RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Reynolds number: 50,000 Max Cl/Cd: 33.93 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg14a147-il-50000.txt Download as CSV file: xf-rg14a147-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RG 14A-1.4/7.0 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5167 0.10089 0.09417 0.0006 1.0000 0.2364
-8.000 -0.5186 0.09788 0.09125 0.0001 1.0000 0.2477
-7.750 -0.5191 0.09481 0.08826 0.0001 1.0000 0.2607
-7.500 -0.5130 0.09127 0.08478 0.0010 1.0000 0.2757
-7.250 -0.5080 0.08799 0.08156 0.0019 1.0000 0.2916
-7.000 -0.5144 0.08561 0.07929 0.0028 1.0000 0.3111
-6.750 -0.5022 0.08195 0.07559 0.0049 1.0000 0.3338
-6.500 -0.4990 0.07913 0.07285 0.0074 1.0000 0.3637
-6.250 -0.5072 0.07711 0.07095 0.0093 1.0000 0.3949
-6.000 -0.4860 0.07330 0.06715 0.0137 1.0000 0.4338
-5.250 -0.4663 0.04890 0.04153 -0.0296 1.0000 0.1454
-5.000 -0.4440 0.04354 0.03531 -0.0308 1.0000 0.1209
-4.750 -0.4241 0.03969 0.03113 -0.0302 1.0000 0.1167
-4.500 -0.3998 0.03606 0.02656 -0.0295 1.0000 0.1080
-4.250 -0.3768 0.03273 0.02291 -0.0286 1.0000 0.1054
-4.000 -0.3518 0.02982 0.01952 -0.0275 1.0000 0.1045
-3.750 -0.3257 0.02733 0.01652 -0.0262 1.0000 0.1066
-3.500 -0.3012 0.02533 0.01434 -0.0251 1.0000 0.1195
-3.250 -0.2751 0.02331 0.01216 -0.0237 1.0000 0.1344
-3.000 -0.1566 0.01612 0.00842 -0.0316 1.0000 1.0000
-2.750 -0.1443 0.01583 0.00766 -0.0294 1.0000 1.0000
-2.500 -0.1314 0.01561 0.00706 -0.0273 1.0000 1.0000
-2.250 -0.1177 0.01544 0.00656 -0.0251 1.0000 1.0000
-2.000 -0.1035 0.01532 0.00614 -0.0231 1.0000 1.0000
-1.750 -0.0888 0.01523 0.00579 -0.0210 1.0000 1.0000
-1.500 -0.0733 0.01519 0.00550 -0.0191 1.0000 1.0000
-1.250 -0.0569 0.01518 0.00522 -0.0174 1.0000 1.0000
-1.000 -0.0397 0.01520 0.00505 -0.0158 1.0000 1.0000
-0.750 -0.0219 0.01525 0.00493 -0.0144 1.0000 1.0000
-0.500 -0.0037 0.01533 0.00486 -0.0131 1.0000 1.0000
-0.250 0.0148 0.01545 0.00485 -0.0119 1.0000 1.0000
0.000 0.0334 0.01560 0.00486 -0.0108 1.0000 1.0000
0.250 0.0520 0.01579 0.00495 -0.0097 1.0000 1.0000
0.500 0.0705 0.01602 0.00511 -0.0087 1.0000 1.0000
0.750 0.0889 0.01629 0.00533 -0.0078 1.0000 1.0000
1.000 0.1071 0.01661 0.00561 -0.0070 1.0000 1.0000
1.250 0.1249 0.01698 0.00597 -0.0063 1.0000 1.0000
1.500 0.1424 0.01741 0.00640 -0.0056 1.0000 1.0000
1.750 0.1594 0.01790 0.00690 -0.0051 1.0000 1.0000
2.000 0.1759 0.01847 0.00750 -0.0047 1.0000 1.0000
2.250 0.1918 0.01913 0.00820 -0.0044 1.0000 1.0000
2.500 0.2071 0.01988 0.00900 -0.0042 1.0000 1.0000
2.750 0.2477 0.02103 0.01035 -0.0091 0.9870 1.0000
3.000 0.3257 0.02232 0.01194 -0.0201 0.9558 1.0000
3.250 0.3945 0.02297 0.01294 -0.0285 0.9204 1.0000
3.500 0.4701 0.02305 0.01358 -0.0366 0.8828 1.0000
3.750 0.5442 0.02243 0.01349 -0.0426 0.8418 1.0000
4.000 0.6003 0.02154 0.01304 -0.0440 0.7960 1.0000
4.250 0.6378 0.02086 0.01271 -0.0418 0.7445 1.0000
4.500 0.6643 0.02052 0.01247 -0.0379 0.6857 1.0000
4.750 0.6852 0.02042 0.01232 -0.0334 0.6176 1.0000
5.000 0.7020 0.02069 0.01234 -0.0288 0.5349 1.0000
5.250 0.7144 0.02146 0.01265 -0.0241 0.4317 1.0000
5.500 0.7237 0.02307 0.01344 -0.0199 0.3101 1.0000
5.750 0.7354 0.02550 0.01510 -0.0169 0.2059 1.0000
6.000 0.7565 0.02820 0.01743 -0.0152 0.1516 1.0000
6.250 0.7817 0.03093 0.02008 -0.0141 0.1264 1.0000
6.500 0.8091 0.03417 0.02345 -0.0133 0.1149 1.0000
6.750 0.8312 0.03733 0.02732 -0.0116 0.1078 1.0000
7.000 0.8502 0.04048 0.03074 -0.0104 0.0988 1.0000
7.250 0.8663 0.04446 0.03538 -0.0087 0.0968 1.0000
7.500 0.8789 0.04893 0.04043 -0.0072 0.0972 1.0000
7.750 0.8877 0.05376 0.04577 -0.0058 0.0987 1.0000
8.000 0.8967 0.05894 0.05125 -0.0048 0.1007 1.0000
8.250 0.8837 0.06481 0.05792 -0.0042 0.1069 1.0000
8.500 0.8790 0.07066 0.06402 -0.0043 0.1118 1.0000
8.750 0.8159 0.06787 0.06175 -0.0005 0.1151 1.0000
9.000 0.7650 0.07347 0.06752 -0.0019 0.1176 1.0000
9.250 0.7248 0.08152 0.07554 -0.0075 0.1220 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 14A-1.4/7.0 AIRFOIL (rg14a147-il)