RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Reynolds number: 200,000 Max Cl/Cd: 56.07 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg14a147-il-200000-n5.txt Download as CSV file: xf-rg14a147-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5175 0.09638 0.09282 -0.0139 1.0000 0.0221
-8.750 -0.5163 0.09225 0.08873 -0.0161 1.0000 0.0222
-8.500 -0.5155 0.08824 0.08476 -0.0183 1.0000 0.0222
-8.250 -0.4454 0.07048 0.06721 -0.0200 1.0000 0.0121
-8.000 -0.4516 0.06575 0.06251 -0.0218 1.0000 0.0118
-7.750 -0.5275 0.07293 0.06955 -0.0241 1.0000 0.0123
-7.500 -0.5301 0.06794 0.06454 -0.0271 1.0000 0.0119
-7.250 -0.5298 0.06303 0.05958 -0.0295 1.0000 0.0114
-7.000 -0.5271 0.05806 0.05453 -0.0313 1.0000 0.0112
-6.750 -0.5223 0.05327 0.04960 -0.0323 1.0000 0.0108
-6.500 -0.5158 0.04836 0.04450 -0.0326 1.0000 0.0104
-6.250 -0.5070 0.04361 0.03951 -0.0324 1.0000 0.0101
-6.000 -0.4963 0.03881 0.03440 -0.0317 1.0000 0.0097
-5.750 -0.4836 0.03421 0.02942 -0.0305 1.0000 0.0094
-5.500 -0.4687 0.03001 0.02476 -0.0290 1.0000 0.0091
-5.250 -0.4516 0.02631 0.02057 -0.0274 1.0000 0.0090
-5.000 -0.4322 0.02334 0.01710 -0.0258 1.0000 0.0091
-4.750 -0.4113 0.02101 0.01430 -0.0244 1.0000 0.0093
-4.500 -0.3893 0.01911 0.01204 -0.0231 1.0000 0.0098
-4.250 -0.3669 0.01770 0.01037 -0.0219 1.0000 0.0108
-4.000 -0.3439 0.01729 0.00980 -0.0209 1.0000 0.0128
-3.750 -0.3220 0.01528 0.00760 -0.0196 1.0000 0.0140
-3.500 -0.2999 0.01415 0.00637 -0.0185 1.0000 0.0151
-3.250 -0.2770 0.01337 0.00551 -0.0176 0.9999 0.0169
-3.000 -0.2411 0.01271 0.00472 -0.0193 0.9956 0.0214
-2.750 -0.2063 0.01206 0.00402 -0.0209 0.9908 0.0337
-2.500 -0.1720 0.01126 0.00348 -0.0226 0.9855 0.0995
-2.250 -0.1385 0.01044 0.00321 -0.0245 0.9800 0.2345
-2.000 -0.1086 0.00941 0.00300 -0.0257 0.9730 0.4461
-1.750 -0.0809 0.00853 0.00293 -0.0256 0.9658 0.6540
-1.500 -0.0551 0.00816 0.00303 -0.0243 0.9579 0.8164
-1.250 -0.0253 0.00807 0.00295 -0.0239 0.9492 0.8793
-1.000 0.0128 0.00803 0.00288 -0.0255 0.9427 0.9199
-0.750 0.0538 0.00800 0.00279 -0.0279 0.9340 0.9480
-0.500 0.1002 0.00796 0.00266 -0.0316 0.9264 0.9681
-0.250 0.1473 0.00791 0.00254 -0.0355 0.9166 0.9822
0.000 0.1960 0.00784 0.00241 -0.0398 0.9034 0.9913
0.250 0.2428 0.00777 0.00228 -0.0438 0.8857 0.9997
0.500 0.2753 0.00771 0.00215 -0.0446 0.8613 1.0000
0.750 0.3030 0.00771 0.00206 -0.0444 0.8342 1.0000
1.000 0.3279 0.00775 0.00200 -0.0436 0.8047 1.0000
1.250 0.3516 0.00783 0.00197 -0.0426 0.7737 1.0000
1.500 0.3748 0.00796 0.00199 -0.0414 0.7449 1.0000
1.750 0.3975 0.00811 0.00204 -0.0403 0.7158 1.0000
2.000 0.4199 0.00826 0.00210 -0.0390 0.6812 1.0000
2.500 0.4643 0.00868 0.00225 -0.0365 0.6083 1.0000
2.750 0.4867 0.00894 0.00241 -0.0353 0.5687 1.0000
3.000 0.5091 0.00923 0.00256 -0.0342 0.5271 1.0000
3.250 0.5315 0.00956 0.00273 -0.0331 0.4831 1.0000
3.500 0.5542 0.00991 0.00294 -0.0321 0.4376 1.0000
3.750 0.5770 0.01029 0.00318 -0.0312 0.3916 1.0000
4.000 0.5998 0.01072 0.00350 -0.0304 0.3439 1.0000
4.250 0.6218 0.01127 0.00382 -0.0295 0.2821 1.0000
4.500 0.6437 0.01190 0.00420 -0.0287 0.2180 1.0000
4.750 0.6664 0.01250 0.00463 -0.0279 0.1694 1.0000
5.000 0.6890 0.01317 0.00511 -0.0272 0.1224 1.0000
5.250 0.7113 0.01389 0.00568 -0.0265 0.0823 1.0000
5.500 0.7336 0.01467 0.00640 -0.0257 0.0533 1.0000
5.750 0.7558 0.01549 0.00715 -0.0249 0.0333 1.0000
6.000 0.7777 0.01639 0.00808 -0.0240 0.0214 1.0000
6.250 0.7992 0.01742 0.00928 -0.0229 0.0162 1.0000
6.500 0.8197 0.01855 0.01047 -0.0218 0.0120 1.0000
6.750 0.8412 0.01956 0.01165 -0.0208 0.0102 1.0000
7.000 0.8614 0.02086 0.01311 -0.0196 0.0092 1.0000
7.250 0.8813 0.02228 0.01470 -0.0184 0.0084 1.0000
7.500 0.9008 0.02388 0.01648 -0.0172 0.0080 1.0000
7.750 0.9190 0.02587 0.01869 -0.0159 0.0076 1.0000
8.000 0.9356 0.02839 0.02159 -0.0145 0.0073 1.0000
8.250 0.9491 0.03154 0.02514 -0.0128 0.0071 1.0000
8.500 0.9596 0.03491 0.02896 -0.0110 0.0070 1.0000
8.750 0.9658 0.03859 0.03309 -0.0090 0.0070 1.0000
9.000 0.9676 0.04240 0.03734 -0.0069 0.0070 1.0000
9.250 0.9643 0.04632 0.04164 -0.0047 0.0070 1.0000
9.500 0.9566 0.05007 0.04573 -0.0026 0.0069 1.0000
9.750 0.9435 0.05325 0.04913 -0.0002 0.0069 1.0000
10.000 0.9283 0.05684 0.05293 0.0010 0.0070 1.0000
10.250 0.9083 0.06158 0.05783 -0.0002 0.0069 1.0000
10.500 0.8927 0.06693 0.06338 -0.0037 0.0069 1.0000
10.750 0.8817 0.07284 0.06943 -0.0082 0.0070 1.0000
11.000 0.8701 0.08087 0.07760 -0.0148 0.0072 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 14A-1.4/7.0 AIRFOIL (rg14a147-il)