GOE 332 (PFALZ 61) AIRFOIL (goe332-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 332 (PFALZ 61) AIRFOIL (goe332-il) Reynolds number: 100,000 Max Cl/Cd: 59.44 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe332-il-100000-n5.txt Download as CSV file: xf-goe332-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 332 (PFALZ 61) 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.1637 0.10647 0.10158 -0.0603 0.9683 0.0611
-8.750 -0.1557 0.10378 0.09890 -0.0662 0.9560 0.0645
-8.500 -0.1554 0.10163 0.09677 -0.0746 0.9394 0.0652
-8.250 -0.1313 0.09620 0.09134 -0.0751 0.9364 0.0661
-8.000 -0.1119 0.09247 0.08759 -0.0756 0.9278 0.0674
-7.750 -0.0928 0.08900 0.08409 -0.0785 0.9202 0.0691
-7.500 -0.0785 0.08585 0.08091 -0.0815 0.9091 0.0710
-7.250 -0.0656 0.08270 0.07773 -0.0858 0.8972 0.0736
-6.750 -0.0308 0.07498 0.06991 -0.0992 0.8767 0.0779
-6.500 -0.0121 0.07216 0.06705 -0.0994 0.8682 0.0800
-6.250 0.0090 0.06898 0.06380 -0.1031 0.8597 0.0823
-6.000 0.0286 0.06584 0.06056 -0.1078 0.8494 0.0850
-5.500 0.0817 0.05405 0.04824 -0.1254 0.8310 0.0658
-5.250 0.1038 0.05188 0.04603 -0.1260 0.8247 0.0671
-4.750 0.1597 0.04182 0.03524 -0.1349 0.8089 0.0547
-4.500 0.1828 0.03934 0.03259 -0.1361 0.8008 0.0540
-4.250 0.2101 0.03650 0.02947 -0.1379 0.7939 0.0533
-4.000 0.2376 0.03365 0.02628 -0.1394 0.7872 0.0527
-3.750 0.2644 0.03099 0.02323 -0.1405 0.7798 0.0525
-3.500 0.2952 0.02854 0.02026 -0.1418 0.7744 0.0537
-3.250 0.3210 0.02674 0.01804 -0.1419 0.7666 0.0544
-3.000 0.3498 0.02518 0.01608 -0.1423 0.7604 0.0545
-2.750 0.3782 0.02391 0.01445 -0.1425 0.7543 0.0548
-2.500 0.4049 0.02295 0.01319 -0.1423 0.7471 0.0553
-2.250 0.4342 0.02200 0.01206 -0.1426 0.7416 0.0560
-2.000 0.4596 0.02140 0.01139 -0.1422 0.7345 0.0572
-1.750 0.4870 0.02094 0.01084 -0.1421 0.7281 0.0595
-1.500 0.5153 0.02047 0.01025 -0.1421 0.7224 0.0623
-1.250 0.5406 0.02007 0.00976 -0.1416 0.7150 0.0643
-1.000 0.5690 0.01960 0.00922 -0.1415 0.7093 0.0663
-0.750 0.5942 0.01934 0.00898 -0.1410 0.7023 0.0691
-0.500 0.6211 0.01909 0.00869 -0.1407 0.6956 0.0733
-0.250 0.6487 0.01887 0.00846 -0.1405 0.6896 0.0809
0.000 0.6736 0.01877 0.00842 -0.1399 0.6820 0.0950
0.250 0.7022 0.01856 0.00821 -0.1398 0.6761 0.1275
0.500 0.7270 0.01853 0.00822 -0.1393 0.6685 0.1670
0.750 0.7545 0.01831 0.00815 -0.1393 0.6620 0.2239
1.000 0.7783 0.01662 0.00822 -0.1382 0.6558 1.0000
1.250 0.8037 0.01681 0.00824 -0.1376 0.6486 1.0000
1.500 0.8323 0.01693 0.00814 -0.1376 0.6430 1.0000
1.750 0.8559 0.01718 0.00831 -0.1368 0.6351 1.0000
2.000 0.8829 0.01734 0.00832 -0.1366 0.6287 1.0000
2.250 0.9081 0.01756 0.00845 -0.1360 0.6217 1.0000
2.500 0.9334 0.01777 0.00857 -0.1355 0.6144 1.0000
2.750 0.9604 0.01794 0.00863 -0.1353 0.6082 1.0000
3.000 0.9836 0.01821 0.00888 -0.1345 0.6000 1.0000
3.250 1.0111 0.01836 0.00893 -0.1343 0.5938 1.0000
3.500 1.0333 0.01866 0.00925 -0.1334 0.5853 1.0000
3.750 1.0595 0.01885 0.00936 -0.1331 0.5784 1.0000
4.000 1.0823 0.01914 0.00967 -0.1322 0.5701 1.0000
4.250 1.1074 0.01935 0.00983 -0.1317 0.5626 1.0000
4.500 1.1301 0.01964 0.01014 -0.1308 0.5543 1.0000
4.750 1.1545 0.01988 0.01036 -0.1302 0.5465 1.0000
5.000 1.1767 0.02019 0.01069 -0.1293 0.5379 1.0000
5.250 1.2007 0.02042 0.01091 -0.1286 0.5299 1.0000
5.500 1.2221 0.02077 0.01131 -0.1275 0.5210 1.0000
5.750 1.2461 0.02101 0.01153 -0.1269 0.5133 1.0000
6.000 1.2663 0.02140 0.01198 -0.1256 0.5039 1.0000
6.500 1.3092 0.02208 0.01272 -0.1236 0.4866 1.0000
6.750 1.3315 0.02240 0.01303 -0.1227 0.4785 1.0000
7.000 1.3503 0.02280 0.01347 -0.1212 0.4683 1.0000
7.250 1.3691 0.02321 0.01393 -0.1198 0.4583 1.0000
7.500 1.3896 0.02355 0.01423 -0.1186 0.4491 1.0000
7.750 1.4055 0.02408 0.01486 -0.1167 0.4390 1.0000
8.000 1.4247 0.02454 0.01531 -0.1154 0.4309 1.0000
8.250 1.4409 0.02508 0.01596 -0.1137 0.4220 1.0000
8.500 1.4578 0.02561 0.01651 -0.1120 0.4140 1.0000
8.750 1.4718 0.02619 0.01716 -0.1099 0.4054 1.0000
9.000 1.4859 0.02680 0.01782 -0.1079 0.3969 1.0000
9.250 1.4991 0.02743 0.01852 -0.1058 0.3881 1.0000
9.500 1.5106 0.02818 0.01936 -0.1035 0.3793 1.0000
9.750 1.5240 0.02887 0.02008 -0.1016 0.3713 1.0000
10.000 1.5344 0.02974 0.02108 -0.0993 0.3632 1.0000
10.250 1.5467 0.03053 0.02192 -0.0974 0.3556 1.0000
10.500 1.5545 0.03155 0.02308 -0.0950 0.3469 1.0000
10.750 1.5628 0.03249 0.02403 -0.0927 0.3374 1.0000
11.000 1.5640 0.03386 0.02553 -0.0899 0.3260 1.0000
11.250 1.5650 0.03533 0.02708 -0.0873 0.3141 1.0000
11.500 1.5648 0.03697 0.02876 -0.0848 0.3017 1.0000
11.750 1.5637 0.03879 0.03062 -0.0824 0.2893 1.0000
12.000 1.5622 0.04084 0.03278 -0.0804 0.2773 1.0000
12.250 1.5610 0.04301 0.03506 -0.0786 0.2657 1.0000
12.500 1.5592 0.04532 0.03746 -0.0770 0.2543 1.0000
12.750 1.5561 0.04784 0.04004 -0.0754 0.2427 1.0000
13.000 1.5510 0.05066 0.04291 -0.0740 0.2302 1.0000
13.250 1.5447 0.05375 0.04607 -0.0728 0.2170 1.0000
13.500 1.5376 0.05707 0.04944 -0.0718 0.2039 1.0000
13.750 1.5300 0.06057 0.05301 -0.0710 0.1914 1.0000
14.000 1.5216 0.06431 0.05679 -0.0704 0.1796 1.0000
14.250 1.5112 0.06840 0.06091 -0.0700 0.1678 1.0000
14.500 1.4992 0.07285 0.06539 -0.0697 0.1555 1.0000
14.750 1.4867 0.07755 0.07013 -0.0698 0.1431 1.0000
15.000 1.4746 0.08234 0.07499 -0.0700 0.1305 1.0000
15.250 1.4604 0.08755 0.08024 -0.0705 0.1159 1.0000
15.500 1.4426 0.09344 0.08611 -0.0714 0.1014 1.0000
15.750 1.4243 0.09956 0.09220 -0.0725 0.0902 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 332 (PFALZ 61) AIRFOIL (goe332-il)