USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 49 AIRFOIL (usa49-il) Reynolds number: 500,000 Max Cl/Cd: 59.67 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa49-il-500000.txt Download as CSV file: xf-usa49-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: USA 49 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6120 0.12516 0.12287 0.0222 1.0000 0.0117
-10.250 -0.6087 0.12126 0.11898 0.0206 1.0000 0.0120
-7.000 -0.5752 0.05088 0.04828 -0.0267 1.0000 0.0161
-6.750 -0.5646 0.04804 0.04536 -0.0273 1.0000 0.0167
-6.500 -0.5502 0.04523 0.04244 -0.0278 1.0000 0.0174
-6.250 -0.5345 0.04163 0.03867 -0.0282 1.0000 0.0181
-6.000 -0.5178 0.03781 0.03462 -0.0282 1.0000 0.0190
-5.750 -0.4998 0.03397 0.03050 -0.0277 1.0000 0.0202
-5.500 -0.4774 0.03159 0.02781 -0.0266 1.0000 0.0224
-4.250 -0.3812 0.01561 0.00976 -0.0183 1.0000 0.0162
-4.000 -0.3586 0.01420 0.00819 -0.0171 1.0000 0.0163
-3.750 -0.3362 0.01294 0.00686 -0.0161 1.0000 0.0174
-3.500 -0.3132 0.01222 0.00611 -0.0151 1.0000 0.0187
-3.250 -0.2895 0.01136 0.00515 -0.0142 1.0000 0.0188
-3.000 -0.2612 0.01062 0.00436 -0.0143 0.9992 0.0198
-2.750 -0.2248 0.00999 0.00366 -0.0161 0.9966 0.0215
-2.500 -0.1875 0.00957 0.00320 -0.0181 0.9942 0.0234
-2.250 -0.1518 0.00899 0.00253 -0.0197 0.9904 0.0303
-2.000 -0.1151 0.00823 0.00219 -0.0218 0.9869 0.1306
-1.750 -0.0754 0.00741 0.00198 -0.0250 0.9836 0.2891
-1.500 -0.0439 0.00612 0.00182 -0.0264 0.9763 0.6023
-1.000 0.0074 0.00515 0.00187 -0.0241 0.9590 0.9157
-0.750 0.0376 0.00512 0.00185 -0.0239 0.9488 0.9521
-0.500 0.0725 0.00512 0.00179 -0.0251 0.9381 0.9723
-0.250 0.1085 0.00511 0.00173 -0.0265 0.9251 0.9858
0.000 0.1486 0.00510 0.00164 -0.0289 0.9026 0.9969
0.250 0.1764 0.00510 0.00150 -0.0285 0.8647 1.0000
0.500 0.1996 0.00515 0.00137 -0.0272 0.8202 1.0000
0.750 0.2235 0.00530 0.00127 -0.0261 0.7667 1.0000
1.000 0.2481 0.00549 0.00123 -0.0253 0.7167 1.0000
1.250 0.2729 0.00571 0.00122 -0.0247 0.6672 1.0000
1.500 0.2977 0.00596 0.00124 -0.0241 0.6160 1.0000
1.750 0.3229 0.00620 0.00128 -0.0236 0.5717 1.0000
2.000 0.3484 0.00642 0.00133 -0.0231 0.5357 1.0000
2.250 0.3742 0.00662 0.00140 -0.0228 0.5006 1.0000
2.500 0.4000 0.00685 0.00150 -0.0224 0.4599 1.0000
2.750 0.4253 0.00722 0.00159 -0.0221 0.3905 1.0000
3.000 0.4500 0.00777 0.00175 -0.0218 0.3047 1.0000
3.250 0.4739 0.00854 0.00201 -0.0214 0.1904 1.0000
3.500 0.4974 0.00951 0.00244 -0.0211 0.0753 1.0000
3.750 0.5238 0.00984 0.00277 -0.0208 0.0668 1.0000
4.000 0.5503 0.01013 0.00310 -0.0206 0.0636 1.0000
4.250 0.5766 0.01048 0.00348 -0.0204 0.0608 1.0000
4.500 0.6026 0.01089 0.00392 -0.0201 0.0586 1.0000
4.750 0.6282 0.01141 0.00450 -0.0198 0.0566 1.0000
5.000 0.6530 0.01208 0.00524 -0.0193 0.0548 1.0000
5.250 0.6796 0.01229 0.00549 -0.0192 0.0528 1.0000
5.500 0.7058 0.01262 0.00586 -0.0190 0.0495 1.0000
5.750 0.7313 0.01306 0.00635 -0.0187 0.0461 1.0000
6.000 0.7532 0.01431 0.00764 -0.0180 0.0419 1.0000
6.250 0.7815 0.01411 0.00751 -0.0181 0.0393 1.0000
6.500 0.8087 0.01416 0.00760 -0.0181 0.0348 1.0000
6.750 0.8338 0.01465 0.00811 -0.0178 0.0295 1.0000
7.000 0.8623 0.01445 0.00790 -0.0179 0.0242 1.0000
7.250 0.8866 0.01512 0.00862 -0.0175 0.0202 1.0000
7.500 0.9119 0.01558 0.00912 -0.0171 0.0177 1.0000
7.750 0.9320 0.01707 0.01072 -0.0160 0.0154 1.0000
8.000 0.9540 0.01821 0.01202 -0.0152 0.0145 1.0000
8.250 0.9767 0.01918 0.01312 -0.0145 0.0134 1.0000
8.500 0.9998 0.01993 0.01398 -0.0140 0.0122 1.0000
8.750 1.0213 0.02103 0.01520 -0.0132 0.0114 1.0000
9.000 1.0392 0.02297 0.01735 -0.0121 0.0108 1.0000
9.250 1.0489 0.02709 0.02195 -0.0100 0.0103 1.0000
9.500 1.0599 0.03047 0.02575 -0.0083 0.0100 1.0000
9.750 1.0705 0.03343 0.02909 -0.0067 0.0098 1.0000
10.000 1.0723 0.03768 0.03380 -0.0045 0.0097 1.0000
10.250 1.0631 0.04303 0.03961 -0.0021 0.0097 1.0000
10.500 1.0447 0.04848 0.04544 0.0002 0.0098 1.0000
10.750 1.0211 0.05305 0.05025 0.0022 0.0099 1.0000
11.000 0.9970 0.05847 0.05587 0.0009 0.0100 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 49 AIRFOIL (usa49-il)