GOE 474 AIRFOIL (goe474-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 474 AIRFOIL (goe474-il) Reynolds number: 100,000 Max Cl/Cd: 54.5 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe474-il-100000-n5.txt Download as CSV file: xf-goe474-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 474 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4359 0.09128 0.08626 -0.0304 1.0000 0.0273
-8.500 -0.4377 0.08768 0.08272 -0.0312 1.0000 0.0272
-8.250 -0.4410 0.08434 0.07945 -0.0317 1.0000 0.0269
-8.000 -0.4466 0.08123 0.07643 -0.0320 1.0000 0.0267
-7.750 -0.4510 0.07754 0.07282 -0.0332 1.0000 0.0266
-7.500 -0.4543 0.07339 0.06873 -0.0351 1.0000 0.0265
-7.250 -0.4561 0.06919 0.06456 -0.0368 1.0000 0.0263
-7.000 -0.4575 0.06456 0.05994 -0.0385 1.0000 0.0262
-6.750 -0.4576 0.05972 0.05506 -0.0400 1.0000 0.0262
-6.500 -0.4568 0.05415 0.04940 -0.0415 1.0000 0.0262
-6.250 -0.4551 0.04688 0.04187 -0.0431 1.0000 0.0265
-6.000 -0.4510 0.03904 0.03346 -0.0443 1.0000 0.0271
-5.750 -0.4407 0.03375 0.02737 -0.0441 1.0000 0.0281
-5.500 -0.4286 0.02993 0.02314 -0.0432 1.0000 0.0294
-5.250 -0.3990 0.02702 0.01976 -0.0449 0.9960 0.0312
-5.000 -0.3678 0.02481 0.01704 -0.0463 0.9916 0.0345
-4.750 -0.3353 0.02286 0.01450 -0.0475 0.9876 0.0386
-4.500 -0.3049 0.02109 0.01252 -0.0484 0.9833 0.0425
-4.250 -0.2724 0.02008 0.01125 -0.0495 0.9787 0.0496
-4.000 -0.2407 0.01911 0.01017 -0.0505 0.9742 0.0560
-3.750 -0.2090 0.01846 0.00943 -0.0515 0.9690 0.0662
-3.500 -0.1743 0.01785 0.00876 -0.0529 0.9652 0.0779
-3.250 -0.1447 0.01763 0.00850 -0.0533 0.9587 0.0996
-3.000 -0.1104 0.01746 0.00821 -0.0547 0.9541 0.1212
-2.750 -0.0788 0.01731 0.00793 -0.0555 0.9486 0.1361
-2.500 -0.0472 0.01718 0.00769 -0.0564 0.9429 0.1497
-2.250 -0.0121 0.01707 0.00752 -0.0580 0.9390 0.1664
-2.000 0.0170 0.01692 0.00733 -0.0584 0.9327 0.1799
-1.750 0.0497 0.01667 0.00701 -0.0594 0.9277 0.1890
-1.500 0.0852 0.01633 0.00669 -0.0610 0.9243 0.1972
-1.250 0.1125 0.01615 0.00650 -0.0609 0.9174 0.2061
-0.750 0.1798 0.01566 0.00607 -0.0632 0.9088 0.2299
-0.500 0.2070 0.01551 0.00598 -0.0631 0.9014 0.2480
-0.250 0.2404 0.01519 0.00587 -0.0644 0.8971 0.2919
0.250 0.3423 0.01335 0.00583 -0.0735 0.8939 1.0000
0.500 0.3706 0.01338 0.00577 -0.0734 0.8843 1.0000
0.750 0.4009 0.01335 0.00567 -0.0735 0.8738 1.0000
1.000 0.4328 0.01326 0.00553 -0.0739 0.8633 1.0000
1.250 0.4641 0.01319 0.00541 -0.0740 0.8527 1.0000
1.500 0.4902 0.01320 0.00540 -0.0733 0.8402 1.0000
1.750 0.5171 0.01317 0.00537 -0.0726 0.8267 1.0000
2.000 0.5449 0.01306 0.00524 -0.0719 0.8100 1.0000
2.250 0.5690 0.01303 0.00519 -0.0705 0.7898 1.0000
2.500 0.5942 0.01303 0.00520 -0.0694 0.7716 1.0000
2.750 0.6196 0.01305 0.00524 -0.0685 0.7546 1.0000
3.000 0.6447 0.01308 0.00527 -0.0674 0.7344 1.0000
3.250 0.6689 0.01311 0.00530 -0.0661 0.7077 1.0000
3.500 0.6918 0.01318 0.00533 -0.0645 0.6731 1.0000
3.750 0.7148 0.01330 0.00537 -0.0629 0.6314 1.0000
4.000 0.7368 0.01352 0.00546 -0.0613 0.5781 1.0000
4.250 0.7568 0.01392 0.00559 -0.0593 0.4968 1.0000
4.500 0.7734 0.01465 0.00584 -0.0569 0.4181 1.0000
4.750 0.7910 0.01540 0.00631 -0.0550 0.3599 1.0000
5.000 0.8094 0.01613 0.00683 -0.0533 0.3083 1.0000
5.250 0.8282 0.01686 0.00739 -0.0518 0.2644 1.0000
5.500 0.8475 0.01757 0.00796 -0.0504 0.2309 1.0000
5.750 0.8675 0.01825 0.00859 -0.0490 0.2077 1.0000
6.000 0.8874 0.01896 0.00926 -0.0477 0.1894 1.0000
6.250 0.9074 0.01970 0.00997 -0.0464 0.1736 1.0000
6.500 0.9275 0.02042 0.01074 -0.0451 0.1591 1.0000
6.750 0.9478 0.02113 0.01149 -0.0440 0.1452 1.0000
7.000 0.9683 0.02183 0.01228 -0.0428 0.1331 1.0000
7.250 0.9885 0.02250 0.01303 -0.0416 0.1191 1.0000
7.500 1.0079 0.02312 0.01362 -0.0406 0.1000 1.0000
7.750 1.0278 0.02383 0.01447 -0.0394 0.0818 1.0000
8.000 1.0460 0.02475 0.01539 -0.0381 0.0622 1.0000
8.250 1.0625 0.02595 0.01657 -0.0365 0.0465 1.0000
8.500 1.0765 0.02742 0.01802 -0.0346 0.0377 1.0000
8.750 1.0910 0.02882 0.01955 -0.0327 0.0320 1.0000
9.000 1.1021 0.03057 0.02138 -0.0305 0.0291 1.0000
9.250 1.1154 0.03221 0.02328 -0.0285 0.0266 1.0000
9.500 1.1275 0.03386 0.02511 -0.0265 0.0242 1.0000
9.750 1.1368 0.03559 0.02696 -0.0245 0.0222 1.0000
10.000 1.1426 0.03804 0.02952 -0.0221 0.0211 1.0000
10.250 1.1512 0.04023 0.03209 -0.0198 0.0205 1.0000
10.500 1.1572 0.04267 0.03485 -0.0174 0.0199 1.0000
10.750 1.1598 0.04527 0.03776 -0.0149 0.0194 1.0000
11.000 1.1585 0.04803 0.04086 -0.0124 0.0188 1.0000
11.250 1.1536 0.05092 0.04406 -0.0101 0.0182 1.0000
11.500 1.1454 0.05412 0.04755 -0.0081 0.0177 1.0000
11.750 1.1346 0.05757 0.05129 -0.0067 0.0173 1.0000
12.000 1.1205 0.06174 0.05573 -0.0061 0.0173 1.0000
12.250 1.1046 0.06630 0.06054 -0.0065 0.0171 1.0000
12.500 1.0860 0.07169 0.06617 -0.0081 0.0170 1.0000
12.750 1.0636 0.07849 0.07321 -0.0111 0.0174 1.0000
13.000 1.0389 0.08667 0.08159 -0.0161 0.0177 1.0000
13.250 1.0115 0.09702 0.09212 -0.0230 0.0185 1.0000
13.500 0.9838 0.10902 0.10423 -0.0310 0.0193 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 474 AIRFOIL (goe474-il)