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    XF Goe222 Il 50000

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    huy4ngoo
    The document contains the results of an XFOIL analysis calculating the polar performance of the GOE 222 airfoil. It lists the coefficient of lift (CL), drag (CD), pressure drag (CDp), and moment (CM) for Reynolds numbers of 0.05e6 and Mach numbers from -8.25 to 8.25 degrees. The analysis also includes boundary layer transition locations on the airfoil's top and bottom surfaces.

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    XF Goe222 Il 50000

    Uploaded by

    huy4ngoo
    41 views2 pages
    The document contains the results of an XFOIL analysis calculating the polar performance of the GOE 222 airfoil. It lists the coefficient of lift (CL), drag (CD), pressure drag (CDp), and moment (CM) for Reynolds numbers of 0.05e6 and Mach numbers from -8.25 to 8.25 degrees. The analysis also includes boundary layer transition locations on the airfoil's top and bottom surfaces.

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    xf-goe222-il-50000

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    The document contains the results of an XFOIL analysis calculating the polar performance of the GOE 222 airfoil. It lists the coefficient of lift (CL), drag (CD), pressure drag (CDp), and moment (CM) for Reynolds numbers of 0.05e6 and Mach numbers from -8.25 to 8.25 degrees. The analysis also includes boundary layer transition locations on the airfoil's top and bottom surfaces.

    Copyright:

    © All Rights Reserved
    Download as TXT, PDF, TXT or read online from Scribd
    Download as txt, pdf, or txt
    41 views2 pages

    XF Goe222 Il 50000

    Uploaded by

    huy4ngoo
    The document contains the results of an XFOIL analysis calculating the polar performance of the GOE 222 airfoil. It lists the coefficient of lift (CL), drag (CD), pressure drag (CDp), and moment (CM) for Reynolds numbers of 0.05e6 and Mach numbers from -8.25 to 8.25 degrees. The analysis also includes boundary layer transition locations on the airfoil's top and bottom surfaces.

    Copyright:

    © All Rights Reserved
    Download as TXT, PDF, TXT or read online from Scribd
    Download as txt, pdf, or txt
    of 2

    XFOIL Version 6.

    96

    Calculated polar for: GOE 222 (MVA H.33) AIRFOIL

    1 1 Reynolds number fixed Mach number fixed

    xtrf = 1.000 (top) 1.000 (bottom)


    Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000

    alpha CL CD CDp CM Top_Xtr Bot_Xtr


    ------ -------- --------- --------- -------- -------- --------
    -6.500 -0.3964 0.13592 0.13059 -0.0029 1.0000 0.2864
    -6.250 -0.4270 0.13573 0.13050 -0.0010 1.0000 0.2896
    -6.000 -0.4631 0.13557 0.13045 0.0016 1.0000 0.2906
    -5.250 -0.5217 0.09158 0.08633 -0.0380 1.0000 0.1627
    -5.000 -0.4994 0.07961 0.07414 -0.0526 1.0000 0.1582
    -4.750 -0.4015 0.05657 0.04929 -0.0949 1.0000 0.1586
    -4.500 -0.3685 0.05442 0.04678 -0.0996 1.0000 0.1645
    -4.250 -0.3437 0.05381 0.04610 -0.1012 1.0000 0.1710
    -4.000 -0.3117 0.05268 0.04467 -0.1047 1.0000 0.1793
    -3.750 -0.2871 0.05268 0.04464 -0.1061 1.0000 0.1907
    -3.500 -0.2652 0.05311 0.04522 -0.1067 1.0000 0.2056
    -3.250 -0.2444 0.05385 0.04614 -0.1070 1.0000 0.2265
    -2.500 -0.1763 0.06486 0.05791 -0.1009 0.9733 0.3476
    -2.250 -0.1606 0.06792 0.06101 -0.0977 0.9656 0.3717
    -2.000 -0.1350 0.07099 0.06402 -0.0967 0.9544 0.4030
    -1.750 -0.1297 0.07218 0.06526 -0.0928 0.9453 0.4223
    -1.500 -0.1002 0.07514 0.06810 -0.0930 0.9375 0.4563
    -1.250 -0.0954 0.07543 0.06840 -0.0900 0.9292 0.4748
    -1.000 -0.0709 0.07736 0.07030 -0.0889 0.9224 0.4974
    -0.750 -0.0546 0.07740 0.07019 -0.0904 0.9165 0.5122
    -0.500 -0.0253 0.07837 0.07104 -0.0920 0.9077 0.5262
    -0.250 -0.0038 0.07927 0.07185 -0.0932 0.9037 0.5380
    0.000 0.0198 0.07964 0.07206 -0.0959 0.8955 0.5501
    0.250 0.0568 0.08135 0.07361 -0.0997 0.8902 0.5631
    0.500 0.0712 0.08192 0.07411 -0.1002 0.8864 0.5710
    0.750 0.0965 0.08266 0.07473 -0.1025 0.8776 0.5798
    1.000 0.1419 0.08522 0.07711 -0.1080 0.8728 0.5902
    1.250 0.1437 0.08501 0.07687 -0.1067 0.8675 0.5948
    1.500 0.1773 0.08654 0.07822 -0.1111 0.8605 0.6023
    1.750 0.2177 0.08922 0.08080 -0.1152 0.8564 0.6109
    2.000 0.2192 0.08900 0.08053 -0.1140 0.8484 0.6150
    2.250 0.2564 0.09110 0.08250 -0.1183 0.8421 0.6220
    2.500 0.2806 0.09310 0.08444 -0.1202 0.8388 0.6289
    2.750 0.2930 0.09369 0.08497 -0.1206 0.8295 0.6346
    3.000 0.3331 0.09646 0.08764 -0.1248 0.8242 0.6424
    3.250 0.3359 0.09691 0.08808 -0.1236 0.8165 0.6468
    3.500 0.3691 0.09917 0.09024 -0.1271 0.8095 0.6554
    3.750 0.3886 0.10117 0.09221 -0.1281 0.8048 0.6627
    4.000 0.4064 0.10238 0.09336 -0.1290 0.7944 0.6702
    4.250 0.4467 0.10591 0.09681 -0.1330 0.7898 0.6801
    4.500 0.4443 0.10589 0.09680 -0.1310 0.7787 0.6858
    4.750 0.4901 0.10981 0.10066 -0.1355 0.7730 0.6993
    5.000 0.4848 0.10968 0.10054 -0.1331 0.7611 0.7056
    5.250 0.5156 0.11270 0.10355 -0.1355 0.7546 0.7180
    5.500 0.5279 0.11371 0.10458 -0.1354 0.7417 0.7284
    5.750 0.5386 0.11556 0.10647 -0.1353 0.7329 0.7403
    6.000 0.5718 0.11811 0.10907 -0.1375 0.7222 0.7602
    6.250 0.5702 0.11920 0.11027 -0.1358 0.7115 0.7727
    6.500 0.6127 0.12265 0.11389 -0.1387 0.7030 0.8083
    6.750 0.5995 0.12281 0.11423 -0.1357 0.6911 0.8307
    7.000 0.6262 0.12553 0.11707 -0.1373 0.6836 1.0000
    7.250 0.6370 0.12725 0.11873 -0.1384 0.6700 1.0000
    7.500 0.6445 0.12981 0.12122 -0.1393 0.6599 1.0000
    7.750 0.6957 0.13473 0.12594 -0.1443 0.6496 1.0000
    8.000 0.6780 0.13545 0.12667 -0.1424 0.6391 1.0000
    8.250 0.7167 0.14015 0.13122 -0.1457 0.6320 1.0000

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