GOE 630 AIRFOIL (goe630-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 630 AIRFOIL (goe630-il) Reynolds number: 200,000 Max Cl/Cd: 74.88 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe630-il-200000.txt Download as CSV file: xf-goe630-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 630 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.0118 0.10754 0.10405 -0.0856 0.9575 0.0425
-10.000 0.0027 0.10476 0.10125 -0.0926 0.9504 0.0446
-9.750 0.0142 0.10183 0.09831 -0.1025 0.9438 0.0451
-9.500 0.0463 0.09602 0.09247 -0.1026 0.9419 0.0457
-9.250 0.0688 0.09242 0.08884 -0.1052 0.9347 0.0466
-9.000 0.0927 0.08890 0.08528 -0.1097 0.9283 0.0480
-8.750 0.1136 0.08568 0.08201 -0.1142 0.9195 0.0496
-8.500 0.1255 0.08354 0.07981 -0.1217 0.9070 0.0519
-8.250 0.1160 0.08269 0.07891 -0.1268 0.8889 0.0523
-8.000 0.1431 0.07763 0.07383 -0.1250 0.8837 0.0530
-7.750 0.1629 0.07470 0.07083 -0.1254 0.8761 0.0537
-7.500 0.1714 0.07271 0.06881 -0.1248 0.8655 0.0548
-7.250 0.1803 0.07065 0.06670 -0.1254 0.8567 0.0559
-7.000 0.1822 0.06896 0.06500 -0.1249 0.8462 0.0571
-6.750 0.1836 0.06732 0.06334 -0.1255 0.8373 0.0588
-6.500 0.1849 0.06571 0.06157 -0.1377 0.8258 0.0610
-6.250 0.1921 0.06231 0.05823 -0.1345 0.8188 0.0617
-6.000 0.2051 0.06012 0.05602 -0.1319 0.8122 0.0625
-5.750 0.2194 0.05814 0.05400 -0.1315 0.8059 0.0638
-5.500 0.2315 0.05619 0.05203 -0.1320 0.7977 0.0654
-5.250 0.2499 0.05377 0.04950 -0.1349 0.7916 0.0682
-5.000 0.2719 0.04985 0.04535 -0.1437 0.7827 0.0721
-4.750 0.2868 0.04774 0.04324 -0.1422 0.7767 0.0730
-4.500 0.3021 0.04612 0.04160 -0.1412 0.7697 0.0745
-4.250 0.3211 0.04441 0.03983 -0.1415 0.7624 0.0775
-4.000 0.3534 0.04079 0.03579 -0.1475 0.7566 0.0848
-3.750 0.3683 0.03914 0.03422 -0.1461 0.7484 0.0861
-3.500 0.3903 0.03757 0.03256 -0.1461 0.7423 0.0886
-3.250 0.4194 0.03544 0.02998 -0.1487 0.7348 0.0985
-3.000 0.4387 0.03344 0.02802 -0.1481 0.7277 0.1003
-2.750 0.4606 0.03213 0.02667 -0.1477 0.7211 0.1035
-2.500 0.4863 0.03049 0.02470 -0.1483 0.7131 0.1150
-2.000 0.5331 0.02797 0.02193 -0.1477 0.6981 0.1318
-1.750 0.5574 0.02666 0.02057 -0.1475 0.6915 0.1363
-1.500 0.5905 0.02211 0.01484 -0.1466 0.6843 0.0828
-1.250 0.6154 0.02033 0.01284 -0.1460 0.6767 0.0795
-1.000 0.6404 0.01933 0.01159 -0.1452 0.6685 0.0793
-0.750 0.6663 0.01835 0.01034 -0.1444 0.6602 0.0783
-0.500 0.6916 0.01762 0.00941 -0.1436 0.6516 0.0782
-0.250 0.7176 0.01703 0.00864 -0.1429 0.6431 0.0792
0.000 0.7423 0.01666 0.00816 -0.1420 0.6336 0.0818
0.250 0.7687 0.01619 0.00753 -0.1414 0.6251 0.0848
0.500 0.7915 0.01576 0.00715 -0.1403 0.6145 0.0881
0.750 0.8168 0.01549 0.00682 -0.1396 0.6054 0.0925
1.000 0.8403 0.01528 0.00656 -0.1386 0.5951 0.0989
1.250 0.8638 0.01509 0.00640 -0.1376 0.5852 0.1084
1.500 0.8886 0.01491 0.00617 -0.1369 0.5763 0.1231
1.750 0.9102 0.01467 0.00616 -0.1356 0.5657 0.1826
2.000 0.9632 0.01327 0.00624 -0.1413 0.5547 1.0000
2.250 0.9872 0.01347 0.00625 -0.1404 0.5451 1.0000
2.500 1.0101 0.01371 0.00638 -0.1395 0.5353 1.0000
2.750 1.0353 0.01397 0.00643 -0.1389 0.5272 1.0000
3.000 1.0572 0.01422 0.00663 -0.1378 0.5180 1.0000
3.250 1.0828 0.01452 0.00674 -0.1374 0.5106 1.0000
3.500 1.1042 0.01478 0.00699 -0.1362 0.5021 1.0000
3.750 1.1294 0.01509 0.00714 -0.1358 0.4950 1.0000
4.000 1.1509 0.01538 0.00742 -0.1347 0.4872 1.0000
4.250 1.1749 0.01569 0.00762 -0.1340 0.4804 1.0000
4.500 1.1982 0.01603 0.00792 -0.1333 0.4740 1.0000
4.750 1.2208 0.01635 0.00821 -0.1325 0.4677 1.0000
5.000 1.2472 0.01673 0.00845 -0.1324 0.4624 1.0000
5.250 1.2682 0.01707 0.00883 -0.1312 0.4566 1.0000
5.500 1.2909 0.01741 0.00915 -0.1305 0.4511 1.0000
5.750 1.3176 0.01781 0.00944 -0.1305 0.4462 1.0000
6.000 1.3377 0.01818 0.00987 -0.1293 0.4410 1.0000
6.250 1.3597 0.01854 0.01024 -0.1284 0.4360 1.0000
6.500 1.3847 0.01893 0.01059 -0.1281 0.4318 1.0000
6.750 1.4102 0.01941 0.01103 -0.1280 0.4278 1.0000
7.000 1.4294 0.01981 0.01153 -0.1267 0.4235 1.0000
7.250 1.4516 0.02022 0.01197 -0.1259 0.4193 1.0000
7.500 1.4769 0.02063 0.01234 -0.1258 0.4155 1.0000
7.750 1.5022 0.02115 0.01285 -0.1257 0.4118 1.0000
8.000 1.5193 0.02161 0.01343 -0.1241 0.4080 1.0000
8.250 1.5400 0.02207 0.01397 -0.1231 0.4043 1.0000
8.500 1.5639 0.02252 0.01443 -0.1228 0.4010 1.0000
8.750 1.5926 0.02303 0.01490 -0.1233 0.3978 1.0000
9.000 1.6101 0.02361 0.01560 -0.1219 0.3944 1.0000
9.250 1.6256 0.02414 0.01628 -0.1202 0.3910 1.0000
9.500 1.6450 0.02465 0.01687 -0.1191 0.3876 1.0000
9.750 1.6687 0.02510 0.01732 -0.1188 0.3842 1.0000
10.000 1.6951 0.02561 0.01782 -0.1190 0.3799 1.0000
10.250 1.6986 0.02613 0.01853 -0.1152 0.3756 1.0000
10.500 1.7093 0.02647 0.01894 -0.1125 0.3706 1.0000
10.750 1.7383 0.02671 0.01905 -0.1131 0.3650 1.0000
11.000 1.7344 0.02730 0.01986 -0.1081 0.3611 1.0000
11.250 1.7382 0.02784 0.02052 -0.1046 0.3566 1.0000
11.500 1.7534 0.02811 0.02079 -0.1029 0.3515 1.0000
11.750 1.7631 0.02869 0.02143 -0.1005 0.3467 1.0000
12.000 1.7593 0.02949 0.02239 -0.0964 0.3418 1.0000
12.250 1.7674 0.02997 0.02292 -0.0940 0.3365 1.0000
12.500 1.7745 0.03065 0.02364 -0.0916 0.3310 1.0000
12.750 1.7696 0.03180 0.02497 -0.0880 0.3260 1.0000
13.000 1.7751 0.03258 0.02581 -0.0858 0.3206 1.0000
13.250 1.7769 0.03365 0.02694 -0.0833 0.3148 1.0000
13.500 1.7702 0.03523 0.02870 -0.0803 0.3088 1.0000
13.750 1.7778 0.03609 0.02952 -0.0786 0.3025 1.0000
14.000 1.7676 0.03825 0.03192 -0.0759 0.2967 1.0000
14.250 1.7629 0.04008 0.03382 -0.0737 0.2892 1.0000
14.500 1.7542 0.04249 0.03638 -0.0717 0.2813 1.0000
14.750 1.7489 0.04475 0.03867 -0.0701 0.2725 1.0000
15.000 1.7361 0.04803 0.04213 -0.0685 0.2621 1.0000
15.250 1.7242 0.05146 0.04565 -0.0673 0.2503 1.0000
15.500 1.7098 0.05540 0.04966 -0.0664 0.2364 1.0000
15.750 1.6924 0.05991 0.05421 -0.0657 0.2197 1.0000
16.000 1.6682 0.06548 0.05974 -0.0653 0.1995 1.0000
16.250 1.6433 0.07141 0.06564 -0.0652 0.1804 1.0000
16.500 1.6150 0.07803 0.07222 -0.0655 0.1634 1.0000
16.750 1.5854 0.08511 0.07927 -0.0662 0.1473 1.0000
17.000 1.5572 0.09223 0.08637 -0.0672 0.1308 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 630 AIRFOIL (goe630-il)