GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 622 AIRFOIL (goe622-il) Reynolds number: 200,000 Max Cl/Cd: 64.94 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe622-il-200000-n5.txt Download as CSV file: xf-goe622-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 622 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5003 0.08509 0.08169 -0.0184 1.0000 0.0119
-8.250 -0.5026 0.08049 0.07714 -0.0212 1.0000 0.0120
-8.000 -0.5062 0.07555 0.07226 -0.0250 1.0000 0.0121
-7.750 -0.5057 0.06948 0.06621 -0.0311 1.0000 0.0122
-7.500 -0.5039 0.06264 0.05934 -0.0371 1.0000 0.0122
-7.000 -0.5005 0.04637 0.04278 -0.0456 1.0000 0.0121
-6.750 -0.5006 0.03637 0.03223 -0.0471 1.0000 0.0121
-6.500 -0.4927 0.03018 0.02538 -0.0466 1.0000 0.0122
-6.250 -0.4789 0.02621 0.02082 -0.0455 1.0000 0.0125
-6.000 -0.4618 0.02339 0.01745 -0.0442 1.0000 0.0130
-5.750 -0.4425 0.02125 0.01488 -0.0430 1.0000 0.0136
-5.500 -0.4109 0.01949 0.01273 -0.0441 0.9971 0.0146
-5.250 -0.3764 0.01741 0.01033 -0.0460 0.9930 0.0171
-5.000 -0.3427 0.01628 0.00902 -0.0474 0.9888 0.0193
-4.750 -0.3093 0.01532 0.00788 -0.0486 0.9848 0.0226
-4.500 -0.2765 0.01451 0.00694 -0.0498 0.9801 0.0276
-4.250 -0.2423 0.01389 0.00620 -0.0513 0.9762 0.0352
-4.000 -0.2104 0.01334 0.00561 -0.0523 0.9702 0.0460
-3.750 -0.1761 0.01283 0.00509 -0.0538 0.9660 0.0601
-3.500 -0.1452 0.01246 0.00478 -0.0545 0.9589 0.0768
-3.250 -0.1111 0.01214 0.00443 -0.0559 0.9537 0.0928
-3.000 -0.0800 0.01187 0.00412 -0.0566 0.9461 0.1080
-2.750 -0.0464 0.01159 0.00382 -0.0578 0.9401 0.1254
-2.500 -0.0153 0.01133 0.00355 -0.0585 0.9321 0.1405
-2.250 0.0177 0.01105 0.00330 -0.0596 0.9255 0.1588
-2.000 0.0476 0.01075 0.00309 -0.0601 0.9164 0.1907
-1.750 0.0778 0.01032 0.00291 -0.0607 0.9081 0.2600
-1.500 0.1061 0.00981 0.00278 -0.0610 0.8987 0.3724
-1.250 0.1319 0.00923 0.00269 -0.0606 0.8884 0.5202
-1.000 0.1550 0.00863 0.00266 -0.0593 0.8778 0.6858
-0.750 0.1782 0.00825 0.00267 -0.0573 0.8673 0.8375
-0.500 0.2181 0.00812 0.00261 -0.0592 0.8569 0.9454
-0.250 0.2609 0.00808 0.00249 -0.0623 0.8452 0.9983
0.000 0.2878 0.00812 0.00242 -0.0619 0.8311 1.0000
0.250 0.3136 0.00816 0.00236 -0.0613 0.8152 1.0000
0.500 0.3388 0.00820 0.00228 -0.0605 0.7913 1.0000
0.750 0.3636 0.00827 0.00219 -0.0595 0.7594 1.0000
1.000 0.3884 0.00837 0.00211 -0.0586 0.7243 1.0000
1.250 0.4136 0.00850 0.00208 -0.0579 0.6939 1.0000
1.500 0.4390 0.00866 0.00208 -0.0572 0.6671 1.0000
1.750 0.4648 0.00882 0.00214 -0.0567 0.6450 1.0000
2.250 0.5165 0.00917 0.00230 -0.0557 0.6062 1.0000
2.500 0.5423 0.00937 0.00243 -0.0552 0.5860 1.0000
2.750 0.5682 0.00957 0.00256 -0.0548 0.5683 1.0000
3.000 0.5943 0.00976 0.00271 -0.0544 0.5503 1.0000
3.250 0.6201 0.00997 0.00290 -0.0540 0.5294 1.0000
3.500 0.6458 0.01020 0.00307 -0.0536 0.5067 1.0000
3.750 0.6711 0.01045 0.00326 -0.0531 0.4775 1.0000
4.000 0.6959 0.01074 0.00346 -0.0525 0.4407 1.0000
4.250 0.7202 0.01109 0.00370 -0.0519 0.3960 1.0000
4.750 0.7666 0.01212 0.00433 -0.0505 0.2983 1.0000
5.000 0.7883 0.01287 0.00477 -0.0497 0.2371 1.0000
5.250 0.8112 0.01351 0.00523 -0.0491 0.1892 1.0000
5.500 0.8338 0.01420 0.00570 -0.0485 0.1432 1.0000
5.750 0.8566 0.01487 0.00623 -0.0479 0.1155 1.0000
6.000 0.8801 0.01546 0.00677 -0.0473 0.0864 1.0000
6.250 0.8995 0.01663 0.00759 -0.0463 0.0291 1.0000
6.500 0.9209 0.01755 0.00850 -0.0453 0.0192 1.0000
6.750 0.9436 0.01830 0.00939 -0.0444 0.0169 1.0000
7.000 0.9652 0.01917 0.01040 -0.0434 0.0150 1.0000
7.250 0.9849 0.02031 0.01172 -0.0422 0.0128 1.0000
7.500 1.0016 0.02181 0.01344 -0.0405 0.0117 1.0000
7.750 1.0197 0.02306 0.01482 -0.0391 0.0113 1.0000
8.000 1.0376 0.02435 0.01626 -0.0377 0.0109 1.0000
8.250 1.0550 0.02578 0.01784 -0.0362 0.0105 1.0000
8.500 1.0726 0.02718 0.01945 -0.0348 0.0099 1.0000
8.750 1.0900 0.02854 0.02095 -0.0336 0.0091 1.0000
9.000 1.1062 0.03001 0.02256 -0.0322 0.0085 1.0000
9.250 1.1209 0.03181 0.02452 -0.0307 0.0082 1.0000
9.500 1.1341 0.03380 0.02669 -0.0291 0.0079 1.0000
9.750 1.1452 0.03604 0.02913 -0.0274 0.0077 1.0000
10.000 1.1538 0.03861 0.03191 -0.0255 0.0076 1.0000
10.250 1.1586 0.04152 0.03507 -0.0234 0.0074 1.0000
10.500 1.1581 0.04456 0.03839 -0.0208 0.0074 1.0000
10.750 1.1530 0.04762 0.04172 -0.0180 0.0073 1.0000
11.000 1.1445 0.05085 0.04521 -0.0155 0.0073 1.0000
11.500 1.1223 0.05796 0.05280 -0.0130 0.0073 1.0000
11.750 1.1094 0.06212 0.05719 -0.0130 0.0073 1.0000
12.750 1.0529 0.08395 0.07985 -0.0232 0.0073 1.0000
13.000 1.0384 0.09088 0.08696 -0.0278 0.0074 1.0000
13.250 1.0228 0.09870 0.09495 -0.0331 0.0074 1.0000
13.500 1.0062 0.10769 0.10410 -0.0392 0.0075 1.0000
13.750 0.9837 0.11961 0.11620 -0.0471 0.0077 1.0000
14.000 0.9321 0.14342 0.14017 -0.0596 0.0084 1.0000
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Polar data table (+)
Polar graphs
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