EPPLER 678 AIRFOIL (e678-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 678 AIRFOIL (e678-il) Reynolds number: 100,000 Max Cl/Cd: 45.69 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e678-il-100000-n5.txt Download as CSV file: xf-e678-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 678 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1963 0.10029 0.09520 -0.0941 0.9388 0.0293
-9.000 -0.1905 0.09510 0.09002 -0.0984 0.9363 0.0284
-8.500 -0.2182 0.08008 0.07509 -0.1067 0.9233 0.0258
-8.250 -0.2304 0.07564 0.07069 -0.1077 0.9141 0.0255
-8.000 -0.2408 0.06892 0.06391 -0.1138 0.9082 0.0252
-7.750 -0.2618 0.06449 0.05941 -0.1145 0.8959 0.0249
-7.500 -0.2750 0.05967 0.05441 -0.1155 0.8848 0.0247
-7.250 -0.2711 0.05371 0.04807 -0.1194 0.8789 0.0244
-7.000 -0.2740 0.04988 0.04393 -0.1186 0.8676 0.0242
-6.750 -0.2556 0.04477 0.03823 -0.1216 0.8637 0.0242
-6.500 -0.2483 0.04176 0.03482 -0.1205 0.8531 0.0243
-6.250 -0.2213 0.03822 0.03066 -0.1226 0.8496 0.0250
-6.000 -0.2020 0.03580 0.02762 -0.1224 0.8423 0.0260
-5.750 -0.1780 0.03436 0.02609 -0.1230 0.8368 0.0272
-5.500 -0.1466 0.03262 0.02407 -0.1244 0.8342 0.0282
-5.250 -0.1277 0.03142 0.02263 -0.1233 0.8262 0.0289
-5.000 -0.0997 0.03001 0.02098 -0.1237 0.8219 0.0300
-4.750 -0.0677 0.02867 0.01939 -0.1246 0.8193 0.0314
-4.500 -0.0492 0.02801 0.01872 -0.1236 0.8114 0.0329
-4.250 -0.0205 0.02725 0.01788 -0.1242 0.8075 0.0364
-4.000 0.0117 0.02632 0.01690 -0.1254 0.8052 0.0408
-3.750 0.0341 0.02588 0.01637 -0.1250 0.7988 0.0456
-3.500 0.0635 0.02518 0.01561 -0.1258 0.7948 0.0566
-3.250 0.0969 0.02430 0.01487 -0.1275 0.7924 0.0842
-3.000 0.1317 0.02333 0.01427 -0.1295 0.7907 0.1565
-2.750 0.1554 0.02285 0.01426 -0.1301 0.7844 0.2620
-2.500 0.1865 0.02209 0.01412 -0.1315 0.7815 0.4223
-2.250 0.2132 0.02169 0.01417 -0.1311 0.7791 0.5525
-2.000 0.2374 0.02162 0.01427 -0.1294 0.7772 0.6363
-1.750 0.2548 0.02208 0.01469 -0.1273 0.7715 0.6811
-1.500 0.2787 0.02227 0.01477 -0.1262 0.7678 0.7151
-1.250 0.3056 0.02229 0.01466 -0.1255 0.7654 0.7408
-1.000 0.3322 0.02224 0.01451 -0.1247 0.7636 0.7620
-0.750 0.3519 0.02251 0.01469 -0.1229 0.7598 0.7806
-0.500 0.3691 0.02283 0.01495 -0.1209 0.7553 0.7976
-0.250 0.3930 0.02281 0.01484 -0.1198 0.7526 0.8125
0.000 0.4206 0.02264 0.01457 -0.1194 0.7506 0.8250
0.250 0.4512 0.02243 0.01425 -0.1198 0.7491 0.8365
0.750 0.4884 0.02302 0.01474 -0.1169 0.7405 0.8570
1.000 0.5171 0.02297 0.01461 -0.1172 0.7379 0.8662
1.250 0.5465 0.02280 0.01437 -0.1175 0.7360 0.8744
1.500 0.5797 0.02259 0.01409 -0.1185 0.7344 0.8821
1.750 0.6139 0.02230 0.01374 -0.1197 0.7331 0.8894
2.000 0.6187 0.02327 0.01474 -0.1165 0.7254 0.8986
2.250 0.6465 0.02329 0.01473 -0.1167 0.7224 0.9067
2.500 0.6785 0.02309 0.01450 -0.1175 0.7202 0.9143
2.750 0.7145 0.02281 0.01420 -0.1191 0.7185 0.9223
3.000 0.7528 0.02243 0.01379 -0.1209 0.7170 0.9300
3.250 0.7491 0.02369 0.01514 -0.1164 0.7077 0.9445
3.500 0.7845 0.02353 0.01500 -0.1181 0.7048 0.9560
3.750 0.8272 0.02322 0.01468 -0.1210 0.7026 0.9700
4.000 0.8734 0.02280 0.01428 -0.1245 0.7007 1.0000
4.500 0.9138 0.02420 0.01577 -0.1238 0.6880 1.0000
4.750 0.9591 0.02389 0.01549 -0.1271 0.6854 1.0000
5.000 1.0077 0.02354 0.01515 -0.1310 0.6829 1.0000
5.250 1.0016 0.02512 0.01683 -0.1264 0.6724 1.0000
5.500 1.0474 0.02473 0.01650 -0.1297 0.6691 1.0000
5.750 1.0556 0.02577 0.01762 -0.1273 0.6604 1.0000
6.000 1.0883 0.02581 0.01773 -0.1284 0.6547 1.0000
6.250 1.1394 0.02518 0.01716 -0.1324 0.6511 1.0000
6.500 1.1303 0.02672 0.01882 -0.1272 0.6398 1.0000
6.750 1.1784 0.02605 0.01821 -0.1305 0.6354 1.0000
7.000 1.1713 0.02749 0.01979 -0.1256 0.6238 1.0000
7.250 1.2190 0.02668 0.01904 -0.1286 0.6186 1.0000
7.500 1.2118 0.02808 0.02057 -0.1238 0.6062 1.0000
7.750 1.2164 0.02908 0.02168 -0.1208 0.5948 1.0000
8.000 1.2556 0.02835 0.02105 -0.1222 0.5874 1.0000
8.250 1.2521 0.02971 0.02253 -0.1182 0.5738 1.0000
8.500 1.2544 0.03089 0.02383 -0.1150 0.5599 1.0000
8.750 1.2610 0.03190 0.02495 -0.1125 0.5455 1.0000
9.000 1.2715 0.03272 0.02587 -0.1105 0.5303 1.0000
9.250 1.2864 0.03324 0.02649 -0.1089 0.5139 1.0000
9.500 1.3063 0.03342 0.02672 -0.1078 0.4961 1.0000
9.750 1.3199 0.03408 0.02742 -0.1062 0.4748 1.0000
10.000 1.3431 0.03404 0.02733 -0.1053 0.4510 1.0000
10.250 1.3577 0.03471 0.02795 -0.1038 0.4250 1.0000
10.500 1.3696 0.03563 0.02878 -0.1020 0.3985 1.0000
10.750 1.3776 0.03689 0.02995 -0.0999 0.3723 1.0000
11.000 1.3818 0.03855 0.03155 -0.0977 0.3466 1.0000
11.250 1.3849 0.04035 0.03330 -0.0956 0.3221 1.0000
11.500 1.3863 0.04239 0.03527 -0.0934 0.2975 1.0000
11.750 1.3867 0.04461 0.03742 -0.0915 0.2731 1.0000
12.000 1.3852 0.04708 0.03979 -0.0896 0.2488 1.0000
12.250 1.3839 0.04969 0.04233 -0.0879 0.2238 1.0000
12.500 1.3813 0.05253 0.04506 -0.0864 0.1990 1.0000
12.750 1.3779 0.05555 0.04796 -0.0851 0.1748 1.0000
13.000 1.3758 0.05855 0.05087 -0.0840 0.1508 1.0000
13.250 1.3737 0.06165 0.05389 -0.0830 0.1302 1.0000
13.500 1.3741 0.06456 0.05679 -0.0823 0.1126 1.0000
13.750 1.3742 0.06756 0.05979 -0.0816 0.0975 1.0000
14.000 1.3733 0.07075 0.06298 -0.0810 0.0833 1.0000
14.250 1.3709 0.07418 0.06640 -0.0805 0.0700 1.0000
14.750 1.3639 0.08161 0.07384 -0.0800 0.0492 1.0000
15.000 1.3606 0.08542 0.07772 -0.0799 0.0423 1.0000
15.250 1.3563 0.08945 0.08183 -0.0800 0.0375 1.0000
15.500 1.3516 0.09362 0.08611 -0.0804 0.0342 1.0000
15.750 1.3467 0.09792 0.09051 -0.0809 0.0320 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 678 AIRFOIL (e678-il)