GOE 474 AIRFOIL (goe474-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 474 AIRFOIL (goe474-il) Reynolds number: 50,000 Max Cl/Cd: 38.4 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe474-il-50000.txt Download as CSV file: xf-goe474-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4286 0.09790 0.09098 -0.0186 1.0000 0.1994
-7.750 -0.4288 0.09528 0.08843 -0.0183 1.0000 0.2069
-7.500 -0.4554 0.09525 0.08861 -0.0210 1.0000 0.2101
-7.250 -0.4307 0.08946 0.08279 -0.0172 1.0000 0.2187
-7.000 -0.4535 0.08896 0.08246 -0.0212 1.0000 0.2246
-6.750 -0.4379 0.08433 0.07786 -0.0174 1.0000 0.2368
-6.500 -0.4335 0.08103 0.07462 -0.0160 1.0000 0.2479
-6.250 -0.4337 0.07813 0.07180 -0.0152 1.0000 0.2602
-6.000 -0.4342 0.07532 0.06908 -0.0143 1.0000 0.2735
-5.750 -0.4339 0.07254 0.06636 -0.0131 1.0000 0.2879
-5.500 -0.4324 0.06975 0.06361 -0.0114 1.0000 0.3044
-5.250 -0.4376 0.06769 0.06159 -0.0110 1.0000 0.3274
-5.000 -0.4310 0.06445 0.05843 -0.0064 1.0000 0.3479
-4.250 -0.3577 0.04689 0.03931 -0.0327 1.0000 0.1802
-4.000 -0.3364 0.04221 0.03422 -0.0335 1.0000 0.1636
-3.750 -0.3157 0.03850 0.03007 -0.0336 1.0000 0.1588
-3.500 -0.2940 0.03570 0.02673 -0.0334 1.0000 0.1647
-3.250 -0.2729 0.03312 0.02381 -0.0328 1.0000 0.1708
-3.000 -0.2493 0.03076 0.02094 -0.0321 1.0000 0.1782
-2.750 -0.2267 0.02894 0.01870 -0.0314 1.0000 0.1963
-2.500 -0.2045 0.02706 0.01669 -0.0304 1.0000 0.2181
-2.250 -0.1812 0.02553 0.01485 -0.0296 1.0000 0.2538
-2.000 -0.1607 0.02451 0.01392 -0.0286 1.0000 0.2935
-1.750 -0.1395 0.02370 0.01304 -0.0276 1.0000 0.3241
-1.500 -0.1150 0.02301 0.01214 -0.0270 1.0000 0.3424
-1.250 -0.0889 0.02244 0.01149 -0.0268 1.0000 0.3624
-1.000 -0.0632 0.02200 0.01097 -0.0265 1.0000 0.3822
-0.750 -0.0384 0.02155 0.01057 -0.0262 1.0000 0.4052
-0.500 -0.0147 0.02105 0.01026 -0.0257 1.0000 0.4450
-0.250 0.0230 0.01885 0.00983 -0.0272 1.0000 1.0000
0.000 0.0460 0.01922 0.00959 -0.0265 1.0000 1.0000
0.250 0.0662 0.01959 0.00961 -0.0257 1.0000 1.0000
0.500 0.0858 0.01998 0.00975 -0.0248 1.0000 1.0000
0.750 0.1051 0.02039 0.00996 -0.0240 1.0000 1.0000
1.000 0.1242 0.02084 0.01025 -0.0233 1.0000 1.0000
1.250 0.1432 0.02131 0.01061 -0.0225 1.0000 1.0000
1.500 0.1620 0.02182 0.01102 -0.0219 1.0000 1.0000
1.750 0.1807 0.02237 0.01149 -0.0212 1.0000 1.0000
2.000 0.1991 0.02295 0.01202 -0.0207 1.0000 1.0000
2.250 0.2173 0.02358 0.01262 -0.0202 1.0000 1.0000
2.500 0.2352 0.02426 0.01329 -0.0197 1.0000 1.0000
2.750 0.2527 0.02499 0.01404 -0.0193 1.0000 1.0000
3.000 0.2698 0.02580 0.01487 -0.0190 1.0000 1.0000
3.250 0.3277 0.02746 0.01665 -0.0266 0.9793 1.0000
3.500 0.3921 0.02891 0.01830 -0.0346 0.9512 1.0000
3.750 0.4728 0.02982 0.01948 -0.0442 0.9143 1.0000
4.000 0.5391 0.02985 0.01985 -0.0499 0.8755 1.0000
4.250 0.6142 0.02912 0.01954 -0.0558 0.8389 1.0000
4.500 0.6987 0.02687 0.01786 -0.0606 0.7978 1.0000
4.750 0.7505 0.02460 0.01597 -0.0592 0.7465 1.0000
5.000 0.7930 0.02227 0.01384 -0.0556 0.6741 1.0000
5.250 0.8194 0.02134 0.01249 -0.0504 0.5546 1.0000
5.500 0.8382 0.02244 0.01267 -0.0466 0.4488 1.0000
5.750 0.8616 0.02426 0.01393 -0.0451 0.3841 1.0000
6.000 0.8865 0.02601 0.01545 -0.0441 0.3412 1.0000
6.250 0.9121 0.02773 0.01704 -0.0433 0.3091 1.0000
6.500 0.9352 0.02938 0.01864 -0.0421 0.2802 1.0000
6.750 0.9558 0.03098 0.02020 -0.0407 0.2516 1.0000
7.000 0.9753 0.03277 0.02204 -0.0390 0.2240 1.0000
7.250 0.9935 0.03482 0.02413 -0.0373 0.1962 1.0000
7.500 1.0100 0.03732 0.02689 -0.0352 0.1718 1.0000
7.750 1.0294 0.04015 0.02961 -0.0339 0.1517 1.0000
8.000 1.0433 0.04295 0.03275 -0.0318 0.1368 1.0000
8.250 1.0544 0.04638 0.03676 -0.0292 0.1283 1.0000
8.500 1.0641 0.04994 0.04067 -0.0271 0.1209 1.0000
8.750 1.0697 0.05341 0.04459 -0.0247 0.1150 1.0000
9.000 1.0853 0.05756 0.04870 -0.0237 0.1106 1.0000
9.250 1.0827 0.06218 0.05385 -0.0211 0.1103 1.0000
9.500 1.0754 0.06694 0.05904 -0.0187 0.1102 1.0000
9.750 1.0670 0.07176 0.06418 -0.0167 0.1103 1.0000
10.000 1.0549 0.07666 0.06933 -0.0150 0.1105 1.0000
10.250 0.9822 0.08297 0.07608 -0.0134 0.1200 1.0000
10.500 0.9623 0.08871 0.08187 -0.0142 0.1220 1.0000
10.750 0.9035 0.10080 0.09394 -0.0237 0.1332 1.0000
11.000 0.8962 0.10809 0.10120 -0.0268 0.1364 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 474 AIRFOIL (goe474-il)