GOE 445 AIRFOIL (goe445-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 445 AIRFOIL (goe445-il) Reynolds number: 200,000 Max Cl/Cd: 42.41 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe445-il-200000.txt Download as CSV file: xf-goe445-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 445 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6381 0.08617 0.08275 -0.0112 1.0000 0.0326
-8.500 -0.6454 0.08117 0.07774 -0.0151 1.0000 0.0326
-8.250 -0.6536 0.07708 0.07360 -0.0165 1.0000 0.0326
-8.000 -0.6587 0.07326 0.06961 -0.0180 1.0000 0.0330
-7.750 -0.6607 0.06976 0.06593 -0.0183 1.0000 0.0332
-7.500 -0.6610 0.06633 0.06226 -0.0180 1.0000 0.0333
-7.250 -0.6678 0.05824 0.05409 -0.0183 1.0000 0.0343
-7.000 -0.6590 0.05409 0.05005 -0.0179 1.0000 0.0359
-6.750 -0.6502 0.05094 0.04682 -0.0171 1.0000 0.0383
-6.500 -0.6404 0.04755 0.04320 -0.0162 1.0000 0.0411
-6.250 -0.6341 0.04473 0.03957 -0.0138 1.0000 0.0466
-6.000 -0.6201 0.04030 0.03537 -0.0135 1.0000 0.0498
-5.750 -0.6085 0.03779 0.03249 -0.0119 1.0000 0.0609
-5.500 -0.5950 0.03570 0.03004 -0.0103 1.0000 0.0733
-5.250 -0.5774 0.03302 0.02739 -0.0092 1.0000 0.0803
-5.000 -0.5617 0.03061 0.02478 -0.0077 1.0000 0.0927
-4.500 -0.5045 0.02271 0.01527 -0.0010 1.0000 0.0323
-4.250 -0.4820 0.01975 0.01202 0.0006 1.0000 0.0298
-4.000 -0.4582 0.01774 0.00979 0.0021 1.0000 0.0287
-3.750 -0.4350 0.01616 0.00807 0.0036 1.0000 0.0294
-3.500 -0.4120 0.01533 0.00713 0.0049 1.0000 0.0329
-3.250 -0.3931 0.01372 0.00553 0.0068 1.0000 0.0361
-3.000 -0.3728 0.01279 0.00450 0.0084 1.0000 0.0397
-2.750 -0.3515 0.01201 0.00362 0.0100 1.0000 0.0491
-2.500 -0.3469 0.00909 0.00283 0.0136 1.0000 0.4934
-2.250 -0.3385 0.00798 0.00285 0.0184 1.0000 0.7483
-2.000 -0.2620 0.00832 0.00352 0.0112 1.0000 0.9528
-1.750 -0.2022 0.00851 0.00343 0.0048 1.0000 0.9711
-1.500 -0.1491 0.00854 0.00329 -0.0007 1.0000 0.9836
-1.250 -0.0931 0.00850 0.00312 -0.0069 1.0000 0.9947
-1.000 -0.0545 0.00836 0.00290 -0.0097 1.0000 1.0000
-0.750 -0.0410 0.00823 0.00273 -0.0073 1.0000 1.0000
-0.500 -0.0274 0.00814 0.00262 -0.0049 1.0000 1.0000
-0.250 -0.0137 0.00808 0.00255 -0.0025 1.0000 1.0000
0.000 0.0000 0.00806 0.00253 0.0000 1.0000 1.0000
0.250 0.0137 0.00808 0.00255 0.0025 1.0000 1.0000
0.500 0.0274 0.00814 0.00262 0.0049 1.0000 1.0000
0.750 0.0410 0.00823 0.00273 0.0073 1.0000 1.0000
1.000 0.0545 0.00836 0.00290 0.0097 1.0000 1.0000
1.250 0.0932 0.00850 0.00312 0.0068 0.9947 1.0000
1.500 0.1490 0.00854 0.00329 0.0007 0.9837 1.0000
1.750 0.2020 0.00851 0.00343 -0.0047 0.9712 1.0000
2.000 0.2620 0.00832 0.00352 -0.0112 0.9529 1.0000
2.250 0.3384 0.00798 0.00285 -0.0184 0.7477 1.0000
2.500 0.3463 0.00914 0.00283 -0.0135 0.4835 1.0000
2.750 0.3515 0.01200 0.00361 -0.0100 0.0496 1.0000
3.000 0.3727 0.01278 0.00449 -0.0084 0.0399 1.0000
3.250 0.3930 0.01370 0.00551 -0.0068 0.0363 1.0000
3.500 0.4120 0.01529 0.00710 -0.0049 0.0328 1.0000
3.750 0.4349 0.01616 0.00807 -0.0036 0.0293 1.0000
4.000 0.4581 0.01774 0.00978 -0.0021 0.0288 1.0000
4.250 0.4819 0.01974 0.01201 -0.0006 0.0298 1.0000
4.500 0.5044 0.02268 0.01523 0.0010 0.0323 1.0000
7.250 0.6186 0.04755 0.04403 0.0196 0.0362 1.0000
7.500 0.6244 0.05156 0.04799 0.0203 0.0346 1.0000
9.000 0.6336 0.09191 0.08844 0.0049 0.0323 1.0000
9.250 0.6321 0.09667 0.09316 0.0011 0.0319 1.0000
9.500 0.6300 0.10155 0.09799 -0.0030 0.0307 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 445 AIRFOIL (goe445-il)