GOE 374 AIRFOIL (goe374-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 374 AIRFOIL (goe374-il) Reynolds number: 200,000 Max Cl/Cd: 74.55 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe374-il-200000-n5.txt Download as CSV file: xf-goe374-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 374 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3225 0.10317 0.09979 -0.0219 1.0000 0.0144
-8.000 -0.3213 0.10093 0.09760 -0.0225 1.0000 0.0144
-7.750 -0.3221 0.09874 0.09548 -0.0228 1.0000 0.0144
-7.500 -0.3186 0.09619 0.09297 -0.0240 1.0000 0.0145
-7.250 -0.3137 0.09346 0.09029 -0.0254 1.0000 0.0145
-7.000 -0.3083 0.09072 0.08759 -0.0267 1.0000 0.0145
-6.750 -0.3030 0.08795 0.08485 -0.0276 1.0000 0.0145
-6.500 -0.2981 0.08527 0.08220 -0.0281 1.0000 0.0145
-6.250 -0.2867 0.08208 0.07901 -0.0302 0.9986 0.0145
-6.000 -0.2737 0.07562 0.07260 -0.0314 0.9939 0.0149
-5.750 -0.2503 0.07104 0.06801 -0.0353 0.9880 0.0152
-5.500 -0.2236 0.06674 0.06367 -0.0403 0.9807 0.0156
-5.250 -0.1939 0.06256 0.05943 -0.0459 0.9730 0.0161
-5.000 -0.1594 0.05833 0.05511 -0.0523 0.9662 0.0166
-4.750 -0.1269 0.05434 0.05103 -0.0575 0.9562 0.0176
-4.500 -0.0931 0.05060 0.04717 -0.0625 0.9463 0.0190
-4.250 -0.0447 0.04747 0.04378 -0.0688 0.9385 0.0214
-4.000 -0.0089 0.04419 0.04028 -0.0722 0.9281 0.0216
-3.750 0.0229 0.04076 0.03666 -0.0745 0.9175 0.0217
-3.250 0.0758 0.03344 0.02905 -0.0779 0.8943 0.0202
-3.000 0.1059 0.03016 0.02550 -0.0789 0.8814 0.0175
-2.750 0.1357 0.02749 0.02253 -0.0793 0.8674 0.0173
-2.500 0.1649 0.02550 0.02023 -0.0791 0.8521 0.0198
-2.250 0.1931 0.02323 0.01761 -0.0787 0.8358 0.0199
-2.000 0.2213 0.02116 0.01513 -0.0779 0.8186 0.0203
-1.750 0.2491 0.01950 0.01304 -0.0770 0.8001 0.0230
-1.500 0.2743 0.01845 0.01178 -0.0766 0.7784 0.0259
-1.250 0.3012 0.01709 0.01003 -0.0756 0.7555 0.0269
-1.000 0.3277 0.01616 0.00868 -0.0746 0.7292 0.0314
-0.750 0.3532 0.01524 0.00747 -0.0738 0.7017 0.0328
-0.500 0.3788 0.01456 0.00650 -0.0729 0.6758 0.0342
0.000 0.4298 0.01381 0.00521 -0.0713 0.6314 0.0380
0.250 0.4551 0.01345 0.00468 -0.0706 0.6118 0.0385
0.500 0.4802 0.01312 0.00421 -0.0699 0.5933 0.0391
0.750 0.5048 0.01264 0.00370 -0.0692 0.5746 0.0417
1.000 0.5298 0.01247 0.00347 -0.0686 0.5553 0.0435
1.250 0.5547 0.01235 0.00325 -0.0680 0.5361 0.0430
1.500 0.5799 0.01228 0.00309 -0.0674 0.5169 0.0426
1.750 0.6051 0.01229 0.00297 -0.0668 0.4985 0.0423
2.000 0.6302 0.01235 0.00290 -0.0662 0.4807 0.0422
2.250 0.6552 0.01245 0.00286 -0.0656 0.4641 0.0422
2.500 0.6803 0.01257 0.00288 -0.0651 0.4490 0.0426
2.750 0.7055 0.01271 0.00292 -0.0646 0.4351 0.0432
3.000 0.7308 0.01286 0.00299 -0.0641 0.4219 0.0446
3.250 0.7562 0.01301 0.00309 -0.0636 0.4100 0.0474
3.500 0.7987 0.01154 0.00339 -0.0673 0.3964 1.0000
3.750 0.8236 0.01178 0.00354 -0.0668 0.3838 1.0000
4.000 0.8485 0.01202 0.00371 -0.0662 0.3712 1.0000
4.250 0.8733 0.01226 0.00392 -0.0657 0.3599 1.0000
4.500 0.8985 0.01247 0.00413 -0.0652 0.3496 1.0000
4.750 0.9235 0.01270 0.00436 -0.0648 0.3395 1.0000
5.000 0.9482 0.01295 0.00463 -0.0642 0.3291 1.0000
5.250 0.9728 0.01320 0.00490 -0.0637 0.3179 1.0000
5.500 0.9975 0.01345 0.00518 -0.0632 0.3062 1.0000
5.750 1.0218 0.01373 0.00549 -0.0627 0.2944 1.0000
6.000 1.0459 0.01403 0.00585 -0.0621 0.2819 1.0000
6.250 1.0690 0.01440 0.00619 -0.0614 0.2622 1.0000
6.500 1.0904 0.01492 0.00657 -0.0605 0.2304 1.0000
6.750 1.1114 0.01551 0.00703 -0.0596 0.1971 1.0000
7.000 1.1311 0.01627 0.00762 -0.0586 0.1591 1.0000
7.250 1.1506 0.01707 0.00827 -0.0576 0.1299 1.0000
7.500 1.1673 0.01819 0.00910 -0.0562 0.0860 1.0000
7.750 1.1762 0.02018 0.01058 -0.0539 0.0179 1.0000
8.000 1.1948 0.02105 0.01164 -0.0525 0.0145 1.0000
8.250 1.2118 0.02207 0.01285 -0.0509 0.0122 1.0000
8.500 1.2269 0.02327 0.01436 -0.0490 0.0108 1.0000
8.750 1.2402 0.02457 0.01590 -0.0470 0.0102 1.0000
9.000 1.2545 0.02563 0.01715 -0.0453 0.0095 1.0000
9.250 1.2664 0.02682 0.01851 -0.0433 0.0085 1.0000
9.500 1.2745 0.02818 0.02005 -0.0409 0.0079 1.0000
9.750 1.2773 0.02966 0.02167 -0.0377 0.0076 1.0000
10.000 1.2769 0.03130 0.02345 -0.0345 0.0074 1.0000
10.250 1.2752 0.03315 0.02543 -0.0316 0.0072 1.0000
10.500 1.2708 0.03539 0.02777 -0.0291 0.0070 1.0000
10.750 1.2659 0.03793 0.03041 -0.0270 0.0069 1.0000
11.000 1.2631 0.04051 0.03309 -0.0255 0.0068 1.0000
11.250 1.2589 0.04356 0.03621 -0.0242 0.0067 1.0000
11.500 1.2583 0.04657 0.03926 -0.0229 0.0065 1.0000
11.750 1.2613 0.04916 0.04204 -0.0221 0.0065 1.0000
12.000 1.2630 0.05188 0.04494 -0.0217 0.0063 1.0000
12.250 1.2638 0.05479 0.04803 -0.0216 0.0062 1.0000
12.500 1.2636 0.05797 0.05139 -0.0217 0.0060 1.0000
12.750 1.2611 0.06149 0.05513 -0.0224 0.0056 1.0000
13.000 1.2583 0.06523 0.05905 -0.0230 0.0055 1.0000
13.250 1.2540 0.06922 0.06321 -0.0242 0.0052 1.0000
13.500 1.2487 0.07356 0.06774 -0.0253 0.0052 1.0000
13.750 1.2419 0.07823 0.07258 -0.0268 0.0051 1.0000
14.000 1.2341 0.08324 0.07779 -0.0287 0.0051 1.0000
14.250 1.2252 0.08859 0.08333 -0.0309 0.0051 1.0000
14.500 1.2156 0.09427 0.08919 -0.0336 0.0051 1.0000
14.750 1.2052 0.10032 0.09543 -0.0366 0.0051 1.0000
15.000 1.1939 0.10675 0.10203 -0.0400 0.0051 1.0000
15.250 1.1827 0.11340 0.10886 -0.0436 0.0051 1.0000
15.500 1.1718 0.12027 0.11588 -0.0476 0.0051 1.0000
15.750 1.1601 0.12757 0.12335 -0.0518 0.0051 1.0000
16.000 1.1483 0.13518 0.13112 -0.0564 0.0051 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 374 AIRFOIL (goe374-il)