GOE 510 AIRFOIL (goe510-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 510 AIRFOIL (goe510-il) Reynolds number: 1,000,000 Max Cl/Cd: 99.46 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe510-il-1000000-n5.txt Download as CSV file: xf-goe510-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 510 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.250 -0.7022 0.02808 0.02499 -0.1404 0.9610 0.0278
-13.000 -0.6861 0.02560 0.02228 -0.1414 0.9591 0.0279
-12.750 -0.6652 0.02355 0.02005 -0.1426 0.9577 0.0282
-12.500 -0.6608 0.02265 0.01906 -0.1389 0.9511 0.0284
-12.250 -0.6385 0.02166 0.01798 -0.1389 0.9481 0.0286
-12.000 -0.6101 0.02088 0.01712 -0.1398 0.9461 0.0288
-11.750 -0.5797 0.02003 0.01617 -0.1412 0.9441 0.0289
-11.500 -0.5634 0.01939 0.01546 -0.1394 0.9388 0.0291
-11.250 -0.5430 0.01886 0.01486 -0.1383 0.9339 0.0293
-11.000 -0.5155 0.01822 0.01413 -0.1387 0.9308 0.0295
-10.750 -0.4849 0.01763 0.01346 -0.1397 0.9281 0.0297
-10.500 -0.4703 0.01714 0.01289 -0.1373 0.9210 0.0299
-10.250 -0.4440 0.01657 0.01223 -0.1373 0.9152 0.0301
-10.000 -0.4196 0.01605 0.01161 -0.1369 0.9087 0.0304
-9.750 -0.3949 0.01554 0.01101 -0.1364 0.8997 0.0306
-9.500 -0.3706 0.01503 0.01039 -0.1359 0.8904 0.0309
-9.250 -0.3445 0.01456 0.00980 -0.1357 0.8805 0.0312
-9.000 -0.3209 0.01419 0.00933 -0.1349 0.8700 0.0315
-8.750 -0.2966 0.01386 0.00889 -0.1343 0.8604 0.0317
-8.500 -0.2731 0.01351 0.00841 -0.1334 0.8491 0.0320
-8.250 -0.2508 0.01321 0.00800 -0.1322 0.8361 0.0321
-8.000 -0.2289 0.01292 0.00760 -0.1310 0.8236 0.0322
-7.750 -0.2082 0.01250 0.00708 -0.1296 0.8132 0.0326
-7.500 -0.1865 0.01222 0.00672 -0.1283 0.8026 0.0329
-7.250 -0.1644 0.01201 0.00644 -0.1270 0.7913 0.0332
-7.000 -0.1426 0.01185 0.00621 -0.1256 0.7790 0.0335
-6.750 -0.1215 0.01172 0.00599 -0.1241 0.7641 0.0338
-6.500 -0.1003 0.01157 0.00576 -0.1226 0.7497 0.0340
-6.250 -0.0784 0.01145 0.00557 -0.1213 0.7360 0.0344
-6.000 -0.0570 0.01131 0.00536 -0.1198 0.7219 0.0347
-5.750 -0.0363 0.01119 0.00515 -0.1182 0.7063 0.0350
-5.500 -0.0161 0.01107 0.00494 -0.1164 0.6897 0.0354
-5.250 0.0039 0.01097 0.00474 -0.1147 0.6728 0.0357
-5.000 0.0246 0.01087 0.00455 -0.1130 0.6582 0.0361
-4.750 0.0456 0.01078 0.00438 -0.1115 0.6456 0.0364
-4.500 0.0681 0.01071 0.00425 -0.1102 0.6346 0.0368
-4.250 0.0902 0.01061 0.00409 -0.1089 0.6248 0.0371
-4.000 0.1113 0.01050 0.00390 -0.1074 0.6137 0.0374
-3.750 0.1335 0.01032 0.00368 -0.1061 0.6044 0.0379
-3.500 0.1548 0.01024 0.00356 -0.1046 0.5939 0.0384
-3.250 0.1774 0.01016 0.00344 -0.1034 0.5840 0.0389
-3.000 0.1986 0.01010 0.00334 -0.1018 0.5747 0.0393
-2.750 0.2188 0.01006 0.00325 -0.1001 0.5626 0.0398
-2.500 0.2394 0.01001 0.00315 -0.0984 0.5505 0.0403
-2.250 0.2588 0.00998 0.00306 -0.0965 0.5350 0.0408
-2.000 0.2782 0.00998 0.00298 -0.0946 0.5179 0.0413
-1.750 0.2972 0.01001 0.00292 -0.0927 0.4986 0.0418
-1.500 0.3160 0.01006 0.00288 -0.0907 0.4784 0.0423
-1.250 0.3356 0.01013 0.00286 -0.0889 0.4617 0.0428
-1.000 0.3561 0.01013 0.00280 -0.0873 0.4499 0.0432
-0.750 0.3764 0.01010 0.00272 -0.0856 0.4404 0.0440
-0.500 0.3975 0.01009 0.00269 -0.0841 0.4314 0.0448
-0.250 0.4187 0.01011 0.00268 -0.0826 0.4238 0.0454
0.000 0.4409 0.01013 0.00267 -0.0814 0.4174 0.0463
0.250 0.4623 0.01017 0.00268 -0.0800 0.4108 0.0471
0.500 0.4845 0.01019 0.00269 -0.0788 0.4055 0.0479
0.750 0.5067 0.01022 0.00271 -0.0775 0.4005 0.0488
1.000 0.5282 0.01028 0.00274 -0.0762 0.3950 0.0494
1.250 0.5497 0.01032 0.00276 -0.0748 0.3903 0.0502
1.500 0.5719 0.01032 0.00276 -0.0736 0.3862 0.0516
1.750 0.5940 0.01035 0.00279 -0.0724 0.3820 0.0529
2.000 0.6156 0.01040 0.00283 -0.0711 0.3778 0.0544
2.250 0.6369 0.01047 0.00288 -0.0697 0.3733 0.0559
2.500 0.6593 0.01051 0.00293 -0.0686 0.3698 0.0574
2.750 0.6812 0.01055 0.00298 -0.0673 0.3663 0.0600
3.000 0.7025 0.01061 0.00305 -0.0660 0.3626 0.0648
3.250 0.7227 0.01064 0.00312 -0.0645 0.3589 0.0843
3.500 0.7407 0.01054 0.00324 -0.0625 0.3556 0.1709
3.750 0.7620 0.01056 0.00335 -0.0613 0.3520 0.2050
4.000 0.7830 0.01063 0.00347 -0.0599 0.3482 0.2304
4.250 0.8035 0.01073 0.00359 -0.0585 0.3439 0.2450
4.500 0.8237 0.01085 0.00373 -0.0571 0.3402 0.2599
4.750 0.8458 0.01093 0.00384 -0.0560 0.3373 0.2708
5.000 0.8674 0.01103 0.00397 -0.0548 0.3338 0.2804
5.250 0.8882 0.01116 0.00411 -0.0535 0.3294 0.2910
5.750 0.9283 0.01143 0.00442 -0.0507 0.3208 0.3174
6.000 0.9488 0.01154 0.00457 -0.0494 0.3168 0.3324
6.250 0.9685 0.01168 0.00475 -0.0480 0.3124 0.3510
6.500 0.9851 0.01174 0.00496 -0.0460 0.3080 0.4329
7.000 1.1518 0.01158 0.00596 -0.0718 0.2797 0.9888
7.250 1.1729 0.01194 0.00628 -0.0708 0.2711 0.9970
7.500 1.2035 0.01239 0.00664 -0.0721 0.2577 0.9993
7.750 1.2251 0.01286 0.00701 -0.0715 0.2418 1.0000
8.000 1.2370 0.01327 0.00736 -0.0688 0.2298 1.0000
8.250 1.2483 0.01373 0.00774 -0.0660 0.2165 1.0000
8.500 1.2629 0.01410 0.00808 -0.0639 0.2086 1.0000
8.750 1.2743 0.01464 0.00853 -0.0613 0.1946 1.0000
9.000 1.2858 0.01520 0.00903 -0.0588 0.1822 1.0000
9.250 1.2961 0.01586 0.00960 -0.0562 0.1666 1.0000
9.500 1.3023 0.01673 0.01035 -0.0531 0.1460 1.0000
9.750 1.3075 0.01772 0.01121 -0.0500 0.1254 1.0000
10.250 1.3084 0.02033 0.01355 -0.0430 0.0803 1.0000
10.500 1.2903 0.02284 0.01583 -0.0376 0.0326 1.0000
10.750 1.2942 0.02417 0.01716 -0.0351 0.0192 1.0000
11.000 1.3052 0.02511 0.01811 -0.0335 0.0170 1.0000
11.250 1.3161 0.02608 0.01911 -0.0319 0.0158 1.0000
11.500 1.3264 0.02713 0.02018 -0.0304 0.0147 1.0000
11.750 1.3363 0.02822 0.02132 -0.0290 0.0140 1.0000
12.000 1.3466 0.02934 0.02248 -0.0277 0.0135 1.0000
12.250 1.3562 0.03055 0.02373 -0.0264 0.0130 1.0000
12.500 1.3651 0.03184 0.02505 -0.0252 0.0124 1.0000
12.750 1.3729 0.03327 0.02652 -0.0239 0.0121 1.0000
13.000 1.3795 0.03483 0.02813 -0.0227 0.0116 1.0000
13.250 1.3846 0.03659 0.02994 -0.0216 0.0111 1.0000
13.500 1.3916 0.03822 0.03162 -0.0207 0.0108 1.0000
13.750 1.3981 0.03996 0.03342 -0.0198 0.0107 1.0000
14.000 1.4031 0.04187 0.03538 -0.0190 0.0104 1.0000
14.250 1.4073 0.04391 0.03748 -0.0183 0.0101 1.0000
14.500 1.4106 0.04611 0.03974 -0.0176 0.0099 1.0000
14.750 1.4133 0.04843 0.04212 -0.0171 0.0096 1.0000
15.000 1.4150 0.05089 0.04464 -0.0166 0.0094 1.0000
15.250 1.4151 0.05355 0.04737 -0.0162 0.0092 1.0000
15.500 1.4146 0.05636 0.05025 -0.0159 0.0091 1.0000
15.750 1.4112 0.05957 0.05353 -0.0157 0.0088 1.0000
16.000 1.4081 0.06281 0.05685 -0.0156 0.0087 1.0000
16.250 1.4055 0.06602 0.06013 -0.0156 0.0085 1.0000
16.500 1.4019 0.06943 0.06363 -0.0158 0.0085 1.0000
16.750 1.3977 0.07298 0.06727 -0.0161 0.0083 1.0000
17.000 1.3943 0.07649 0.07085 -0.0164 0.0082 1.0000
17.250 1.3874 0.08050 0.07495 -0.0169 0.0082 1.0000
17.500 1.3829 0.08421 0.07873 -0.0174 0.0080 1.0000
17.750 1.3782 0.08798 0.08258 -0.0180 0.0078 1.0000
18.000 1.3706 0.09221 0.08690 -0.0188 0.0077 1.0000
18.250 1.3632 0.09642 0.09120 -0.0196 0.0077 1.0000
18.500 1.3566 0.10053 0.09539 -0.0205 0.0076 1.0000
18.750 1.3487 0.10487 0.09981 -0.0215 0.0075 1.0000
19.000 1.3437 0.10882 0.10383 -0.0225 0.0073 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 510 AIRFOIL (goe510-il)