GOE 147 (MVA H.6) AIRFOIL (goe147-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 147 (MVA H.6) AIRFOIL (goe147-il) Reynolds number: 200,000 Max Cl/Cd: 66.1 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe147-il-200000-n5.txt Download as CSV file: xf-goe147-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 147 (MVA H.6) 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4549 0.09393 0.09066 0.0058 1.0000 0.0164
-7.750 -0.4515 0.08997 0.08675 0.0036 1.0000 0.0161
-7.500 -0.4486 0.08575 0.08257 0.0007 1.0000 0.0157
-7.250 -0.4418 0.08049 0.07735 -0.0042 1.0000 0.0150
-7.000 -0.4327 0.07408 0.07095 -0.0108 1.0000 0.0144
-6.750 -0.4205 0.06519 0.06204 -0.0202 1.0000 0.0137
-6.500 -0.4027 0.05455 0.05129 -0.0310 1.0000 0.0133
-6.250 -0.3731 0.03593 0.03199 -0.0483 1.0000 0.0129
-6.000 -0.3442 0.02688 0.02199 -0.0544 1.0000 0.0130
-5.750 -0.3181 0.02276 0.01727 -0.0561 1.0000 0.0135
-5.500 -0.2919 0.02092 0.01517 -0.0566 1.0000 0.0145
-5.250 -0.2653 0.01966 0.01372 -0.0567 1.0000 0.0159
-5.000 -0.2384 0.01825 0.01204 -0.0568 1.0000 0.0170
-4.750 -0.2115 0.01697 0.01054 -0.0566 1.0000 0.0180
-4.500 -0.1847 0.01598 0.00936 -0.0563 1.0000 0.0190
-4.250 -0.1583 0.01504 0.00826 -0.0561 1.0000 0.0208
-4.000 -0.1229 0.01412 0.00721 -0.0578 0.9870 0.0244
-3.750 -0.0849 0.01341 0.00629 -0.0597 0.9642 0.0281
-3.500 -0.0464 0.01254 0.00528 -0.0616 0.9299 0.0371
-3.250 -0.0119 0.01222 0.00510 -0.0626 0.8856 0.0805
-3.000 0.0162 0.01259 0.00510 -0.0617 0.8392 0.0990
-2.750 0.0417 0.01285 0.00513 -0.0607 0.7961 0.1087
-2.500 0.0674 0.01294 0.00502 -0.0599 0.7592 0.1174
-2.250 0.0940 0.01294 0.00481 -0.0593 0.7283 0.1225
-2.000 0.1210 0.01291 0.00453 -0.0588 0.7004 0.1263
-1.750 0.1480 0.01269 0.00419 -0.0584 0.6752 0.1289
-1.500 0.1753 0.01255 0.00391 -0.0580 0.6528 0.1298
-1.250 0.2028 0.01244 0.00366 -0.0577 0.6329 0.1309
-1.000 0.2304 0.01234 0.00345 -0.0575 0.6148 0.1320
-0.750 0.2580 0.01227 0.00327 -0.0572 0.5976 0.1332
-0.500 0.2857 0.01221 0.00311 -0.0570 0.5809 0.1346
-0.250 0.3135 0.01216 0.00298 -0.0568 0.5649 0.1361
0.000 0.3413 0.01214 0.00288 -0.0566 0.5490 0.1383
0.250 0.3691 0.01213 0.00280 -0.0565 0.5318 0.1408
0.500 0.3968 0.01212 0.00275 -0.0563 0.5130 0.1437
0.750 0.4245 0.01215 0.00271 -0.0561 0.4930 0.1468
1.000 0.4523 0.01218 0.00270 -0.0560 0.4709 0.1508
1.250 0.4799 0.01223 0.00270 -0.0559 0.4470 0.1565
1.500 0.5075 0.01230 0.00272 -0.0557 0.4203 0.1665
1.750 0.5349 0.01242 0.00277 -0.0556 0.3914 0.1835
2.000 0.5622 0.01256 0.00285 -0.0555 0.3657 0.2067
2.250 0.5894 0.01271 0.00295 -0.0554 0.3445 0.2329
2.500 0.6166 0.01281 0.00309 -0.0554 0.3273 0.2717
3.250 0.6920 0.01192 0.00353 -0.0534 0.2924 1.0000
3.500 0.7193 0.01217 0.00372 -0.0532 0.2828 1.0000
3.750 0.7465 0.01244 0.00392 -0.0531 0.2733 1.0000
4.000 0.7738 0.01267 0.00416 -0.0530 0.2641 1.0000
4.250 0.8009 0.01295 0.00439 -0.0528 0.2556 1.0000
4.500 0.8281 0.01320 0.00464 -0.0527 0.2470 1.0000
4.750 0.8551 0.01348 0.00493 -0.0525 0.2392 1.0000
5.000 0.8820 0.01376 0.00521 -0.0524 0.2310 1.0000
5.250 0.9088 0.01405 0.00552 -0.0522 0.2233 1.0000
5.500 0.9354 0.01437 0.00585 -0.0520 0.2151 1.0000
5.750 0.9621 0.01467 0.00622 -0.0518 0.2073 1.0000
6.000 0.9884 0.01503 0.00659 -0.0516 0.1996 1.0000
6.250 1.0148 0.01537 0.00699 -0.0514 0.1921 1.0000
6.500 1.0407 0.01577 0.00742 -0.0512 0.1854 1.0000
6.750 1.0668 0.01614 0.00791 -0.0509 0.1793 1.0000
7.000 1.0924 0.01656 0.00834 -0.0506 0.1714 1.0000
7.250 1.1181 0.01694 0.00877 -0.0504 0.1609 1.0000
7.500 1.1432 0.01737 0.00921 -0.0502 0.1474 1.0000
7.750 1.1675 0.01793 0.00972 -0.0499 0.1318 1.0000
8.000 1.1916 0.01854 0.01037 -0.0496 0.1175 1.0000
8.250 1.2149 0.01928 0.01108 -0.0492 0.0886 1.0000
8.500 1.2284 0.02172 0.01308 -0.0481 0.0259 1.0000
8.750 1.2491 0.02287 0.01435 -0.0473 0.0198 1.0000
9.000 1.2684 0.02417 0.01583 -0.0464 0.0165 1.0000
9.250 1.2879 0.02532 0.01720 -0.0455 0.0150 1.0000
9.500 1.3057 0.02661 0.01870 -0.0446 0.0138 1.0000
9.750 1.3215 0.02803 0.02030 -0.0436 0.0126 1.0000
10.000 1.3314 0.02999 0.02249 -0.0422 0.0114 1.0000
10.250 1.3381 0.03204 0.02473 -0.0406 0.0108 1.0000
10.500 1.3451 0.03388 0.02673 -0.0390 0.0104 1.0000
10.750 1.3468 0.03588 0.02889 -0.0370 0.0101 1.0000
11.000 1.3442 0.03819 0.03136 -0.0352 0.0099 1.0000
11.250 1.3407 0.04104 0.03437 -0.0345 0.0097 1.0000
11.500 1.3368 0.04443 0.03791 -0.0349 0.0095 1.0000
11.750 1.3325 0.04824 0.04188 -0.0359 0.0093 1.0000
12.000 1.3279 0.05235 0.04614 -0.0373 0.0091 1.0000
12.250 1.3230 0.05663 0.05056 -0.0389 0.0089 1.0000
12.500 1.3175 0.06106 0.05512 -0.0405 0.0088 1.0000
12.750 1.3115 0.06559 0.05977 -0.0421 0.0086 1.0000
13.000 1.3049 0.07022 0.06452 -0.0438 0.0084 1.0000
13.250 1.2975 0.07504 0.06944 -0.0457 0.0082 1.0000
13.500 1.2900 0.07986 0.07437 -0.0474 0.0081 1.0000
13.750 1.2821 0.08477 0.07938 -0.0492 0.0079 1.0000
14.000 1.2741 0.08970 0.08440 -0.0510 0.0078 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 147 (MVA H.6) AIRFOIL (goe147-il)