EPPLER 422 AIRFOIL (e422-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 422 AIRFOIL (e422-il) Reynolds number: 1,000,000 Max Cl/Cd: 139.81 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e422-il-1000000-n5.txt Download as CSV file: xf-e422-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 422 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.2200 0.05220 0.04841 -0.1047 0.5600 0.0059
-9.500 -0.2567 0.04228 0.03839 -0.1143 0.5582 0.0058
-9.250 -0.2789 0.02794 0.02360 -0.1280 0.5560 0.0057
-9.000 -0.2642 0.02298 0.01826 -0.1312 0.5498 0.0057
-8.750 -0.2445 0.02105 0.01610 -0.1316 0.5437 0.0058
-8.500 -0.2230 0.01967 0.01452 -0.1316 0.5370 0.0058
-8.250 -0.2010 0.01850 0.01314 -0.1313 0.5301 0.0058
-8.000 -0.1774 0.01757 0.01207 -0.1311 0.5249 0.0058
-7.750 -0.1534 0.01673 0.01107 -0.1307 0.5192 0.0058
-7.500 -0.1290 0.01603 0.01023 -0.1303 0.5136 0.0058
-7.250 -0.1040 0.01537 0.00944 -0.1299 0.5091 0.0059
-7.000 -0.0786 0.01479 0.00873 -0.1296 0.5040 0.0059
-6.750 -0.0530 0.01428 0.00810 -0.1292 0.4984 0.0059
-6.500 -0.0271 0.01382 0.00754 -0.1288 0.4935 0.0059
-6.250 -0.0013 0.01324 0.00685 -0.1284 0.4892 0.0060
-6.000 0.0246 0.01271 0.00620 -0.1280 0.4847 0.0061
-5.750 0.0510 0.01229 0.00568 -0.1276 0.4803 0.0063
-5.500 0.0778 0.01197 0.00528 -0.1272 0.4765 0.0064
-5.250 0.1051 0.01169 0.00495 -0.1269 0.4732 0.0066
-5.000 0.1324 0.01145 0.00464 -0.1266 0.4692 0.0068
-4.750 0.1597 0.01124 0.00437 -0.1263 0.4649 0.0070
-4.250 0.2146 0.01087 0.00388 -0.1258 0.4577 0.0075
-4.000 0.2423 0.01069 0.00367 -0.1255 0.4548 0.0079
-3.750 0.2700 0.01055 0.00348 -0.1253 0.4516 0.0081
-3.500 0.2975 0.01038 0.00326 -0.1250 0.4484 0.0087
-3.250 0.3250 0.01027 0.00311 -0.1247 0.4448 0.0093
-3.000 0.3527 0.01017 0.00297 -0.1244 0.4419 0.0101
-2.750 0.3806 0.01005 0.00283 -0.1242 0.4393 0.0111
-2.500 0.4084 0.00994 0.00271 -0.1240 0.4365 0.0129
-2.250 0.4362 0.00985 0.00261 -0.1238 0.4337 0.0158
-2.000 0.4638 0.00976 0.00251 -0.1235 0.4310 0.0206
-1.750 0.4911 0.00966 0.00242 -0.1233 0.4282 0.0311
-1.500 0.5184 0.00953 0.00235 -0.1230 0.4256 0.0481
-1.250 0.5458 0.00936 0.00228 -0.1228 0.4236 0.0768
-1.000 0.5734 0.00925 0.00225 -0.1226 0.4211 0.1021
-0.750 0.6009 0.00921 0.00225 -0.1224 0.4186 0.1213
-0.500 0.6285 0.00918 0.00226 -0.1222 0.4162 0.1372
-0.250 0.6560 0.00918 0.00227 -0.1220 0.4139 0.1487
0.000 0.6833 0.00921 0.00229 -0.1218 0.4113 0.1582
0.250 0.7108 0.00923 0.00231 -0.1215 0.4090 0.1660
0.500 0.7386 0.00924 0.00233 -0.1214 0.4072 0.1728
0.750 0.7663 0.00925 0.00236 -0.1212 0.4051 0.1810
1.000 0.7939 0.00929 0.00238 -0.1210 0.4029 0.1867
1.250 0.8212 0.00931 0.00242 -0.1208 0.4008 0.1962
1.500 0.8485 0.00935 0.00247 -0.1206 0.3986 0.2036
1.750 0.8755 0.00940 0.00252 -0.1203 0.3963 0.2128
2.000 0.9023 0.00947 0.00259 -0.1200 0.3937 0.2223
2.250 0.9295 0.00950 0.00265 -0.1198 0.3919 0.2339
2.500 0.9568 0.00953 0.00271 -0.1196 0.3902 0.2457
2.750 0.9840 0.00956 0.00278 -0.1194 0.3883 0.2590
3.000 1.0109 0.00960 0.00286 -0.1192 0.3863 0.2765
3.250 1.0376 0.00964 0.00295 -0.1190 0.3842 0.2978
3.500 1.0640 0.00970 0.00304 -0.1187 0.3819 0.3204
3.750 1.0900 0.00975 0.00315 -0.1183 0.3794 0.3528
4.000 1.1158 0.00981 0.00328 -0.1180 0.3769 0.3896
4.250 1.1422 0.00980 0.00339 -0.1177 0.3755 0.4370
4.500 1.1616 0.00920 0.00360 -0.1162 0.3740 0.7957
4.750 1.2116 0.00908 0.00379 -0.1210 0.3715 1.0000
5.000 1.2368 0.00920 0.00389 -0.1205 0.3691 1.0000
5.250 1.2618 0.00932 0.00401 -0.1200 0.3667 1.0000
5.500 1.2864 0.00946 0.00413 -0.1194 0.3643 1.0000
5.750 1.3106 0.00962 0.00427 -0.1187 0.3617 1.0000
6.000 1.3359 0.00973 0.00438 -0.1183 0.3598 1.0000
6.250 1.3610 0.00984 0.00452 -0.1178 0.3575 1.0000
6.500 1.3857 0.00997 0.00465 -0.1173 0.3544 1.0000
6.750 1.4096 0.01013 0.00479 -0.1166 0.3507 1.0000
7.000 1.4326 0.01032 0.00497 -0.1158 0.3468 1.0000
7.250 1.4569 0.01046 0.00512 -0.1152 0.3443 1.0000
7.500 1.4810 0.01060 0.00527 -0.1146 0.3413 1.0000
7.750 1.5044 0.01076 0.00544 -0.1139 0.3379 1.0000
8.000 1.5268 0.01095 0.00563 -0.1131 0.3344 1.0000
8.250 1.5485 0.01117 0.00584 -0.1121 0.3307 1.0000
8.500 1.5715 0.01132 0.00602 -0.1113 0.3272 1.0000
8.750 1.5925 0.01151 0.00622 -0.1102 0.3229 1.0000
9.000 1.6113 0.01175 0.00645 -0.1087 0.3185 1.0000
9.250 1.6303 0.01197 0.00669 -0.1073 0.3150 1.0000
9.500 1.6503 0.01218 0.00692 -0.1060 0.3108 1.0000
9.750 1.6684 0.01245 0.00719 -0.1045 0.3057 1.0000
10.000 1.6853 0.01276 0.00751 -0.1028 0.3009 1.0000
10.250 1.7040 0.01303 0.00779 -0.1015 0.2956 1.0000
10.500 1.7194 0.01341 0.00817 -0.0996 0.2890 1.0000
10.750 1.7358 0.01377 0.00854 -0.0980 0.2829 1.0000
11.000 1.7496 0.01424 0.00899 -0.0960 0.2751 1.0000
11.250 1.7638 0.01470 0.00946 -0.0941 0.2684 1.0000
11.500 1.7748 0.01531 0.01006 -0.0918 0.2598 1.0000
11.750 1.7854 0.01596 0.01070 -0.0896 0.2505 1.0000
12.000 1.7921 0.01681 0.01152 -0.0870 0.2402 1.0000
12.250 1.7969 0.01782 0.01251 -0.0844 0.2303 1.0000
12.500 1.8023 0.01886 0.01355 -0.0820 0.2205 1.0000
12.750 1.8035 0.02024 0.01492 -0.0794 0.2101 1.0000
13.000 1.8036 0.02182 0.01649 -0.0770 0.2012 1.0000
13.250 1.8036 0.02354 0.01822 -0.0749 0.1923 1.0000
13.500 1.8022 0.02551 0.02020 -0.0731 0.1842 1.0000
14.000 1.7940 0.03033 0.02505 -0.0698 0.1684 1.0000
14.250 1.7860 0.03328 0.02802 -0.0684 0.1614 1.0000
14.500 1.7842 0.03577 0.03057 -0.0675 0.1568 1.0000
14.750 1.7723 0.03930 0.03412 -0.0663 0.1499 1.0000
15.000 1.7658 0.04242 0.03730 -0.0656 0.1449 1.0000
15.250 1.7531 0.04628 0.04120 -0.0648 0.1390 1.0000
15.500 1.7433 0.04994 0.04491 -0.0643 0.1345 1.0000
15.750 1.7326 0.05380 0.04881 -0.0639 0.1291 1.0000
16.000 1.7181 0.05816 0.05322 -0.0636 0.1240 1.0000
16.250 1.7103 0.06189 0.05700 -0.0636 0.1200 1.0000
16.500 1.6982 0.06619 0.06134 -0.0636 0.1152 1.0000
16.750 1.6876 0.07038 0.06559 -0.0637 0.1109 1.0000
17.000 1.6804 0.07422 0.06947 -0.0640 0.1067 1.0000
17.250 1.6673 0.07893 0.07420 -0.0644 0.1018 1.0000
17.500 1.6615 0.08273 0.07807 -0.0649 0.0986 1.0000
17.750 1.6546 0.08673 0.08209 -0.0654 0.0942 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 422 AIRFOIL (e422-il)