EPPLER 297 AIRFOIL (e297-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 297 AIRFOIL (e297-il) Reynolds number: 100,000 Max Cl/Cd: 34.85 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e297-il-100000-n5.txt Download as CSV file: xf-e297-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 297 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6623 0.10009 0.09534 -0.0131 1.0000 0.0158
-11.750 -0.6852 0.08803 0.08327 -0.0207 1.0000 0.0150
-11.500 -0.7472 0.07039 0.06528 -0.0326 1.0000 0.0122
-11.000 -0.7833 0.06035 0.05487 -0.0365 1.0000 0.0123
-10.750 -0.7968 0.05669 0.05104 -0.0367 1.0000 0.0125
-10.500 -0.8099 0.05319 0.04732 -0.0359 1.0000 0.0125
-10.250 -0.8216 0.04977 0.04359 -0.0341 1.0000 0.0123
-10.000 -0.8270 0.04777 0.04149 -0.0322 1.0000 0.0130
-9.750 -0.8305 0.04552 0.03897 -0.0299 1.0000 0.0134
-9.500 -0.8287 0.04350 0.03673 -0.0280 1.0000 0.0145
-9.250 -0.8245 0.04079 0.03370 -0.0260 1.0000 0.0150
-9.000 -0.8167 0.03807 0.03062 -0.0241 1.0000 0.0156
-8.750 -0.8050 0.03532 0.02750 -0.0224 1.0000 0.0164
-8.500 -0.7896 0.03284 0.02468 -0.0209 1.0000 0.0174
-8.250 -0.7721 0.03076 0.02231 -0.0196 1.0000 0.0185
-8.000 -0.7587 0.02882 0.02030 -0.0181 1.0000 0.0214
-7.750 -0.7432 0.02759 0.01898 -0.0166 1.0000 0.0248
-7.500 -0.7280 0.02613 0.01732 -0.0147 1.0000 0.0279
-7.250 -0.7189 0.02478 0.01598 -0.0121 1.0000 0.0326
-7.000 -0.7094 0.02393 0.01502 -0.0093 1.0000 0.0386
-6.750 -0.7051 0.02290 0.01396 -0.0057 1.0000 0.0431
-6.500 -0.6974 0.02224 0.01314 -0.0025 1.0000 0.0520
-6.250 -0.6852 0.02122 0.01213 -0.0004 0.9977 0.0628
-6.000 -0.6594 0.02006 0.01104 -0.0010 0.9910 0.0847
-5.750 -0.6333 0.01900 0.01018 -0.0017 0.9848 0.1255
-5.500 -0.6071 0.01797 0.00945 -0.0025 0.9794 0.1872
-5.250 -0.5834 0.01699 0.00885 -0.0028 0.9733 0.2667
-5.000 -0.5572 0.01605 0.00835 -0.0035 0.9686 0.3624
-4.750 -0.5352 0.01536 0.00804 -0.0028 0.9625 0.4526
-4.500 -0.5086 0.01494 0.00790 -0.0026 0.9579 0.5393
-4.250 -0.4778 0.01488 0.00802 -0.0027 0.9545 0.6073
-4.000 -0.4522 0.01501 0.00816 -0.0018 0.9489 0.6522
-3.750 -0.4223 0.01520 0.00829 -0.0018 0.9445 0.6850
-3.500 -0.3904 0.01539 0.00836 -0.0023 0.9410 0.7116
-3.250 -0.3662 0.01555 0.00836 -0.0014 0.9348 0.7336
-3.000 -0.3368 0.01575 0.00846 -0.0013 0.9304 0.7513
-2.750 -0.3044 0.01593 0.00853 -0.0019 0.9270 0.7664
-2.500 -0.2807 0.01608 0.00859 -0.0009 0.9207 0.7797
-2.250 -0.2514 0.01620 0.00859 -0.0009 0.9162 0.7921
-2.000 -0.2230 0.01629 0.00860 -0.0009 0.9117 0.8039
-1.750 -0.1958 0.01643 0.00869 -0.0005 0.9060 0.8127
-1.500 -0.1655 0.01650 0.00869 -0.0008 0.9016 0.8216
-1.250 -0.1411 0.01653 0.00864 -0.0001 0.8955 0.8315
-1.000 -0.1109 0.01662 0.00870 -0.0004 0.8903 0.8380
-0.750 -0.0810 0.01661 0.00865 -0.0007 0.8860 0.8463
-0.500 -0.0558 0.01668 0.00871 -0.0001 0.8787 0.8531
-0.250 -0.0270 0.01666 0.00866 -0.0002 0.8735 0.8604
0.000 0.0000 0.01670 0.00871 0.0000 0.8665 0.8665
0.250 0.0271 0.01665 0.00866 0.0002 0.8604 0.8735
0.500 0.0558 0.01668 0.00871 0.0001 0.8531 0.8787
0.750 0.0810 0.01660 0.00864 0.0007 0.8463 0.8860
1.000 0.1108 0.01661 0.00870 0.0004 0.8381 0.8903
1.250 0.1410 0.01653 0.00864 0.0002 0.8315 0.8956
1.500 0.1655 0.01649 0.00868 0.0008 0.8215 0.9016
1.750 0.1958 0.01643 0.00868 0.0005 0.8127 0.9060
2.000 0.2227 0.01629 0.00859 0.0010 0.8039 0.9118
2.250 0.2514 0.01620 0.00859 0.0010 0.7921 0.9163
2.500 0.2806 0.01608 0.00859 0.0009 0.7798 0.9208
2.750 0.3043 0.01593 0.00852 0.0019 0.7665 0.9271
3.000 0.3367 0.01574 0.00845 0.0014 0.7513 0.9305
3.250 0.3660 0.01554 0.00835 0.0014 0.7336 0.9350
3.500 0.3903 0.01538 0.00835 0.0023 0.7117 0.9411
3.750 0.4222 0.01519 0.00828 0.0018 0.6850 0.9446
4.000 0.4519 0.01499 0.00815 0.0019 0.6518 0.9492
4.250 0.4779 0.01487 0.00801 0.0027 0.6074 0.9547
4.500 0.5086 0.01493 0.00789 0.0026 0.5393 0.9581
4.750 0.5350 0.01535 0.00803 0.0029 0.4529 0.9628
5.000 0.5572 0.01605 0.00833 0.0034 0.3617 0.9688
5.250 0.5834 0.01698 0.00884 0.0028 0.2666 0.9735
5.500 0.6071 0.01796 0.00944 0.0025 0.1870 0.9798
5.750 0.6335 0.01900 0.01017 0.0017 0.1250 0.9851
6.000 0.6596 0.02005 0.01104 0.0010 0.0847 0.9914
6.250 0.6855 0.02122 0.01212 0.0004 0.0626 0.9981
6.500 0.6966 0.02221 0.01312 0.0027 0.0523 1.0000
6.750 0.7043 0.02288 0.01394 0.0058 0.0434 1.0000
7.000 0.7089 0.02391 0.01500 0.0094 0.0386 1.0000
7.250 0.7187 0.02478 0.01597 0.0122 0.0328 1.0000
7.500 0.7282 0.02613 0.01732 0.0146 0.0279 1.0000
7.750 0.7436 0.02759 0.01898 0.0166 0.0247 1.0000
8.000 0.7592 0.02884 0.02032 0.0180 0.0213 1.0000
8.250 0.7728 0.03083 0.02237 0.0194 0.0185 1.0000
8.500 0.7904 0.03295 0.02481 0.0208 0.0172 1.0000
8.750 0.8059 0.03537 0.02755 0.0222 0.0164 1.0000
9.000 0.8175 0.03817 0.03073 0.0240 0.0155 1.0000
9.250 0.8254 0.04079 0.03370 0.0258 0.0147 1.0000
9.500 0.8302 0.04337 0.03658 0.0277 0.0143 1.0000
9.750 0.8315 0.04567 0.03913 0.0297 0.0135 1.0000
10.000 0.8297 0.04737 0.04102 0.0318 0.0127 1.0000
10.250 0.8182 0.05100 0.04502 0.0343 0.0131 1.0000
10.500 0.8111 0.05330 0.04744 0.0356 0.0125 1.0000
10.750 0.7981 0.05657 0.05082 0.0362 0.0120 1.0000
11.000 0.7846 0.06035 0.05483 0.0361 0.0120 1.0000
11.250 0.7670 0.06499 0.05966 0.0349 0.0120 1.0000
11.500 0.7488 0.07037 0.06521 0.0323 0.0120 1.0000
11.750 0.6849 0.08873 0.08397 0.0200 0.0153 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 297 AIRFOIL (e297-il)