USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 100,000 Max Cl/Cd: 51.06 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa50-il-100000.txt Download as CSV file: xf-usa50-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: USA 50 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4091 0.10013 0.09540 -0.0205 1.0000 0.0768
-9.000 -0.4215 0.09742 0.09277 -0.0235 1.0000 0.0775
-8.750 -0.4333 0.09451 0.08994 -0.0262 1.0000 0.0778
-8.500 -0.4088 0.08777 0.08317 -0.0215 1.0000 0.0808
-8.250 -0.4052 0.08410 0.07953 -0.0210 1.0000 0.0837
-8.000 -0.4081 0.08050 0.07597 -0.0216 1.0000 0.0865
-7.750 -0.4167 0.07708 0.07263 -0.0230 1.0000 0.0889
-7.500 -0.5073 0.08441 0.07978 -0.0221 1.0000 0.0809
-7.250 -0.5034 0.08106 0.07646 -0.0219 1.0000 0.0836
-7.000 -0.5025 0.07739 0.07281 -0.0238 1.0000 0.0869
-6.750 -0.5084 0.07392 0.06911 -0.0318 1.0000 0.0912
-6.500 -0.5023 0.06877 0.06405 -0.0306 1.0000 0.0931
-6.250 -0.4930 0.06548 0.06084 -0.0282 1.0000 0.0966
-6.000 -0.4907 0.06278 0.05761 -0.0326 1.0000 0.1054
-5.750 -0.4791 0.05773 0.05287 -0.0299 1.0000 0.1084
-5.500 -0.4705 0.05478 0.04955 -0.0311 1.0000 0.1196
-5.250 -0.4582 0.05109 0.04599 -0.0290 1.0000 0.1240
-5.000 -0.4469 0.04783 0.04257 -0.0286 1.0000 0.1362
-4.750 -0.4344 0.04491 0.03954 -0.0276 1.0000 0.1502
-4.500 -0.4211 0.04224 0.03680 -0.0263 1.0000 0.1677
-4.000 -0.3645 0.03141 0.02415 -0.0248 1.0000 0.0737
-3.750 -0.3435 0.02831 0.02056 -0.0231 1.0000 0.0701
-3.500 -0.3187 0.02588 0.01736 -0.0209 1.0000 0.0632
-3.250 -0.2956 0.02348 0.01459 -0.0194 1.0000 0.0625
-3.000 -0.2712 0.02171 0.01240 -0.0179 1.0000 0.0638
-2.750 -0.2476 0.01972 0.01035 -0.0169 1.0000 0.0698
-2.500 -0.2240 0.01838 0.00887 -0.0157 1.0000 0.0820
-2.250 -0.2009 0.01686 0.00741 -0.0143 1.0000 0.1006
-2.000 -0.1805 0.01561 0.00636 -0.0128 1.0000 0.1470
-1.750 -0.0908 0.01175 0.00547 -0.0242 1.0000 1.0000
-1.500 -0.0719 0.01178 0.00520 -0.0225 1.0000 1.0000
-1.250 -0.0529 0.01185 0.00497 -0.0209 1.0000 1.0000
-1.000 -0.0339 0.01193 0.00486 -0.0194 1.0000 1.0000
-0.750 -0.0149 0.01205 0.00480 -0.0180 1.0000 1.0000
-0.500 0.0039 0.01219 0.00481 -0.0166 1.0000 1.0000
-0.250 0.0225 0.01237 0.00487 -0.0153 1.0000 1.0000
0.000 0.0405 0.01260 0.00497 -0.0140 1.0000 1.0000
0.250 0.0578 0.01288 0.00518 -0.0127 1.0000 1.0000
0.500 0.0740 0.01324 0.00548 -0.0114 1.0000 1.0000
0.750 0.1020 0.01373 0.00593 -0.0128 0.9960 1.0000
1.000 0.1628 0.01419 0.00637 -0.0203 0.9813 1.0000
1.250 0.2236 0.01431 0.00650 -0.0274 0.9625 1.0000
1.500 0.2843 0.01423 0.00648 -0.0341 0.9451 1.0000
1.750 0.3366 0.01404 0.00640 -0.0389 0.9261 1.0000
2.000 0.3916 0.01372 0.00620 -0.0441 0.9091 1.0000
2.250 0.4490 0.01332 0.00595 -0.0495 0.8919 1.0000
2.500 0.4966 0.01304 0.00581 -0.0529 0.8680 1.0000
2.750 0.5411 0.01283 0.00578 -0.0556 0.8411 1.0000
3.000 0.5804 0.01275 0.00580 -0.0572 0.8117 1.0000
3.250 0.6135 0.01277 0.00587 -0.0574 0.7785 1.0000
3.500 0.6420 0.01290 0.00599 -0.0566 0.7419 1.0000
3.750 0.6625 0.01307 0.00604 -0.0537 0.6854 1.0000
4.000 0.6816 0.01335 0.00625 -0.0511 0.6358 1.0000
4.250 0.6984 0.01369 0.00639 -0.0479 0.5763 1.0000
4.500 0.7117 0.01418 0.00659 -0.0443 0.4952 1.0000
4.750 0.7187 0.01511 0.00676 -0.0397 0.3209 1.0000
5.000 0.7165 0.01859 0.00849 -0.0348 0.0769 1.0000
5.250 0.7315 0.02008 0.01004 -0.0320 0.0627 1.0000
5.500 0.7478 0.02141 0.01145 -0.0297 0.0534 1.0000
5.750 0.7662 0.02284 0.01295 -0.0276 0.0481 1.0000
6.000 0.7879 0.02457 0.01473 -0.0260 0.0457 1.0000
6.250 0.8131 0.02667 0.01692 -0.0249 0.0446 1.0000
6.500 0.8398 0.02924 0.01970 -0.0239 0.0445 1.0000
6.750 0.8642 0.03225 0.02314 -0.0224 0.0456 1.0000
7.000 0.8843 0.03551 0.02687 -0.0203 0.0471 1.0000
7.250 0.9006 0.03873 0.03055 -0.0181 0.0480 1.0000
7.500 0.9131 0.04190 0.03419 -0.0155 0.0476 1.0000
7.750 0.9235 0.04609 0.03876 -0.0131 0.0497 1.0000
8.000 0.9287 0.05118 0.04464 -0.0094 0.0608 1.0000
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