AH 79-100 A AIRFOIL (ah79100a-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: AH 79-100 A AIRFOIL (ah79100a-il) Reynolds number: 500,000 Max Cl/Cd: 116.19 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ah79100a-il-500000.txt Download as CSV file: xf-ah79100a-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: AH 79-100 A 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.2879 0.08606 0.08396 -0.0406 1.0000 0.0255
-9.500 -0.2998 0.08181 0.07976 -0.0410 1.0000 0.0264
-9.250 -0.3131 0.07923 0.07723 -0.0393 1.0000 0.0266
-8.500 -0.4848 0.04207 0.03951 -0.0837 0.9890 0.0157
-8.250 -0.4642 0.03294 0.02979 -0.0959 0.9822 0.0153
-8.000 -0.4335 0.02686 0.02305 -0.1033 0.9796 0.0154
-7.750 -0.4067 0.02354 0.01924 -0.1060 0.9742 0.0157
-7.500 -0.3764 0.01956 0.01479 -0.1092 0.9715 0.0162
-7.250 -0.3412 0.01853 0.01368 -0.1116 0.9699 0.0173
-7.000 -0.3052 0.01747 0.01248 -0.1139 0.9686 0.0184
-6.750 -0.2747 0.01625 0.01106 -0.1148 0.9643 0.0195
-6.500 -0.2422 0.01542 0.01004 -0.1159 0.9601 0.0205
-6.250 -0.2099 0.01363 0.00814 -0.1175 0.9566 0.0226
-6.000 -0.1778 0.01309 0.00753 -0.1185 0.9527 0.0248
-5.750 -0.1497 0.01266 0.00702 -0.1185 0.9465 0.0265
-5.500 -0.1202 0.01177 0.00608 -0.1192 0.9421 0.0298
-5.250 -0.0913 0.01145 0.00572 -0.1194 0.9371 0.0326
-5.000 -0.0636 0.01124 0.00545 -0.1193 0.9309 0.0347
-4.750 -0.0355 0.01041 0.00455 -0.1195 0.9259 0.0387
-4.500 -0.0085 0.01011 0.00421 -0.1192 0.9193 0.0426
-4.250 0.0191 0.00968 0.00371 -0.1191 0.9131 0.0477
-4.000 0.0468 0.00936 0.00336 -0.1190 0.9069 0.0563
-3.750 0.0739 0.00893 0.00299 -0.1188 0.9001 0.0775
-3.500 0.1018 0.00851 0.00269 -0.1189 0.8946 0.1201
-3.250 0.1282 0.00807 0.00249 -0.1188 0.8873 0.1895
-3.000 0.1557 0.00773 0.00232 -0.1188 0.8814 0.2601
-2.750 0.1827 0.00750 0.00224 -0.1186 0.8741 0.3190
-2.500 0.2102 0.00734 0.00217 -0.1184 0.8671 0.3674
-2.250 0.2375 0.00725 0.00214 -0.1182 0.8594 0.4008
-2.000 0.2652 0.00721 0.00207 -0.1180 0.8524 0.4231
-1.750 0.2928 0.00717 0.00204 -0.1177 0.8449 0.4434
-1.500 0.3204 0.00713 0.00201 -0.1175 0.8375 0.4664
-1.250 0.3480 0.00711 0.00199 -0.1173 0.8294 0.4829
-1.000 0.3756 0.00710 0.00194 -0.1170 0.8210 0.5013
-0.750 0.4027 0.00706 0.00194 -0.1167 0.8111 0.5189
-0.500 0.4304 0.00706 0.00191 -0.1164 0.8026 0.5323
-0.250 0.4578 0.00705 0.00190 -0.1162 0.7934 0.5473
0.000 0.4853 0.00704 0.00190 -0.1159 0.7837 0.5616
0.250 0.5125 0.00703 0.00189 -0.1156 0.7733 0.5752
0.500 0.5397 0.00703 0.00189 -0.1153 0.7616 0.5900
0.750 0.5667 0.00703 0.00190 -0.1149 0.7490 0.6055
1.000 0.5937 0.00704 0.00192 -0.1146 0.7366 0.6211
1.250 0.6207 0.00705 0.00195 -0.1143 0.7241 0.6368
1.500 0.6473 0.00707 0.00199 -0.1139 0.7096 0.6545
1.750 0.6735 0.00709 0.00203 -0.1134 0.6938 0.6762
2.000 0.6996 0.00711 0.00209 -0.1129 0.6779 0.7017
2.250 0.7252 0.00712 0.00217 -0.1123 0.6621 0.7343
2.500 0.7496 0.00710 0.00225 -0.1114 0.6450 0.7845
2.750 0.7791 0.00684 0.00228 -0.1115 0.6244 1.0000
3.000 0.8053 0.00703 0.00236 -0.1111 0.6027 1.0000
3.250 0.8315 0.00723 0.00248 -0.1107 0.5843 1.0000
3.500 0.8577 0.00742 0.00261 -0.1103 0.5673 1.0000
3.750 0.8839 0.00762 0.00275 -0.1100 0.5489 1.0000
4.000 0.9098 0.00783 0.00291 -0.1096 0.5299 1.0000
4.250 0.9352 0.00808 0.00308 -0.1091 0.5097 1.0000
4.500 0.9606 0.00832 0.00326 -0.1086 0.4871 1.0000
4.750 0.9853 0.00861 0.00345 -0.1080 0.4611 1.0000
5.000 1.0100 0.00891 0.00368 -0.1074 0.4354 1.0000
5.250 1.0346 0.00922 0.00391 -0.1068 0.4099 1.0000
5.500 1.0585 0.00959 0.00418 -0.1061 0.3808 1.0000
5.750 1.0806 0.01011 0.00450 -0.1052 0.3375 1.0000
6.000 1.1015 0.01076 0.00490 -0.1041 0.2876 1.0000
6.250 1.1222 0.01144 0.00533 -0.1030 0.2397 1.0000
6.500 1.1430 0.01213 0.00580 -0.1020 0.1977 1.0000
6.750 1.1632 0.01286 0.00632 -0.1009 0.1577 1.0000
7.000 1.1836 0.01357 0.00686 -0.0998 0.1255 1.0000
7.250 1.2039 0.01426 0.00741 -0.0987 0.0984 1.0000
7.500 1.2237 0.01499 0.00800 -0.0975 0.0729 1.0000
7.750 1.2420 0.01584 0.00869 -0.0961 0.0480 1.0000
8.000 1.2584 0.01688 0.00959 -0.0943 0.0287 1.0000
8.250 1.2760 0.01774 0.01047 -0.0927 0.0229 1.0000
8.500 1.2936 0.01857 0.01132 -0.0911 0.0201 1.0000
8.750 1.3090 0.01952 0.01234 -0.0892 0.0183 1.0000
9.000 1.3262 0.02025 0.01315 -0.0875 0.0171 1.0000
9.250 1.3414 0.02106 0.01402 -0.0856 0.0162 1.0000
9.500 1.3531 0.02195 0.01498 -0.0831 0.0155 1.0000
9.750 1.3609 0.02308 0.01618 -0.0802 0.0149 1.0000
10.000 1.3634 0.02460 0.01781 -0.0766 0.0143 1.0000
10.250 1.3676 0.02613 0.01944 -0.0735 0.0139 1.0000
10.500 1.3769 0.02731 0.02072 -0.0713 0.0137 1.0000
10.750 1.3854 0.02861 0.02213 -0.0690 0.0135 1.0000
11.000 1.3931 0.03005 0.02367 -0.0669 0.0132 1.0000
11.250 1.4007 0.03156 0.02528 -0.0649 0.0128 1.0000
11.500 1.4081 0.03312 0.02695 -0.0630 0.0125 1.0000
11.750 1.4154 0.03470 0.02863 -0.0613 0.0121 1.0000
12.000 1.4216 0.03642 0.03045 -0.0597 0.0117 1.0000
12.250 1.4267 0.03837 0.03250 -0.0581 0.0115 1.0000
12.500 1.4312 0.04042 0.03464 -0.0567 0.0112 1.0000
12.750 1.4342 0.04276 0.03709 -0.0553 0.0110 1.0000
13.000 1.4359 0.04543 0.03988 -0.0540 0.0108 1.0000
13.250 1.4359 0.04857 0.04316 -0.0527 0.0106 1.0000
13.500 1.4337 0.05219 0.04697 -0.0516 0.0105 1.0000
13.750 1.4278 0.05644 0.05145 -0.0506 0.0104 1.0000
14.000 1.4200 0.06069 0.05591 -0.0501 0.0103 1.0000
14.250 1.4128 0.06441 0.05982 -0.0502 0.0103 1.0000
14.500 1.4038 0.06875 0.06436 -0.0507 0.0102 1.0000
14.750 1.3927 0.07357 0.06939 -0.0517 0.0102 1.0000
15.000 1.3805 0.07885 0.07486 -0.0532 0.0102 1.0000
15.250 1.3669 0.08466 0.08087 -0.0553 0.0102 1.0000
15.500 1.3514 0.09113 0.08753 -0.0580 0.0102 1.0000
15.750 1.3345 0.09823 0.09482 -0.0614 0.0102 1.0000
16.000 1.3165 0.10594 0.10272 -0.0654 0.0102 1.0000
16.250 1.2968 0.11441 0.11137 -0.0701 0.0103 1.0000
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