USA 51 AIRFOIL (usa51-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 51 AIRFOIL (usa51-il) Reynolds number: 1,000,000 Max Cl/Cd: 86.25 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa51-il-1000000-n5.txt Download as CSV file: xf-usa51-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 51 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4748 0.10808 0.10643 -0.0169 1.0000 0.0031
-10.750 -0.4630 0.10720 0.10555 -0.0174 1.0000 0.0035
-10.250 -0.8198 0.02804 0.02536 -0.0487 0.9973 0.0031
-10.000 -0.7994 0.02543 0.02247 -0.0495 0.9955 0.0032
-9.750 -0.7800 0.02282 0.01954 -0.0498 0.9933 0.0033
-9.500 -0.7559 0.02140 0.01794 -0.0502 0.9913 0.0034
-9.250 -0.7304 0.02003 0.01638 -0.0507 0.9895 0.0035
-9.000 -0.7035 0.01854 0.01468 -0.0514 0.9878 0.0035
-8.750 -0.6733 0.01750 0.01348 -0.0526 0.9865 0.0037
-8.500 -0.6480 0.01648 0.01230 -0.0525 0.9832 0.0038
-8.250 -0.6198 0.01555 0.01122 -0.0530 0.9807 0.0040
-8.000 -0.5900 0.01470 0.01019 -0.0538 0.9788 0.0042
-7.750 -0.5583 0.01404 0.00942 -0.0549 0.9774 0.0044
-7.500 -0.5298 0.01342 0.00869 -0.0553 0.9748 0.0046
-7.250 -0.5046 0.01250 0.00763 -0.0550 0.9710 0.0049
-7.000 -0.4748 0.01191 0.00697 -0.0556 0.9676 0.0052
-6.750 -0.4433 0.01141 0.00640 -0.0565 0.9649 0.0055
-6.500 -0.4185 0.01098 0.00591 -0.0559 0.9589 0.0058
-6.250 -0.3886 0.01055 0.00541 -0.0563 0.9538 0.0062
-6.000 -0.3613 0.01018 0.00498 -0.0562 0.9466 0.0066
-5.750 -0.3324 0.00983 0.00453 -0.0564 0.9383 0.0069
-5.500 -0.3059 0.00939 0.00403 -0.0561 0.9259 0.0077
-5.250 -0.2783 0.00912 0.00371 -0.0560 0.9034 0.0088
-5.000 -0.2521 0.00895 0.00334 -0.0554 0.8559 0.0096
-4.750 -0.2284 0.00888 0.00307 -0.0544 0.8148 0.0103
-4.500 -0.2043 0.00871 0.00276 -0.0535 0.7888 0.0122
-4.250 -0.1796 0.00861 0.00260 -0.0528 0.7662 0.0144
-4.000 -0.1543 0.00852 0.00242 -0.0522 0.7466 0.0166
-3.750 -0.1288 0.00841 0.00228 -0.0516 0.7296 0.0208
-3.500 -0.1028 0.00837 0.00216 -0.0512 0.7131 0.0237
-3.250 -0.0766 0.00835 0.00206 -0.0508 0.6963 0.0250
-3.000 -0.0504 0.00834 0.00197 -0.0504 0.6787 0.0256
-2.750 -0.0247 0.00822 0.00176 -0.0499 0.6604 0.0275
-2.500 0.0010 0.00816 0.00159 -0.0493 0.6377 0.0294
-2.250 0.0258 0.00818 0.00147 -0.0486 0.5991 0.0314
-2.000 0.0499 0.00829 0.00137 -0.0479 0.5456 0.0331
-1.750 0.0739 0.00846 0.00130 -0.0471 0.4872 0.0354
-1.500 0.0994 0.00853 0.00124 -0.0466 0.4559 0.0413
-1.250 0.1246 0.00839 0.00118 -0.0461 0.4393 0.0818
-1.000 0.1505 0.00830 0.00113 -0.0457 0.4268 0.1136
-0.750 0.1765 0.00821 0.00110 -0.0454 0.4168 0.1488
-0.500 0.2023 0.00810 0.00108 -0.0450 0.4067 0.1944
-0.250 0.2282 0.00801 0.00106 -0.0446 0.3953 0.2374
0.000 0.2536 0.00788 0.00105 -0.0442 0.3856 0.2973
0.250 0.2786 0.00768 0.00105 -0.0437 0.3771 0.3798
0.500 0.3035 0.00749 0.00106 -0.0432 0.3680 0.4633
0.750 0.3254 0.00710 0.00108 -0.0421 0.3592 0.6128
1.000 0.3464 0.00677 0.00113 -0.0407 0.3508 0.7440
1.250 0.3651 0.00642 0.00122 -0.0384 0.3432 0.8821
1.500 0.4180 0.00645 0.00135 -0.0439 0.3306 0.9566
1.750 0.4608 0.00660 0.00142 -0.0473 0.3157 0.9715
2.000 0.4963 0.00678 0.00151 -0.0490 0.2965 0.9825
2.250 0.5320 0.00698 0.00160 -0.0509 0.2748 0.9897
2.500 0.5652 0.00726 0.00172 -0.0522 0.2422 0.9941
2.750 0.5934 0.00798 0.00201 -0.0528 0.1512 0.9970
3.000 0.6249 0.00830 0.00219 -0.0539 0.1253 0.9991
3.250 0.6539 0.00848 0.00233 -0.0543 0.1176 1.0000
3.500 0.6774 0.00862 0.00245 -0.0534 0.1140 1.0000
3.750 0.7009 0.00877 0.00258 -0.0525 0.1107 1.0000
4.000 0.7245 0.00893 0.00274 -0.0517 0.1074 1.0000
4.250 0.7483 0.00908 0.00288 -0.0509 0.1040 1.0000
4.500 0.7719 0.00925 0.00303 -0.0501 0.0992 1.0000
4.750 0.7953 0.00945 0.00320 -0.0492 0.0943 1.0000
5.000 0.8191 0.00962 0.00337 -0.0484 0.0902 1.0000
5.250 0.8430 0.00978 0.00353 -0.0477 0.0865 1.0000
5.500 0.8659 0.01004 0.00373 -0.0468 0.0762 1.0000
5.750 0.8834 0.01081 0.00427 -0.0450 0.0254 1.0000
6.000 0.9060 0.01111 0.00458 -0.0440 0.0178 1.0000
6.250 0.9284 0.01143 0.00489 -0.0430 0.0132 1.0000
6.500 0.9511 0.01171 0.00521 -0.0421 0.0111 1.0000
6.750 0.9736 0.01201 0.00553 -0.0412 0.0097 1.0000
7.000 0.9953 0.01239 0.00592 -0.0401 0.0081 1.0000
7.250 1.0174 0.01273 0.00630 -0.0391 0.0074 1.0000
7.500 1.0395 0.01306 0.00665 -0.0381 0.0065 1.0000
7.750 1.0611 0.01343 0.00705 -0.0371 0.0061 1.0000
8.000 1.0819 0.01388 0.00752 -0.0359 0.0054 1.0000
8.250 1.1024 0.01435 0.00805 -0.0347 0.0050 1.0000
8.500 1.1232 0.01479 0.00854 -0.0335 0.0048 1.0000
8.750 1.1434 0.01526 0.00909 -0.0323 0.0045 1.0000
9.000 1.1633 0.01574 0.00962 -0.0311 0.0042 1.0000
9.250 1.1840 0.01614 0.01003 -0.0301 0.0038 1.0000
9.500 1.2021 0.01674 0.01068 -0.0286 0.0035 1.0000
9.750 1.2181 0.01749 0.01151 -0.0268 0.0033 1.0000
10.000 1.2356 0.01808 0.01217 -0.0253 0.0032 1.0000
10.250 1.2520 0.01870 0.01287 -0.0236 0.0031 1.0000
10.500 1.2683 0.01930 0.01354 -0.0219 0.0029 1.0000
10.750 1.2809 0.02011 0.01443 -0.0197 0.0029 1.0000
11.000 1.2919 0.02080 0.01521 -0.0172 0.0028 1.0000
11.250 1.2988 0.02166 0.01618 -0.0141 0.0027 1.0000
11.500 1.3081 0.02243 0.01702 -0.0115 0.0026 1.0000
11.750 1.3147 0.02339 0.01806 -0.0087 0.0025 1.0000
12.000 1.3234 0.02426 0.01901 -0.0065 0.0024 1.0000
12.250 1.3254 0.02561 0.02047 -0.0037 0.0024 1.0000
12.500 1.3373 0.02639 0.02129 -0.0023 0.0023 1.0000
12.750 1.3375 0.02803 0.02304 0.0001 0.0023 1.0000
13.000 1.3388 0.02975 0.02487 0.0018 0.0022 1.0000
13.250 1.3355 0.03206 0.02731 0.0033 0.0022 1.0000
13.500 1.3306 0.03481 0.03017 0.0041 0.0021 1.0000
13.750 1.3238 0.03810 0.03361 0.0043 0.0020 1.0000
14.000 1.3205 0.04127 0.03689 0.0039 0.0020 1.0000
14.250 1.3154 0.04492 0.04067 0.0031 0.0020 1.0000
14.500 1.3092 0.04892 0.04480 0.0018 0.0020 1.0000
14.750 1.2998 0.05364 0.04967 0.0001 0.0019 1.0000
15.000 1.2864 0.05907 0.05523 -0.0021 0.0020 1.0000
15.250 1.2822 0.06333 0.05959 -0.0039 0.0020 1.0000
15.500 1.2686 0.06899 0.06539 -0.0062 0.0019 1.0000
15.750 1.2570 0.07437 0.07087 -0.0084 0.0019 1.0000
16.000 1.2385 0.08091 0.07755 -0.0111 0.0019 1.0000
16.250 1.2292 0.08614 0.08289 -0.0134 0.0019 1.0000
16.500 1.2155 0.09218 0.08903 -0.0160 0.0019 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 51 AIRFOIL (usa51-il)