RAF 30 MOD AIRFOIL (raf30md-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 30 MOD AIRFOIL (raf30md-il) Reynolds number: 50,000 Max Cl/Cd: 25.1 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf30md-il-50000-n5.txt Download as CSV file: xf-raf30md-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 30 MOD 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6705 0.09236 0.08554 -0.0031 1.0000 0.0424
-9.250 -0.6792 0.08518 0.07841 -0.0086 1.0000 0.0412
-9.000 -0.6961 0.07831 0.07155 -0.0138 1.0000 0.0396
-8.750 -0.7199 0.07212 0.06520 -0.0166 1.0000 0.0381
-8.500 -0.7325 0.06706 0.05987 -0.0177 1.0000 0.0375
-8.250 -0.7352 0.06267 0.05525 -0.0177 1.0000 0.0376
-8.000 -0.7311 0.05896 0.05141 -0.0174 1.0000 0.0387
-7.750 -0.7246 0.05567 0.04795 -0.0169 1.0000 0.0404
-7.500 -0.7185 0.05185 0.04379 -0.0163 1.0000 0.0419
-7.250 -0.7099 0.04792 0.03941 -0.0154 1.0000 0.0431
-7.000 -0.6981 0.04405 0.03503 -0.0144 1.0000 0.0439
-6.750 -0.6830 0.04040 0.03082 -0.0131 1.0000 0.0450
-6.500 -0.6646 0.03729 0.02706 -0.0118 1.0000 0.0475
-6.250 -0.6454 0.03445 0.02388 -0.0109 1.0000 0.0515
-6.000 -0.6232 0.03205 0.02120 -0.0099 1.0000 0.0554
-5.750 -0.5989 0.02986 0.01859 -0.0088 1.0000 0.0607
-5.500 -0.5763 0.02806 0.01673 -0.0079 1.0000 0.0688
-5.250 -0.5514 0.02637 0.01471 -0.0067 1.0000 0.0755
-5.000 -0.5292 0.02504 0.01334 -0.0057 1.0000 0.0885
-4.750 -0.5073 0.02364 0.01184 -0.0045 1.0000 0.0996
-4.500 -0.4856 0.02227 0.01039 -0.0034 1.0000 0.1148
-4.250 -0.4649 0.02065 0.00901 -0.0024 1.0000 0.1538
-4.000 -0.4481 0.01884 0.00798 -0.0011 1.0000 0.2853
-3.750 -0.4353 0.01750 0.00764 0.0017 1.0000 0.4742
-3.500 -0.4234 0.01687 0.00777 0.0065 1.0000 0.6496
-3.250 -0.4049 0.01672 0.00782 0.0104 1.0000 0.7445
-3.000 -0.3788 0.01668 0.00770 0.0126 1.0000 0.8071
-2.750 -0.3418 0.01675 0.00758 0.0126 1.0000 0.8596
-2.500 -0.2993 0.01674 0.00725 0.0105 1.0000 0.8961
-2.250 -0.2528 0.01667 0.00687 0.0072 1.0000 0.9237
-2.000 -0.2023 0.01654 0.00644 0.0026 1.0000 0.9474
-1.750 -0.1508 0.01634 0.00599 -0.0025 1.0000 0.9691
-1.500 -0.0992 0.01605 0.00549 -0.0079 1.0000 0.9897
-1.250 -0.0668 0.01576 0.00510 -0.0099 1.0000 1.0000
-1.000 -0.0534 0.01554 0.00485 -0.0081 1.0000 1.0000
-0.750 -0.0401 0.01537 0.00465 -0.0061 1.0000 1.0000
-0.500 -0.0269 0.01525 0.00451 -0.0041 1.0000 1.0000
-0.250 -0.0136 0.01518 0.00442 -0.0020 1.0000 1.0000
0.000 0.0000 0.01515 0.00439 0.0000 1.0000 1.0000
0.250 0.0136 0.01518 0.00442 0.0020 1.0000 1.0000
0.500 0.0269 0.01525 0.00451 0.0041 1.0000 1.0000
0.750 0.0401 0.01537 0.00465 0.0061 1.0000 1.0000
1.000 0.0533 0.01554 0.00485 0.0081 1.0000 1.0000
1.250 0.0668 0.01576 0.00510 0.0099 1.0000 1.0000
1.500 0.0992 0.01605 0.00550 0.0079 0.9897 1.0000
1.750 0.1507 0.01634 0.00600 0.0025 0.9691 1.0000
2.000 0.2024 0.01654 0.00644 -0.0026 0.9473 1.0000
2.250 0.2529 0.01667 0.00688 -0.0072 0.9237 1.0000
2.500 0.2994 0.01674 0.00725 -0.0106 0.8961 1.0000
2.750 0.3418 0.01675 0.00758 -0.0126 0.8596 1.0000
3.000 0.3789 0.01668 0.00771 -0.0126 0.8072 1.0000
3.250 0.4049 0.01672 0.00783 -0.0105 0.7449 1.0000
3.500 0.4235 0.01687 0.00778 -0.0065 0.6494 1.0000
3.750 0.4354 0.01750 0.00764 -0.0017 0.4742 1.0000
4.000 0.4482 0.01884 0.00798 0.0011 0.2851 1.0000
4.250 0.4650 0.02065 0.00901 0.0024 0.1539 1.0000
4.500 0.4856 0.02227 0.01039 0.0034 0.1148 1.0000
4.750 0.5073 0.02364 0.01184 0.0045 0.0996 1.0000
5.000 0.5292 0.02504 0.01334 0.0057 0.0883 1.0000
5.250 0.5513 0.02637 0.01472 0.0067 0.0754 1.0000
5.500 0.5762 0.02806 0.01673 0.0079 0.0688 1.0000
5.750 0.5988 0.02986 0.01860 0.0088 0.0606 1.0000
6.000 0.6231 0.03204 0.02119 0.0099 0.0554 1.0000
6.250 0.6454 0.03445 0.02388 0.0109 0.0515 1.0000
6.500 0.6645 0.03729 0.02705 0.0118 0.0476 1.0000
6.750 0.6828 0.04041 0.03084 0.0131 0.0450 1.0000
7.000 0.6980 0.04402 0.03500 0.0144 0.0439 1.0000
7.250 0.7097 0.04791 0.03941 0.0155 0.0431 1.0000
7.500 0.7182 0.05186 0.04379 0.0164 0.0419 1.0000
7.750 0.7243 0.05567 0.04796 0.0170 0.0405 1.0000
8.000 0.7297 0.05915 0.05164 0.0174 0.0390 1.0000
8.250 0.7350 0.06264 0.05522 0.0178 0.0377 1.0000
8.500 0.7321 0.06706 0.05987 0.0177 0.0376 1.0000
8.750 0.7171 0.07240 0.06552 0.0164 0.0384 1.0000
9.000 0.6955 0.07838 0.07162 0.0137 0.0398 1.0000
9.250 0.6787 0.08534 0.07856 0.0084 0.0414 1.0000
9.500 0.6708 0.09228 0.08545 0.0031 0.0423 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 30 MOD AIRFOIL (raf30md-il)