GOE 780 AIRFOIL (goe780-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 780 AIRFOIL (goe780-il) Reynolds number: 200,000 Max Cl/Cd: 34.32 at α=2.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe780-il-200000-n5.txt Download as CSV file: xf-goe780-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 780 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5007 0.07831 0.07509 -0.0594 1.0000 0.0095
-9.500 -0.5126 0.07407 0.07083 -0.0606 1.0000 0.0102
-9.250 -0.5242 0.06904 0.06573 -0.0632 0.9814 0.0081
-9.000 -0.5249 0.06391 0.06042 -0.0660 0.9605 0.0096
-8.750 -0.5285 0.05921 0.05551 -0.0668 0.9415 0.0089
-8.500 -0.5309 0.05501 0.05108 -0.0659 0.9238 0.0084
-8.250 -0.5289 0.05144 0.04724 -0.0642 0.9091 0.0092
-8.000 -0.5272 0.04798 0.04350 -0.0618 0.8970 0.0092
-7.750 -0.5250 0.04456 0.03976 -0.0587 0.8868 0.0094
-7.500 -0.5231 0.04102 0.03588 -0.0550 0.8784 0.0094
-7.250 -0.5174 0.03803 0.03255 -0.0519 0.8719 0.0090
-7.000 -0.5127 0.03444 0.02853 -0.0477 0.8657 0.0090
-6.750 -0.5022 0.03152 0.02520 -0.0445 0.8607 0.0088
-6.500 -0.4877 0.02851 0.02172 -0.0418 0.8565 0.0087
-6.250 -0.4684 0.02566 0.01836 -0.0398 0.8521 0.0088
-6.000 -0.4426 0.02317 0.01543 -0.0392 0.8485 0.0090
-5.750 -0.4137 0.02148 0.01340 -0.0393 0.8456 0.0104
-5.500 -0.3823 0.01972 0.01138 -0.0401 0.8430 0.0108
-5.250 -0.3541 0.01763 0.00910 -0.0404 0.8401 0.0128
-5.000 -0.3338 0.01663 0.00799 -0.0391 0.8367 0.0145
-4.750 -0.3138 0.01598 0.00719 -0.0377 0.8335 0.0177
-4.250 -0.2749 0.01489 0.00595 -0.0346 0.8270 0.0374
-4.000 -0.2553 0.01449 0.00542 -0.0333 0.8237 0.0510
-3.750 -0.2381 0.01390 0.00502 -0.0315 0.8208 0.0837
-3.500 -0.2194 0.01342 0.00462 -0.0300 0.8182 0.1392
-3.250 -0.2161 0.01227 0.00428 -0.0257 0.8150 0.3387
-3.000 -0.2208 0.01125 0.00405 -0.0193 0.8108 0.5209
-2.750 -0.1275 0.01091 0.00467 -0.0328 0.8115 0.7904
-2.500 -0.0962 0.01129 0.00497 -0.0331 0.8091 0.8268
-2.250 -0.0733 0.01162 0.00521 -0.0318 0.8067 0.8518
-2.000 -0.0452 0.01198 0.00547 -0.0315 0.8045 0.8674
-1.750 -0.0066 0.01258 0.00595 -0.0333 0.8029 0.8805
-1.500 0.0227 0.01284 0.00616 -0.0336 0.8002 0.8888
-1.000 0.0730 0.01299 0.00621 -0.0331 0.7950 0.9003
-0.750 0.1099 0.01313 0.00631 -0.0353 0.7929 0.9022
-0.500 0.1426 0.01320 0.00634 -0.0367 0.7908 0.9043
-0.250 0.1765 0.01336 0.00647 -0.0382 0.7891 0.9089
0.000 0.1948 0.01337 0.00651 -0.0366 0.7863 0.9148
0.250 0.2311 0.01351 0.00668 -0.0387 0.7838 0.9171
0.500 0.2642 0.01361 0.00684 -0.0402 0.7813 0.9202
0.750 0.2914 0.01363 0.00689 -0.0404 0.7775 0.9242
1.000 0.3128 0.01359 0.00688 -0.0392 0.7709 0.9286
1.250 0.3440 0.01355 0.00692 -0.0401 0.7641 0.9303
1.500 0.3753 0.01354 0.00698 -0.0411 0.7608 0.9323
1.750 0.4021 0.01352 0.00709 -0.0412 0.7527 0.9353
2.000 0.4181 0.01307 0.00657 -0.0380 0.7182 0.9403
2.250 0.4413 0.01286 0.00638 -0.0368 0.6744 0.9431
2.500 0.4311 0.01472 0.00643 -0.0292 0.2863 0.9484
2.750 0.4260 0.01623 0.00715 -0.0234 0.0732 0.9552
3.000 0.4494 0.01690 0.00765 -0.0230 0.0214 0.9573
3.250 0.4720 0.01748 0.00837 -0.0223 0.0166 0.9598
3.500 0.4916 0.01805 0.00913 -0.0210 0.0154 0.9630
3.750 0.5034 0.01867 0.00986 -0.0180 0.0145 0.9675
4.000 0.5257 0.01943 0.01072 -0.0175 0.0126 0.9695
4.250 0.5461 0.02039 0.01175 -0.0165 0.0111 0.9724
4.500 0.5661 0.02150 0.01293 -0.0152 0.0108 0.9762
4.750 0.5906 0.02287 0.01439 -0.0147 0.0104 0.9786
5.000 0.6214 0.02434 0.01603 -0.0153 0.0103 0.9795
5.250 0.6524 0.02752 0.01944 -0.0157 0.0086 0.9792
5.500 0.6800 0.02886 0.02108 -0.0155 0.0082 0.9803
5.750 0.7062 0.03108 0.02366 -0.0148 0.0080 0.9814
6.000 0.7277 0.03398 0.02689 -0.0137 0.0083 0.9829
6.500 0.7752 0.04061 0.03423 -0.0094 0.0185 0.9856
6.750 0.7908 0.04361 0.03762 -0.0068 0.0191 0.9878
7.000 0.8048 0.04671 0.04096 -0.0056 0.0184 0.9898
7.250 0.8144 0.05014 0.04458 -0.0040 0.0179 0.9921
7.500 0.8168 0.05458 0.04927 -0.0017 0.0174 0.9946
7.750 0.8134 0.06061 0.05557 0.0006 0.0170 0.9972
8.000 0.8191 0.06426 0.05954 0.0023 0.0169 0.9997
8.250 0.8124 0.06684 0.06229 0.0065 0.0169 1.0000
8.500 0.8013 0.06923 0.06483 0.0115 0.0169 1.0000
8.750 0.7888 0.07148 0.06720 0.0164 0.0169 1.0000
9.000 0.7725 0.07372 0.06955 0.0217 0.0169 1.0000
9.250 0.7521 0.07518 0.07108 0.0277 0.0169 1.0000
9.500 0.7319 0.07687 0.07283 0.0329 0.0169 1.0000
9.750 0.7105 0.07890 0.07493 0.0375 0.0169 1.0000
10.000 0.6913 0.08140 0.07751 0.0407 0.0169 1.0000
10.250 0.6732 0.08425 0.08042 0.0429 0.0169 1.0000
10.500 0.6553 0.08765 0.08388 0.0443 0.0169 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 780 AIRFOIL (goe780-il)