1.0000
18.250 1.5954 0.12489 0.12081 -0.1092 0.0096 1.0000
18.500 1.5883 0.12997 0.12607 -0.1112 0.0096 1.0000
18.750 1.5793 0.13557 0.13187 -0.1138 0.0096 1.0000
19.000 1.5688 0.14158 0.13808 -0.1169 0.0097 1.0000
19.250 1.5545 0.14860 0.14533 -0.1208 0.0098 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AH 79-100 C AIRFOIL (ah79100c-il)