EPPLER 521 AIRFOIL (e521-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 521 AIRFOIL (e521-il) Reynolds number: 200,000 Max Cl/Cd: 43.22 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e521-il-200000-n5.txt Download as CSV file: xf-e521-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 521 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
------ -------- --------- --------- -------- -------- --------
-17.000 -0.8563 0.12893 0.12507 0.0047 1.0000 0.0091
-16.750 -0.8850 0.11740 0.11326 -0.0017 1.0000 0.0087
-16.500 -0.9013 0.10922 0.10496 -0.0065 1.0000 0.0088
-16.250 -0.9176 0.10136 0.09693 -0.0110 1.0000 0.0088
-16.000 -0.9318 0.09420 0.08961 -0.0150 1.0000 0.0087
-15.750 -0.9433 0.08770 0.08299 -0.0189 1.0000 0.0089
-15.500 -0.9551 0.08148 0.07663 -0.0225 1.0000 0.0091
-15.250 -0.9653 0.07575 0.07075 -0.0256 1.0000 0.0091
-15.000 -0.9742 0.07048 0.06536 -0.0284 1.0000 0.0095
-14.750 -0.9835 0.06549 0.06020 -0.0306 1.0000 0.0094
-14.500 -0.9889 0.06129 0.05586 -0.0325 1.0000 0.0096
-14.250 -0.9928 0.05753 0.05196 -0.0340 1.0000 0.0100
-14.000 -0.9974 0.05390 0.04818 -0.0350 1.0000 0.0104
-13.750 -1.0012 0.05053 0.04466 -0.0356 1.0000 0.0106
-13.500 -1.0025 0.04761 0.04159 -0.0359 1.0000 0.0111
-13.250 -1.0032 0.04486 0.03866 -0.0359 1.0000 0.0115
-13.000 -1.0029 0.04231 0.03596 -0.0356 1.0000 0.0120
-12.750 -1.0049 0.03971 0.03327 -0.0352 1.0000 0.0124
-12.500 -1.0031 0.03757 0.03101 -0.0346 1.0000 0.0131
-12.250 -0.9976 0.03582 0.02917 -0.0340 1.0000 0.0141
-12.000 -0.9903 0.03424 0.02748 -0.0333 1.0000 0.0154
-11.750 -0.9820 0.03275 0.02584 -0.0324 1.0000 0.0165
-11.500 -0.9789 0.03092 0.02397 -0.0312 1.0000 0.0177
-11.250 -0.9709 0.02952 0.02251 -0.0301 1.0000 0.0190
-11.000 -0.9604 0.02833 0.02123 -0.0291 1.0000 0.0209
-10.750 -0.9483 0.02726 0.02003 -0.0280 1.0000 0.0228
-10.500 -0.9412 0.02587 0.01864 -0.0266 1.0000 0.0251
-10.250 -0.9290 0.02487 0.01759 -0.0254 1.0000 0.0277
-10.000 -0.9152 0.02400 0.01659 -0.0243 1.0000 0.0304
-9.750 -0.9051 0.02295 0.01555 -0.0228 1.0000 0.0338
-9.500 -0.8854 0.02211 0.01465 -0.0227 0.9668 0.0386
-9.250 -0.8616 0.02114 0.01363 -0.0234 0.9443 0.0447
-9.000 -0.8406 0.02046 0.01283 -0.0231 0.9295 0.0512
-8.750 -0.8240 0.01975 0.01211 -0.0219 0.9177 0.0593
-8.500 -0.8064 0.01913 0.01146 -0.0207 0.9091 0.0684
-8.250 -0.7884 0.01857 0.01084 -0.0196 0.9016 0.0788
-8.000 -0.7702 0.01804 0.01028 -0.0184 0.8955 0.0910
-7.750 -0.7516 0.01750 0.00976 -0.0173 0.8891 0.1053
-7.500 -0.7327 0.01698 0.00927 -0.0162 0.8840 0.1218
-7.250 -0.7129 0.01650 0.00882 -0.0152 0.8795 0.1402
-7.000 -0.6926 0.01603 0.00838 -0.0143 0.8746 0.1599
-6.750 -0.6725 0.01558 0.00796 -0.0133 0.8702 0.1816
-6.500 -0.6521 0.01514 0.00757 -0.0124 0.8666 0.2053
-6.250 -0.6304 0.01472 0.00721 -0.0117 0.8623 0.2305
-6.000 -0.6091 0.01431 0.00687 -0.0109 0.8583 0.2578
-5.750 -0.5877 0.01391 0.00654 -0.0100 0.8548 0.2859
-5.500 -0.5657 0.01354 0.00623 -0.0093 0.8515 0.3159
-5.250 -0.5432 0.01316 0.00596 -0.0087 0.8476 0.3473
-5.000 -0.5208 0.01281 0.00570 -0.0080 0.8439 0.3798
-4.750 -0.4982 0.01248 0.00547 -0.0072 0.8407 0.4148
-4.500 -0.4752 0.01217 0.00527 -0.0064 0.8378 0.4503
-4.250 -0.4515 0.01189 0.00511 -0.0059 0.8339 0.4872
-4.000 -0.4275 0.01165 0.00501 -0.0052 0.8303 0.5233
-3.750 -0.4030 0.01148 0.00496 -0.0046 0.8271 0.5579
-3.500 -0.3778 0.01140 0.00495 -0.0039 0.8243 0.5895
-3.250 -0.3515 0.01137 0.00499 -0.0036 0.8204 0.6166
-3.000 -0.3250 0.01139 0.00504 -0.0032 0.8165 0.6392
-2.750 -0.2983 0.01143 0.00508 -0.0028 0.8130 0.6576
-2.500 -0.2715 0.01149 0.00509 -0.0024 0.8099 0.6733
-2.250 -0.2443 0.01155 0.00515 -0.0022 0.8054 0.6865
-2.000 -0.2172 0.01160 0.00518 -0.0020 0.8010 0.6983
-1.750 -0.1902 0.01165 0.00516 -0.0017 0.7972 0.7095
-1.500 -0.1630 0.01171 0.00521 -0.0015 0.7931 0.7189
-1.250 -0.1358 0.01177 0.00527 -0.0013 0.7878 0.7272
-1.000 -0.1089 0.01179 0.00523 -0.0010 0.7833 0.7364
-0.750 -0.0815 0.01184 0.00528 -0.0007 0.7788 0.7425
-0.500 -0.0545 0.01186 0.00529 -0.0005 0.7728 0.7503
-0.250 -0.0271 0.01187 0.00529 -0.0002 0.7679 0.7557
0.000 0.0000 0.01188 0.00531 0.0000 0.7620 0.7620
0.250 0.0272 0.01187 0.00529 0.0002 0.7557 0.7679
0.500 0.0545 0.01186 0.00529 0.0005 0.7503 0.7728
0.750 0.0815 0.01184 0.00528 0.0007 0.7425 0.7788
1.000 0.1089 0.01179 0.00523 0.0010 0.7364 0.7833
1.250 0.1358 0.01177 0.00527 0.0013 0.7272 0.7878
1.500 0.1630 0.01171 0.00521 0.0014 0.7189 0.7931
1.750 0.1902 0.01164 0.00516 0.0017 0.7095 0.7972
2.000 0.2172 0.01160 0.00518 0.0020 0.6983 0.8011
2.250 0.2443 0.01155 0.00515 0.0022 0.6865 0.8054
2.500 0.2715 0.01149 0.00509 0.0024 0.6732 0.8099
2.750 0.2983 0.01143 0.00508 0.0028 0.6576 0.8130
3.000 0.3250 0.01139 0.00504 0.0032 0.6392 0.8165
3.250 0.3515 0.01137 0.00499 0.0036 0.6166 0.8204
3.500 0.3778 0.01140 0.00495 0.0039 0.5896 0.8243
3.750 0.4030 0.01148 0.00496 0.0046 0.5579 0.8271
4.000 0.4275 0.01165 0.00501 0.0052 0.5234 0.8303
4.250 0.4515 0.01189 0.00511 0.0059 0.4871 0.8339
4.500 0.4752 0.01217 0.00527 0.0064 0.4506 0.8378
4.750 0.4981 0.01248 0.00547 0.0072 0.4150 0.8407
5.000 0.5207 0.01281 0.00570 0.0080 0.3800 0.8439
5.250 0.5431 0.01316 0.00596 0.0087 0.3474 0.8476
5.750 0.5876 0.01391 0.00654 0.0101 0.2858 0.8548
6.000 0.6089 0.01430 0.00687 0.0109 0.2579 0.8583
6.250 0.6303 0.01472 0.00721 0.0117 0.2305 0.8624
6.500 0.6519 0.01514 0.00757 0.0124 0.2054 0.8666
6.750 0.6723 0.01557 0.00795 0.0134 0.1817 0.8702
7.000 0.6924 0.01602 0.00838 0.0144 0.1603 0.8746
7.250 0.7126 0.01650 0.00881 0.0153 0.1404 0.8796
7.500 0.7325 0.01698 0.00927 0.0162 0.1221 0.8841
7.750 0.7513 0.01750 0.00976 0.0173 0.1051 0.8892
8.000 0.7699 0.01804 0.01028 0.0184 0.0907 0.8955
8.250 0.7881 0.01857 0.01083 0.0196 0.0787 0.9017
8.500 0.8061 0.01913 0.01145 0.0208 0.0684 0.9093
8.750 0.8237 0.01974 0.01210 0.0220 0.0593 0.9179
9.000 0.8403 0.02045 0.01282 0.0231 0.0512 0.9297
9.250 0.8614 0.02114 0.01362 0.0235 0.0446 0.9446
9.500 0.8852 0.02211 0.01465 0.0227 0.0386 0.9674
9.750 0.9048 0.02294 0.01554 0.0228 0.0340 1.0000
10.000 0.9148 0.02401 0.01659 0.0244 0.0306 1.0000
10.250 0.9287 0.02487 0.01758 0.0255 0.0277 1.0000
10.500 0.9408 0.02588 0.01865 0.0266 0.0252 1.0000
10.750 0.9479 0.02726 0.02003 0.0281 0.0228 1.0000
11.000 0.9608 0.02828 0.02117 0.0290 0.0208 1.0000
11.250 0.9711 0.02949 0.02248 0.0301 0.0190 1.0000
11.500 0.9787 0.03093 0.02397 0.0312 0.0177 1.0000
11.750 0.9818 0.03276 0.02586 0.0324 0.0165 1.0000
12.000 0.9903 0.03423 0.02747 0.0333 0.0153 1.0000
12.250 0.9982 0.03575 0.02910 0.0340 0.0140 1.0000
12.500 1.0031 0.03757 0.03101 0.0346 0.0132 1.0000
12.750 1.0053 0.03967 0.03322 0.0352 0.0125 1.0000
13.000 1.0027 0.04233 0.03598 0.0356 0.0120 1.0000
13.250 1.0036 0.04482 0.03863 0.0359 0.0115 1.0000
13.500 1.0028 0.04757 0.04154 0.0359 0.0111 1.0000
13.750 1.0007 0.05060 0.04474 0.0356 0.0106 1.0000
14.000 0.9973 0.05392 0.04821 0.0350 0.0103 1.0000
14.250 0.9937 0.05742 0.05185 0.0340 0.0101 1.0000
14.500 0.9888 0.06131 0.05589 0.0325 0.0097 1.0000
14.750 0.9828 0.06560 0.06032 0.0306 0.0095 1.0000
15.000 0.9759 0.07028 0.06513 0.0283 0.0093 1.0000
15.250 0.9661 0.07566 0.07064 0.0255 0.0090 1.0000
15.500 0.9558 0.08136 0.07652 0.0226 0.0092 1.0000
15.750 0.9459 0.08727 0.08257 0.0191 0.0090 1.0000
16.000 0.9327 0.09403 0.08946 0.0152 0.0089 1.0000
16.250 0.9169 0.10151 0.09712 0.0110 0.0089 1.0000
16.500 0.9009 0.10938 0.10515 0.0065 0.0090 1.0000
16.750 0.8858 0.11734 0.11323 0.0017 0.0088 1.0000
17.000 0.8555 0.12929 0.12544 -0.0049 0.0092 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 521 AIRFOIL (e521-il)