GOE 143 (MVA H.20) AIRFOIL (goe143-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 143 (MVA H.20) AIRFOIL (goe143-il) Reynolds number: 500,000 Max Cl/Cd: 86.8 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe143-il-500000-n5.txt Download as CSV file: xf-goe143-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 143 (MVA H.20) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4485 0.08761 0.08542 -0.0115 1.0000 0.0120
-8.250 -0.4509 0.08266 0.08050 -0.0147 1.0000 0.0126
-8.000 -0.4570 0.07311 0.07098 -0.0246 1.0000 0.0136
-7.750 -0.4425 0.07021 0.06808 -0.0282 1.0000 0.0139
-7.500 -0.4270 0.06745 0.06532 -0.0314 1.0000 0.0143
-7.000 -0.3985 0.04873 0.04636 -0.0520 0.9975 0.0181
-6.750 -0.3649 0.04684 0.04440 -0.0560 0.9932 0.0185
-6.500 -0.3272 0.02901 0.02572 -0.0713 0.9853 0.0235
-6.250 -0.2933 0.02897 0.02565 -0.0730 0.9755 0.0240
-6.000 -0.2613 0.02858 0.02519 -0.0744 0.9632 0.0245
-5.750 -0.2323 0.02750 0.02399 -0.0753 0.9506 0.0255
-5.500 -0.2063 0.01800 0.01338 -0.0781 0.9385 0.0296
-5.250 -0.1799 0.01630 0.01138 -0.0783 0.9239 0.0306
-5.000 -0.1534 0.01588 0.01085 -0.0780 0.9068 0.0311
-4.750 -0.1272 0.01563 0.01050 -0.0775 0.8881 0.0318
-4.500 -0.1010 0.01508 0.00977 -0.0771 0.8698 0.0325
-4.250 -0.0745 0.01428 0.00874 -0.0767 0.8526 0.0331
-4.000 -0.0476 0.01352 0.00775 -0.0764 0.8376 0.0338
-3.750 -0.0205 0.01286 0.00690 -0.0762 0.8232 0.0344
-3.500 0.0069 0.01227 0.00612 -0.0760 0.8100 0.0350
-3.250 0.0345 0.01178 0.00548 -0.0758 0.7971 0.0357
-3.000 0.0622 0.01145 0.00502 -0.0756 0.7841 0.0365
-2.500 0.1177 0.01087 0.00420 -0.0753 0.7555 0.0374
-2.250 0.1453 0.01029 0.00352 -0.0751 0.7398 0.0385
-2.000 0.1731 0.01001 0.00317 -0.0750 0.7211 0.0394
-1.750 0.2009 0.00985 0.00292 -0.0748 0.6997 0.0403
-1.500 0.2285 0.00973 0.00271 -0.0747 0.6761 0.0413
-1.000 0.2838 0.00959 0.00233 -0.0743 0.6222 0.0435
-0.750 0.3114 0.00956 0.00217 -0.0742 0.5960 0.0448
-0.500 0.3391 0.00957 0.00206 -0.0741 0.5694 0.0462
-0.250 0.3668 0.00962 0.00200 -0.0740 0.5445 0.0474
0.000 0.3947 0.00954 0.00183 -0.0739 0.5227 0.0500
0.250 0.4224 0.00957 0.00176 -0.0738 0.4981 0.0525
0.500 0.4499 0.00966 0.00173 -0.0738 0.4697 0.0549
0.750 0.4775 0.00976 0.00171 -0.0737 0.4440 0.0575
1.000 0.5052 0.00985 0.00171 -0.0736 0.4248 0.0599
1.250 0.5332 0.00988 0.00172 -0.0736 0.4113 0.0657
1.750 0.5889 0.00999 0.00183 -0.0735 0.3840 0.0940
2.000 0.6167 0.01005 0.00189 -0.0735 0.3714 0.1113
2.250 0.6445 0.00993 0.00199 -0.0736 0.3574 0.2277
2.500 0.6667 0.00835 0.00216 -0.0727 0.3433 1.0000
2.750 0.6944 0.00852 0.00224 -0.0726 0.3293 1.0000
3.000 0.7219 0.00870 0.00234 -0.0725 0.3166 1.0000
3.250 0.7494 0.00889 0.00245 -0.0724 0.3024 1.0000
3.500 0.7768 0.00910 0.00258 -0.0723 0.2867 1.0000
3.750 0.8040 0.00932 0.00273 -0.0722 0.2692 1.0000
4.000 0.8310 0.00958 0.00289 -0.0721 0.2496 1.0000
4.250 0.8576 0.00988 0.00308 -0.0719 0.2269 1.0000
4.500 0.8838 0.01026 0.00333 -0.0717 0.1979 1.0000
4.750 0.9095 0.01072 0.00362 -0.0715 0.1665 1.0000
5.000 0.9353 0.01116 0.00393 -0.0713 0.1437 1.0000
5.500 0.9875 0.01186 0.00453 -0.0708 0.1205 1.0000
5.750 1.0135 0.01222 0.00484 -0.0706 0.1108 1.0000
6.000 1.0393 0.01258 0.00516 -0.0703 0.1039 1.0000
6.250 1.0656 0.01284 0.00546 -0.0701 0.0999 1.0000
6.500 1.0915 0.01316 0.00576 -0.0698 0.0941 1.0000
6.750 1.1171 0.01351 0.00610 -0.0696 0.0879 1.0000
7.000 1.1427 0.01384 0.00642 -0.0693 0.0803 1.0000
7.250 1.1676 0.01426 0.00680 -0.0689 0.0671 1.0000
7.500 1.1877 0.01540 0.00770 -0.0680 0.0254 1.0000
7.750 1.2112 0.01599 0.00833 -0.0675 0.0198 1.0000
8.000 1.2352 0.01649 0.00888 -0.0669 0.0173 1.0000
8.250 1.2582 0.01711 0.00955 -0.0663 0.0149 1.0000
8.500 1.2814 0.01767 0.01019 -0.0657 0.0134 1.0000
8.750 1.3045 0.01821 0.01080 -0.0651 0.0123 1.0000
9.000 1.3267 0.01883 0.01148 -0.0644 0.0113 1.0000
9.250 1.3474 0.01963 0.01234 -0.0635 0.0104 1.0000
9.500 1.3674 0.02047 0.01327 -0.0625 0.0097 1.0000
9.750 1.3879 0.02118 0.01408 -0.0616 0.0092 1.0000
10.000 1.4077 0.02193 0.01492 -0.0607 0.0086 1.0000
10.250 1.4267 0.02270 0.01576 -0.0596 0.0081 1.0000
10.500 1.4444 0.02356 0.01668 -0.0585 0.0077 1.0000
10.750 1.4590 0.02466 0.01787 -0.0569 0.0073 1.0000
11.000 1.4703 0.02597 0.01930 -0.0550 0.0070 1.0000
11.250 1.4827 0.02705 0.02050 -0.0533 0.0068 1.0000
11.500 1.4909 0.02822 0.02180 -0.0509 0.0066 1.0000
11.750 1.4960 0.02954 0.02323 -0.0483 0.0064 1.0000
12.000 1.4997 0.03105 0.02487 -0.0460 0.0062 1.0000
12.250 1.5022 0.03280 0.02674 -0.0440 0.0061 1.0000
12.500 1.5038 0.03479 0.02886 -0.0425 0.0059 1.0000
12.750 1.5048 0.03705 0.03123 -0.0414 0.0058 1.0000
13.000 1.5052 0.03957 0.03387 -0.0408 0.0056 1.0000
13.250 1.5041 0.04246 0.03689 -0.0406 0.0055 1.0000
13.500 1.5024 0.04566 0.04020 -0.0408 0.0054 1.0000
13.750 1.4984 0.04935 0.04402 -0.0415 0.0053 1.0000
14.000 1.4924 0.05360 0.04840 -0.0427 0.0052 1.0000
14.250 1.4831 0.05851 0.05346 -0.0444 0.0052 1.0000
14.500 1.4723 0.06386 0.05895 -0.0464 0.0051 1.0000
14.750 1.4591 0.06967 0.06490 -0.0486 0.0050 1.0000
15.000 1.4438 0.07589 0.07125 -0.0510 0.0050 1.0000
15.250 1.4282 0.08216 0.07765 -0.0534 0.0050 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 143 (MVA H.20) AIRFOIL (goe143-il)