RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Reynolds number: 500,000 Max Cl/Cd: 78.93 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rg14a147-il-500000.txt Download as CSV file: xf-rg14a147-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4457 0.11108 0.10887 -0.0075 1.0000 0.0120
-10.500 -0.4442 0.10703 0.10482 -0.0085 1.0000 0.0122
-9.500 -0.5268 0.09909 0.09678 -0.0103 1.0000 0.0119
-9.250 -0.5291 0.09452 0.09223 -0.0119 1.0000 0.0122
-9.000 -0.5281 0.09081 0.08854 -0.0133 1.0000 0.0124
-8.750 -0.5276 0.08690 0.08465 -0.0151 1.0000 0.0124
-8.500 -0.5277 0.08325 0.08103 -0.0167 1.0000 0.0126
-8.250 -0.5291 0.07958 0.07739 -0.0185 1.0000 0.0128
-8.000 -0.5351 0.07558 0.07344 -0.0209 1.0000 0.0128
-7.750 -0.5389 0.07083 0.06869 -0.0250 1.0000 0.0129
-7.500 -0.5380 0.06632 0.06415 -0.0279 1.0000 0.0131
-7.250 -0.5354 0.06178 0.05955 -0.0301 1.0000 0.0132
-7.000 -0.5310 0.05756 0.05527 -0.0314 1.0000 0.0135
-6.750 -0.5249 0.05324 0.05084 -0.0323 1.0000 0.0137
-6.500 -0.5171 0.04928 0.04676 -0.0325 1.0000 0.0141
-6.250 -0.5078 0.04530 0.04263 -0.0322 1.0000 0.0146
-6.000 -0.4970 0.04149 0.03864 -0.0315 1.0000 0.0152
-5.750 -0.4846 0.03779 0.03472 -0.0305 1.0000 0.0160
-5.500 -0.4703 0.03432 0.03100 -0.0290 1.0000 0.0171
-5.250 -0.4466 0.03370 0.03013 -0.0271 1.0000 0.0194
-5.000 -0.4306 0.03146 0.02758 -0.0254 1.0000 0.0196
-4.750 -0.4236 0.02436 0.01993 -0.0238 1.0000 0.0212
-4.500 -0.4056 0.02230 0.01773 -0.0226 1.0000 0.0224
-4.250 -0.3812 0.01743 0.01212 -0.0197 1.0000 0.0121
-4.000 -0.3588 0.01624 0.01076 -0.0184 1.0000 0.0112
-3.750 -0.3327 0.01492 0.00921 -0.0178 0.9995 0.0115
-3.500 -0.3017 0.01179 0.00575 -0.0183 0.9980 0.0129
-3.250 -0.2652 0.01111 0.00505 -0.0203 0.9953 0.0158
-3.000 -0.2293 0.01028 0.00412 -0.0219 0.9924 0.0175
-2.750 -0.1932 0.00968 0.00343 -0.0236 0.9886 0.0196
-2.500 -0.1563 0.00882 0.00259 -0.0254 0.9851 0.0501
-2.250 -0.1193 0.00792 0.00232 -0.0280 0.9824 0.2088
-2.000 -0.0890 0.00697 0.00213 -0.0291 0.9763 0.4190
-1.750 -0.0560 0.00606 0.00199 -0.0306 0.9717 0.6381
-1.500 -0.0281 0.00549 0.00207 -0.0301 0.9662 0.8345
-1.250 0.0046 0.00536 0.00199 -0.0307 0.9597 0.8744
-1.000 0.0368 0.00525 0.00191 -0.0312 0.9531 0.9038
-0.750 0.0668 0.00517 0.00184 -0.0310 0.9439 0.9317
-0.500 0.0992 0.00512 0.00176 -0.0315 0.9328 0.9496
-0.250 0.1356 0.00509 0.00169 -0.0329 0.9207 0.9626
0.000 0.1742 0.00508 0.00163 -0.0349 0.9067 0.9725
0.250 0.2141 0.00509 0.00157 -0.0372 0.8890 0.9796
0.500 0.2549 0.00512 0.00151 -0.0398 0.8660 0.9849
0.750 0.2949 0.00517 0.00146 -0.0423 0.8399 0.9901
1.000 0.3332 0.00524 0.00143 -0.0445 0.8106 0.9957
1.250 0.3694 0.00534 0.00138 -0.0463 0.7785 1.0000
1.500 0.3901 0.00544 0.00137 -0.0448 0.7489 1.0000
1.750 0.4109 0.00556 0.00138 -0.0433 0.7197 1.0000
2.000 0.4317 0.00568 0.00140 -0.0419 0.6848 1.0000
2.250 0.4523 0.00585 0.00143 -0.0404 0.6468 1.0000
2.500 0.4733 0.00605 0.00149 -0.0390 0.6091 1.0000
2.750 0.4949 0.00627 0.00161 -0.0377 0.5698 1.0000
3.000 0.5165 0.00655 0.00171 -0.0365 0.5246 1.0000
3.250 0.5386 0.00685 0.00184 -0.0354 0.4774 1.0000
3.500 0.5604 0.00724 0.00200 -0.0343 0.4164 1.0000
3.750 0.5828 0.00767 0.00219 -0.0333 0.3571 1.0000
4.000 0.6057 0.00811 0.00244 -0.0325 0.2982 1.0000
4.250 0.6287 0.00860 0.00270 -0.0318 0.2398 1.0000
4.500 0.6519 0.00911 0.00299 -0.0312 0.1860 1.0000
4.750 0.6746 0.00974 0.00335 -0.0305 0.1276 1.0000
5.000 0.6967 0.01049 0.00381 -0.0297 0.0666 1.0000
5.250 0.7183 0.01143 0.00454 -0.0286 0.0295 1.0000
5.500 0.7423 0.01202 0.00526 -0.0278 0.0256 1.0000
5.750 0.7664 0.01255 0.00585 -0.0272 0.0215 1.0000
6.000 0.7868 0.01374 0.00717 -0.0258 0.0183 1.0000
6.250 0.8095 0.01451 0.00805 -0.0249 0.0163 1.0000
6.500 0.8329 0.01521 0.00882 -0.0241 0.0141 1.0000
6.750 0.8550 0.01614 0.00983 -0.0232 0.0122 1.0000
7.000 0.8711 0.01871 0.01261 -0.0212 0.0105 1.0000
7.250 0.8890 0.02170 0.01592 -0.0195 0.0099 1.0000
7.500 0.9090 0.02384 0.01837 -0.0182 0.0097 1.0000
7.750 0.9283 0.02584 0.02065 -0.0169 0.0094 1.0000
8.000 0.9498 0.02668 0.02166 -0.0160 0.0084 1.0000
8.250 0.9656 0.02913 0.02443 -0.0144 0.0079 1.0000
8.500 0.9772 0.03225 0.02793 -0.0125 0.0077 1.0000
8.750 0.9845 0.03584 0.03191 -0.0103 0.0075 1.0000
9.000 0.9814 0.04099 0.03752 -0.0074 0.0077 1.0000
9.250 0.9709 0.04640 0.04334 -0.0046 0.0079 1.0000
9.500 0.9542 0.05150 0.04876 -0.0020 0.0081 1.0000
9.750 0.9327 0.05532 0.05276 0.0009 0.0083 1.0000
10.000 0.9104 0.05951 0.05710 0.0014 0.0085 1.0000
10.250 0.8900 0.06469 0.06242 -0.0011 0.0085 1.0000
10.500 0.8731 0.07086 0.06870 -0.0062 0.0087 1.0000
10.750 0.8514 0.08090 0.07883 -0.0150 0.0091 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 14A-1.4/7.0 AIRFOIL (rg14a147-il)