BOEING 737 ROOT AIRFOIL (b737a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 737 ROOT AIRFOIL (b737a-il) Reynolds number: 50,000 Max Cl/Cd: 16.38 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-b737a-il-50000.txt Download as CSV file: xf-b737a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 737 ROOT 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4239 0.10709 0.09862 0.0359 1.0000 0.4953
-9.000 -0.3949 0.10324 0.09476 0.0351 1.0000 0.5003
-8.750 -0.3854 0.10121 0.09277 0.0353 1.0000 0.5101
-8.500 -0.3903 0.09936 0.09097 0.0359 1.0000 0.5176
-8.250 -0.5792 0.07317 0.06478 0.0078 1.0000 0.3399
-8.000 -0.5603 0.07007 0.06167 0.0075 1.0000 0.3357
-7.750 -0.6756 0.05141 0.04238 -0.0037 1.0000 0.3174
-7.500 -0.6694 0.04788 0.03868 -0.0037 1.0000 0.3173
-7.250 -0.6588 0.04491 0.03558 -0.0034 1.0000 0.3181
-7.000 -0.6423 0.04266 0.03330 -0.0028 1.0000 0.3198
-6.750 -0.6252 0.04071 0.03133 -0.0020 1.0000 0.3220
-6.500 -0.6092 0.03883 0.02941 -0.0012 1.0000 0.3248
-6.250 -0.5943 0.03697 0.02747 -0.0002 1.0000 0.3282
-6.000 -0.5806 0.03512 0.02547 0.0008 1.0000 0.3325
-5.750 -0.5684 0.03326 0.02335 0.0019 1.0000 0.3379
-5.500 -0.5513 0.03187 0.02204 0.0031 1.0000 0.3436
-5.250 -0.5348 0.03077 0.02100 0.0046 1.0000 0.3514
-5.000 -0.5212 0.02945 0.01949 0.0059 1.0000 0.3620
-4.750 -0.5038 0.02861 0.01885 0.0074 1.0000 0.3724
-4.500 -0.4877 0.02765 0.01785 0.0086 1.0000 0.3860
-4.250 -0.4702 0.02702 0.01729 0.0099 1.0000 0.4003
-4.000 -0.4521 0.02644 0.01685 0.0110 1.0000 0.4141
-3.750 -0.4335 0.02591 0.01631 0.0119 1.0000 0.4294
-3.500 -0.4144 0.02555 0.01607 0.0128 1.0000 0.4440
-3.250 -0.3947 0.02523 0.01590 0.0136 1.0000 0.4584
-3.000 -0.3746 0.02499 0.01577 0.0142 1.0000 0.4756
-2.750 -0.3542 0.02484 0.01571 0.0146 1.0000 0.4939
-2.500 -0.3345 0.02478 0.01587 0.0154 1.0000 0.5109
-2.250 -0.2992 0.02488 0.01618 0.0132 0.9929 0.5341
-2.000 -0.2233 0.02518 0.01689 0.0050 0.9618 0.5705
-1.750 -0.1532 0.02531 0.01742 -0.0008 0.9281 0.6152
-1.500 -0.0828 0.02536 0.01786 -0.0046 0.8930 0.6746
-1.250 -0.0302 0.02532 0.01809 -0.0034 0.8540 0.7334
-1.000 0.0036 0.02524 0.01812 0.0015 0.8064 0.7842
-0.750 0.0319 0.02514 0.01794 0.0077 0.7438 0.8266
-0.500 0.0616 0.02503 0.01750 0.0131 0.6607 0.8638
-0.250 0.1119 0.02523 0.01710 0.0135 0.5709 0.8954
0.000 0.1730 0.02566 0.01698 0.0096 0.5101 0.9258
0.250 0.2441 0.02608 0.01696 0.0027 0.4676 0.9521
0.500 0.3099 0.02646 0.01702 -0.0042 0.4387 0.9780
0.750 0.3808 0.02678 0.01707 -0.0126 0.4147 1.0000
1.000 0.4023 0.02726 0.01753 -0.0126 0.4032 1.0000
1.250 0.4234 0.02773 0.01787 -0.0122 0.3925 1.0000
1.500 0.4442 0.02841 0.01857 -0.0120 0.3828 1.0000
1.750 0.4642 0.02908 0.01925 -0.0116 0.3737 1.0000
2.000 0.4838 0.02990 0.01993 -0.0109 0.3660 1.0000
2.250 0.5015 0.03084 0.02109 -0.0103 0.3576 1.0000
2.500 0.5193 0.03171 0.02194 -0.0094 0.3504 1.0000
2.750 0.5364 0.03281 0.02294 -0.0084 0.3441 1.0000
3.000 0.5487 0.03413 0.02457 -0.0073 0.3378 1.0000
3.250 0.5610 0.03539 0.02595 -0.0059 0.3320 1.0000
3.500 0.5748 0.03647 0.02696 -0.0043 0.3267 1.0000
3.750 0.5857 0.03802 0.02849 -0.0026 0.3223 1.0000
4.000 0.5872 0.04002 0.03083 -0.0005 0.3188 1.0000
4.250 0.5887 0.04228 0.03330 0.0014 0.3154 1.0000
4.500 0.5923 0.04480 0.03598 0.0026 0.3120 1.0000
4.750 0.5977 0.04769 0.03900 0.0033 0.3093 1.0000
5.000 0.5986 0.05154 0.04302 0.0031 0.3083 1.0000
5.250 0.5692 0.05941 0.05118 0.0001 0.3125 1.0000
5.500 0.5488 0.06691 0.05876 -0.0037 0.3185 1.0000
5.750 0.5487 0.07250 0.06437 -0.0064 0.3249 1.0000
6.000 0.3749 0.09774 0.08974 -0.0348 0.5024 1.0000
6.250 0.3691 0.09959 0.09152 -0.0342 0.4903 1.0000
6.500 0.3961 0.10305 0.09495 -0.0348 0.4803 1.0000
6.750 0.3906 0.10491 0.09675 -0.0343 0.4682 1.0000
7.000 0.4201 0.10858 0.10040 -0.0348 0.4580 1.0000
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Polar data table (+)
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