EPPLER 327 AIRFOIL (e327-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 327 AIRFOIL (e327-il) Reynolds number: 500,000 Max Cl/Cd: 83.07 at α=10° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e327-il-500000.txt Download as CSV file: xf-e327-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 327 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4455 0.08783 0.08565 -0.0251 1.0000 0.0150
-9.500 -0.4480 0.08373 0.08158 -0.0274 1.0000 0.0154
-9.250 -0.4597 0.07774 0.07561 -0.0324 1.0000 0.0158
-9.000 -0.4762 0.07260 0.07044 -0.0349 1.0000 0.0155
-8.750 -0.4935 0.06874 0.06654 -0.0347 1.0000 0.0157
-8.500 -0.5118 0.06564 0.06339 -0.0323 1.0000 0.0155
-8.250 -0.5265 0.06285 0.06055 -0.0290 1.0000 0.0156
-8.000 -0.5356 0.05988 0.05750 -0.0262 1.0000 0.0161
-7.750 -0.5428 0.05671 0.05423 -0.0230 1.0000 0.0166
-7.500 -0.5615 0.05092 0.04789 -0.0160 1.0000 0.0178
-7.250 -0.5501 0.04627 0.04322 -0.0173 0.9924 0.0183
-7.000 -0.5266 0.04358 0.04045 -0.0191 0.9710 0.0186
-6.750 -0.4874 0.04021 0.03689 -0.0241 0.9328 0.0195
-6.500 -0.4613 0.03460 0.03037 -0.0243 0.8807 0.0224
-6.250 -0.4435 0.03292 0.02856 -0.0236 0.8328 0.0230
-6.000 -0.4279 0.03178 0.02723 -0.0220 0.7986 0.0237
-5.750 -0.4120 0.03023 0.02539 -0.0198 0.7720 0.0253
-4.750 -0.3291 0.01779 0.01084 -0.0104 0.6978 0.0146
-4.500 -0.3019 0.01664 0.00956 -0.0100 0.6810 0.0139
-4.250 -0.2752 0.01542 0.00820 -0.0095 0.6656 0.0134
-4.000 -0.2498 0.01450 0.00713 -0.0088 0.6506 0.0132
-3.750 -0.2262 0.01375 0.00626 -0.0077 0.6364 0.0133
-3.500 -0.2035 0.01319 0.00561 -0.0066 0.6229 0.0133
-3.250 -0.1811 0.01270 0.00504 -0.0054 0.6101 0.0136
-3.000 -0.1583 0.01231 0.00456 -0.0043 0.5976 0.0140
-2.750 -0.1368 0.01182 0.00402 -0.0029 0.5859 0.0156
-2.500 -0.1132 0.01157 0.00368 -0.0020 0.5745 0.0178
-2.250 -0.0912 0.01114 0.00327 -0.0007 0.5634 0.0299
-2.000 -0.0727 0.01047 0.00299 0.0010 0.5536 0.1240
-1.750 -0.0688 0.00894 0.00259 0.0049 0.5453 0.4363
-1.500 -0.0731 0.00764 0.00272 0.0118 0.5375 0.7970
-1.250 -0.0509 0.00809 0.00314 0.0139 0.5281 0.8409
-1.000 -0.0268 0.00864 0.00368 0.0158 0.5185 0.8640
-0.750 -0.0008 0.00897 0.00393 0.0169 0.5100 0.8735
-0.500 0.0236 0.00927 0.00414 0.0181 0.5013 0.8840
-0.250 0.0534 0.00979 0.00462 0.0189 0.4927 0.8925
0.000 0.0779 0.01004 0.00478 0.0200 0.4845 0.9010
0.250 0.1139 0.01044 0.00512 0.0192 0.4766 0.9056
0.500 0.1419 0.01050 0.00510 0.0192 0.4690 0.9080
0.750 0.1663 0.01048 0.00500 0.0198 0.4620 0.9103
1.000 0.1882 0.01040 0.00487 0.0208 0.4552 0.9131
1.250 0.2098 0.01037 0.00473 0.0218 0.4492 0.9156
1.500 0.2374 0.01031 0.00466 0.0218 0.4426 0.9166
1.750 0.2647 0.01034 0.00461 0.0219 0.4363 0.9178
2.000 0.2920 0.01035 0.00459 0.0219 0.4303 0.9190
2.250 0.3187 0.01035 0.00456 0.0220 0.4243 0.9201
2.500 0.3446 0.01042 0.00455 0.0222 0.4185 0.9216
2.750 0.3710 0.01040 0.00454 0.0224 0.4131 0.9232
3.000 0.3965 0.01041 0.00451 0.0227 0.4077 0.9247
3.250 0.4213 0.01047 0.00450 0.0231 0.4024 0.9263
3.500 0.4466 0.01044 0.00450 0.0234 0.3972 0.9282
3.750 0.4712 0.01046 0.00450 0.0238 0.3922 0.9300
4.000 0.4969 0.01054 0.00451 0.0240 0.3871 0.9312
4.250 0.5240 0.01052 0.00454 0.0240 0.3821 0.9322
4.500 0.5506 0.01057 0.00458 0.0241 0.3770 0.9333
4.750 0.5766 0.01070 0.00465 0.0242 0.3719 0.9346
5.000 0.6033 0.01071 0.00472 0.0243 0.3672 0.9360
5.250 0.6294 0.01077 0.00479 0.0244 0.3619 0.9377
5.500 0.6546 0.01091 0.00489 0.0247 0.3569 0.9395
5.750 0.6805 0.01093 0.00497 0.0249 0.3520 0.9412
6.000 0.7057 0.01098 0.00504 0.0251 0.3467 0.9430
6.250 0.7299 0.01113 0.00515 0.0255 0.3415 0.9448
6.500 0.7563 0.01114 0.00524 0.0255 0.3362 0.9462
6.750 0.7825 0.01122 0.00535 0.0256 0.3306 0.9477
7.000 0.8081 0.01138 0.00550 0.0258 0.3252 0.9493
7.250 0.8343 0.01144 0.00564 0.0258 0.3193 0.9512
7.500 0.8588 0.01157 0.00577 0.0262 0.3132 0.9536
7.750 0.8834 0.01168 0.00594 0.0265 0.3074 0.9560
8.000 0.9069 0.01177 0.00606 0.0271 0.3007 0.9586
8.250 0.9319 0.01192 0.00624 0.0272 0.2941 0.9607
8.500 0.9589 0.01205 0.00643 0.0270 0.2867 0.9625
8.750 0.9849 0.01224 0.00665 0.0269 0.2791 0.9648
9.000 1.0108 0.01241 0.00687 0.0269 0.2709 0.9673
9.250 1.0350 0.01260 0.00711 0.0271 0.2626 0.9705
9.500 1.0594 0.01286 0.00737 0.0272 0.2532 0.9734
9.750 1.0887 0.01312 0.00769 0.0263 0.2425 0.9752
10.000 1.1173 0.01345 0.00805 0.0254 0.2306 0.9774
10.250 1.1447 0.01384 0.00845 0.0247 0.2178 0.9801
10.500 1.1692 0.01429 0.00890 0.0245 0.2029 0.9838
10.750 1.1973 0.01489 0.00947 0.0232 0.1856 0.9857
11.000 1.2244 0.01560 0.01014 0.0220 0.1681 0.9878
11.250 1.2501 0.01635 0.01086 0.0211 0.1502 0.9908
11.500 1.2732 0.01722 0.01169 0.0204 0.1334 0.9950
11.750 1.2954 0.01822 0.01266 0.0195 0.1174 0.9999
12.000 1.2716 0.01840 0.01286 0.0282 0.1147 1.0000
12.250 1.2735 0.01927 0.01372 0.0312 0.1050 1.0000
12.500 1.2770 0.02031 0.01476 0.0336 0.0958 1.0000
12.750 1.2788 0.02159 0.01604 0.0356 0.0874 1.0000
13.000 1.2774 0.02325 0.01770 0.0374 0.0799 1.0000
13.250 1.2792 0.02491 0.01941 0.0385 0.0733 1.0000
13.500 1.2775 0.02704 0.02157 0.0393 0.0673 1.0000
13.750 1.2756 0.02940 0.02397 0.0398 0.0620 1.0000
14.000 1.2718 0.03208 0.02672 0.0399 0.0573 1.0000
14.250 1.2685 0.03487 0.02956 0.0398 0.0532 1.0000
14.500 1.2581 0.03848 0.03321 0.0395 0.0495 1.0000
14.750 1.2559 0.04133 0.03615 0.0391 0.0464 1.0000
15.000 1.2466 0.04501 0.03989 0.0385 0.0436 1.0000
15.250 1.2335 0.04921 0.04414 0.0376 0.0412 1.0000
15.500 1.2297 0.05242 0.04745 0.0369 0.0387 1.0000
15.750 1.2189 0.05656 0.05166 0.0358 0.0368 1.0000
16.000 1.2064 0.06108 0.05624 0.0345 0.0350 1.0000
16.250 1.1962 0.06544 0.06068 0.0331 0.0334 1.0000
16.500 1.1910 0.06926 0.06460 0.0318 0.0316 1.0000
16.750 1.1823 0.07363 0.06902 0.0302 0.0300 1.0000
17.000 1.1704 0.07855 0.07399 0.0284 0.0287 1.0000
17.250 1.1602 0.08331 0.07883 0.0265 0.0274 1.0000
17.500 1.1564 0.08726 0.08287 0.0249 0.0262 1.0000
17.750 1.1512 0.09150 0.08719 0.0231 0.0247 1.0000
18.000 1.1414 0.09649 0.09220 0.0209 0.0231 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 327 AIRFOIL (e327-il)