GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: GOE 415 AIRFOIL (goe415-il) Reynolds number: 1,000,000 Max Cl/Cd: 109.75 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe415-il-1000000.txt Download as CSV file: xf-goe415-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 415 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4241 0.09030 0.08883 -0.0047 1.0000 0.0077
-7.750 -0.4191 0.08657 0.08512 -0.0071 1.0000 0.0077
-7.500 -0.4142 0.08277 0.08134 -0.0100 1.0000 0.0078
-7.250 -0.4043 0.07837 0.07695 -0.0146 1.0000 0.0078
-7.000 -0.3918 0.07363 0.07222 -0.0200 1.0000 0.0078
-6.750 -0.3775 0.06865 0.06723 -0.0256 1.0000 0.0078
-6.500 -0.3622 0.06353 0.06209 -0.0308 1.0000 0.0078
-6.250 -0.3466 0.05833 0.05685 -0.0354 0.9996 0.0078
-6.000 -0.3189 0.04829 0.04665 -0.0464 0.9896 0.0082
-5.750 -0.2895 0.04414 0.04239 -0.0513 0.9789 0.0084
-5.500 -0.2597 0.04017 0.03827 -0.0554 0.9646 0.0088
-5.250 -0.2319 0.03616 0.03406 -0.0581 0.9470 0.0097
-5.000 -0.2000 0.03276 0.03032 -0.0583 0.9305 0.0114
-4.750 -0.1763 0.02949 0.02672 -0.0584 0.9153 0.0115
-3.500 -0.0557 0.01488 0.01028 -0.0563 0.8198 0.0119
-3.250 -0.0314 0.01252 0.00746 -0.0551 0.7649 0.0107
-3.000 -0.0056 0.01181 0.00636 -0.0543 0.7125 0.0114
-2.750 0.0198 0.01026 0.00449 -0.0536 0.6754 0.0123
-2.500 0.0463 0.00985 0.00391 -0.0532 0.6390 0.0130
-2.250 0.0733 0.00968 0.00360 -0.0529 0.6055 0.0139
-2.000 0.1005 0.00953 0.00330 -0.0527 0.5772 0.0153
-1.750 0.1279 0.00931 0.00296 -0.0524 0.5555 0.0165
-1.500 0.1547 0.00872 0.00222 -0.0521 0.5383 0.0183
-1.250 0.1824 0.00854 0.00199 -0.0520 0.5244 0.0204
-1.000 0.2103 0.00848 0.00187 -0.0519 0.5124 0.0234
-0.750 0.2382 0.00842 0.00174 -0.0518 0.5013 0.0252
-0.250 0.2939 0.00809 0.00125 -0.0515 0.4799 0.0351
0.000 0.3219 0.00803 0.00112 -0.0514 0.4694 0.0427
0.250 0.3481 0.00729 0.00108 -0.0515 0.4585 0.3136
0.500 0.3742 0.00682 0.00112 -0.0514 0.4465 0.5127
0.750 0.4101 0.00566 0.00117 -0.0532 0.4316 1.0000
1.000 0.4374 0.00576 0.00117 -0.0530 0.4171 1.0000
1.250 0.4648 0.00587 0.00119 -0.0528 0.4031 1.0000
1.500 0.4922 0.00599 0.00121 -0.0526 0.3884 1.0000
1.750 0.5195 0.00611 0.00125 -0.0525 0.3742 1.0000
2.000 0.5469 0.00624 0.00130 -0.0523 0.3615 1.0000
2.250 0.5742 0.00638 0.00136 -0.0522 0.3493 1.0000
2.500 0.6015 0.00651 0.00143 -0.0521 0.3378 1.0000
2.750 0.6289 0.00664 0.00151 -0.0520 0.3269 1.0000
3.000 0.6562 0.00679 0.00159 -0.0518 0.3157 1.0000
3.250 0.6834 0.00694 0.00169 -0.0517 0.3050 1.0000
3.500 0.7107 0.00709 0.00178 -0.0516 0.2960 1.0000
3.750 0.7381 0.00722 0.00190 -0.0515 0.2886 1.0000
4.250 0.7926 0.00752 0.00213 -0.0513 0.2714 1.0000
4.500 0.8197 0.00769 0.00226 -0.0512 0.2609 1.0000
4.750 0.8467 0.00787 0.00242 -0.0511 0.2523 1.0000
5.000 0.8737 0.00804 0.00256 -0.0510 0.2412 1.0000
5.250 0.9006 0.00824 0.00271 -0.0509 0.2289 1.0000
5.500 0.9274 0.00845 0.00288 -0.0508 0.2164 1.0000
5.750 0.9535 0.00876 0.00310 -0.0506 0.1955 1.0000
6.000 0.9770 0.00947 0.00349 -0.0502 0.1378 1.0000
6.250 0.9988 0.01050 0.00416 -0.0496 0.0802 1.0000
6.500 1.0192 0.01182 0.00515 -0.0486 0.0143 1.0000
6.750 1.0450 0.01217 0.00559 -0.0483 0.0126 1.0000
7.000 1.0703 0.01259 0.00608 -0.0480 0.0113 1.0000
7.250 1.0950 0.01311 0.00666 -0.0476 0.0101 1.0000
7.500 1.1182 0.01387 0.00754 -0.0469 0.0090 1.0000
7.750 1.1400 0.01484 0.00866 -0.0461 0.0083 1.0000
8.000 1.1645 0.01527 0.00915 -0.0457 0.0078 1.0000
8.250 1.1881 0.01584 0.00978 -0.0452 0.0072 1.0000
8.500 1.2103 0.01658 0.01060 -0.0445 0.0067 1.0000
8.750 1.2317 0.01739 0.01148 -0.0437 0.0063 1.0000
9.000 1.2521 0.01827 0.01244 -0.0429 0.0060 1.0000
9.250 1.2704 0.01934 0.01359 -0.0418 0.0057 1.0000
9.500 1.2821 0.02109 0.01547 -0.0400 0.0054 1.0000
9.750 1.2843 0.02380 0.01831 -0.0371 0.0051 1.0000
10.000 1.3037 0.02447 0.01907 -0.0362 0.0050 1.0000
10.250 1.3179 0.02566 0.02035 -0.0347 0.0048 1.0000
10.500 1.3281 0.02718 0.02198 -0.0328 0.0047 1.0000
10.750 1.3354 0.02870 0.02359 -0.0305 0.0045 1.0000
11.000 1.3408 0.03019 0.02518 -0.0282 0.0044 1.0000
11.250 1.3450 0.03186 0.02696 -0.0262 0.0042 1.0000
11.500 1.3483 0.03372 0.02892 -0.0245 0.0041 1.0000
11.750 1.3497 0.03599 0.03131 -0.0231 0.0040 1.0000
12.000 1.3496 0.03861 0.03406 -0.0219 0.0040 1.0000
12.250 1.3492 0.04140 0.03700 -0.0211 0.0039 1.0000
12.500 1.3468 0.04463 0.04038 -0.0206 0.0039 1.0000
12.750 1.3440 0.04803 0.04392 -0.0206 0.0039 1.0000
13.000 1.3386 0.05195 0.04800 -0.0208 0.0039 1.0000
13.250 1.3320 0.05612 0.05233 -0.0217 0.0039 1.0000
13.500 1.3232 0.06076 0.05713 -0.0230 0.0039 1.0000
13.750 1.3131 0.06574 0.06227 -0.0248 0.0039 1.0000
14.000 1.3007 0.07120 0.06789 -0.0269 0.0040 1.0000
14.250 1.2878 0.07685 0.07369 -0.0294 0.0040 1.0000
14.500 1.2711 0.08327 0.08027 -0.0323 0.0040 1.0000
14.750 1.2566 0.08957 0.08671 -0.0354 0.0040 1.0000
15.000 1.2392 0.09662 0.09391 -0.0390 0.0041 1.0000
15.250 1.2230 0.10382 0.10125 -0.0428 0.0041 1.0000
15.500 1.2073 0.11122 0.10877 -0.0469 0.0041 1.0000
15.750 1.1904 0.11910 0.11678 -0.0513 0.0042 1.0000
16.000 1.1737 0.12732 0.12513 -0.0560 0.0043 1.0000
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Polar data table (+)
Polar graphs
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