USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 46 AIRFOIL (usa46-il) Reynolds number: 500,000 Max Cl/Cd: 75.06 at α=2.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa46-il-500000-n5.txt Download as CSV file: xf-usa46-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 46 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5701 0.08426 0.08207 -0.0129 1.0000 0.0047
-9.250 -0.5984 0.07471 0.07260 -0.0179 1.0000 0.0046
-9.000 -0.6240 0.06670 0.06465 -0.0231 1.0000 0.0045
-8.500 -0.6607 0.02898 0.02591 -0.0571 1.0000 0.0041
-8.250 -0.6450 0.02489 0.02133 -0.0582 1.0000 0.0042
-8.000 -0.6260 0.02209 0.01812 -0.0585 1.0000 0.0044
-7.750 -0.6049 0.02002 0.01572 -0.0583 1.0000 0.0046
-7.500 -0.5828 0.01840 0.01380 -0.0580 1.0000 0.0048
-7.250 -0.5600 0.01709 0.01224 -0.0576 1.0000 0.0050
-7.000 -0.5377 0.01562 0.01058 -0.0572 1.0000 0.0055
-6.750 -0.5140 0.01487 0.00973 -0.0567 1.0000 0.0060
-6.500 -0.4901 0.01407 0.00879 -0.0562 1.0000 0.0065
-6.250 -0.4662 0.01322 0.00774 -0.0555 1.0000 0.0070
-6.000 -0.4421 0.01252 0.00689 -0.0549 1.0000 0.0075
-5.750 -0.4180 0.01178 0.00600 -0.0542 1.0000 0.0081
-5.500 -0.3902 0.01106 0.00515 -0.0543 0.9990 0.0097
-5.250 -0.3557 0.01057 0.00459 -0.0559 0.9957 0.0119
-5.000 -0.3228 0.01005 0.00397 -0.0570 0.9921 0.0166
-4.750 -0.2904 0.00963 0.00358 -0.0581 0.9884 0.0275
-4.500 -0.2584 0.00937 0.00332 -0.0590 0.9843 0.0368
-4.250 -0.2266 0.00916 0.00308 -0.0599 0.9792 0.0423
-4.000 -0.1948 0.00890 0.00282 -0.0607 0.9736 0.0493
-3.750 -0.1621 0.00870 0.00257 -0.0618 0.9677 0.0550
-3.500 -0.1299 0.00845 0.00230 -0.0627 0.9608 0.0609
-3.250 -0.0962 0.00823 0.00208 -0.0639 0.9536 0.0691
-3.000 -0.0625 0.00799 0.00184 -0.0651 0.9435 0.0816
-2.750 -0.0267 0.00775 0.00166 -0.0668 0.9293 0.1020
-2.500 0.0075 0.00761 0.00152 -0.0680 0.9105 0.1228
-2.250 0.0372 0.00755 0.00140 -0.0683 0.8898 0.1358
-2.000 0.0652 0.00751 0.00129 -0.0681 0.8695 0.1451
-1.750 0.0922 0.00750 0.00120 -0.0677 0.8480 0.1529
-1.500 0.1189 0.00747 0.00112 -0.0672 0.8265 0.1615
-1.250 0.1454 0.00746 0.00105 -0.0668 0.8035 0.1720
-1.000 0.1716 0.00745 0.00098 -0.0663 0.7780 0.1854
-0.750 0.1977 0.00744 0.00093 -0.0658 0.7501 0.2019
-0.500 0.2236 0.00739 0.00088 -0.0653 0.7210 0.2337
-0.250 0.2495 0.00727 0.00086 -0.0649 0.6941 0.2969
0.000 0.2752 0.00708 0.00086 -0.0646 0.6694 0.3948
0.250 0.3008 0.00690 0.00088 -0.0642 0.6450 0.4962
0.500 0.3261 0.00676 0.00093 -0.0636 0.6198 0.5888
0.750 0.3507 0.00668 0.00103 -0.0628 0.5925 0.6773
1.000 0.3757 0.00672 0.00112 -0.0621 0.5661 0.7380
1.250 0.4013 0.00678 0.00119 -0.0615 0.5440 0.7736
1.500 0.4268 0.00680 0.00125 -0.0609 0.5270 0.8073
1.750 0.4518 0.00674 0.00131 -0.0600 0.5107 0.8461
2.000 0.4904 0.00664 0.00137 -0.0623 0.4878 1.0000
2.250 0.5164 0.00688 0.00147 -0.0619 0.4473 1.0000
2.500 0.5415 0.00726 0.00159 -0.0614 0.3837 1.0000
2.750 0.5667 0.00765 0.00176 -0.0609 0.3346 1.0000
3.000 0.5921 0.00804 0.00196 -0.0605 0.2931 1.0000
3.250 0.6180 0.00835 0.00216 -0.0602 0.2607 1.0000
3.500 0.6436 0.00872 0.00236 -0.0598 0.2179 1.0000
3.750 0.6667 0.00950 0.00270 -0.0592 0.1236 1.0000
4.000 0.6897 0.01035 0.00319 -0.0586 0.0469 1.0000
4.250 0.7145 0.01090 0.00362 -0.0580 0.0181 1.0000
4.500 0.7402 0.01133 0.00409 -0.0575 0.0123 1.0000
4.750 0.7659 0.01173 0.00458 -0.0570 0.0109 1.0000
5.000 0.7914 0.01218 0.00511 -0.0565 0.0098 1.0000
5.250 0.8166 0.01263 0.00561 -0.0560 0.0084 1.0000
5.500 0.8405 0.01332 0.00638 -0.0553 0.0073 1.0000
5.750 0.8637 0.01417 0.00733 -0.0545 0.0068 1.0000
6.000 0.8874 0.01487 0.00812 -0.0537 0.0064 1.0000
6.250 0.9103 0.01572 0.00908 -0.0528 0.0060 1.0000
6.500 0.9327 0.01669 0.01015 -0.0518 0.0057 1.0000
6.750 0.9555 0.01754 0.01108 -0.0510 0.0052 1.0000
7.000 0.9789 0.01816 0.01173 -0.0505 0.0048 1.0000
7.250 0.9991 0.01949 0.01315 -0.0494 0.0044 1.0000
7.500 1.0201 0.02083 0.01470 -0.0483 0.0042 1.0000
7.750 1.0410 0.02223 0.01630 -0.0472 0.0040 1.0000
8.000 1.0604 0.02407 0.01838 -0.0459 0.0038 1.0000
8.250 1.0783 0.02632 0.02094 -0.0444 0.0036 1.0000
8.500 1.0938 0.02897 0.02392 -0.0427 0.0035 1.0000
8.750 1.1061 0.03211 0.02744 -0.0408 0.0034 1.0000
9.000 1.1143 0.03568 0.03141 -0.0387 0.0033 1.0000
9.250 1.1168 0.03984 0.03598 -0.0363 0.0033 1.0000
9.500 1.1127 0.04445 0.04098 -0.0336 0.0033 1.0000
9.750 1.1024 0.04907 0.04592 -0.0310 0.0033 1.0000
10.000 1.0862 0.05283 0.04991 -0.0279 0.0034 1.0000
10.250 1.0646 0.05689 0.05417 -0.0260 0.0034 1.0000
10.500 1.0451 0.06142 0.05887 -0.0265 0.0034 1.0000
10.750 1.0245 0.06726 0.06486 -0.0295 0.0035 1.0000
11.000 1.0057 0.07438 0.07212 -0.0349 0.0035 1.0000
11.250 0.9849 0.08392 0.08179 -0.0425 0.0035 1.0000
11.500 0.9634 0.09630 0.09428 -0.0515 0.0036 1.0000
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Polar data table (+)
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