GOE 746 AIRFOIL (goe746-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 746 AIRFOIL (goe746-il) Reynolds number: 200,000 Max Cl/Cd: 73.45 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe746-il-200000.txt Download as CSV file: xf-goe746-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 746 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 -0.4185 0.11078 0.10704 0.0153 0.6840 0.0303
-8.500 -0.4184 0.10802 0.10428 0.0124 0.6781 0.0304
-8.250 -0.4201 0.10524 0.10147 0.0096 0.6734 0.0305
-8.000 -0.4091 0.09911 0.09536 0.0104 0.6673 0.0309
-7.750 -0.3948 0.09501 0.09118 0.0117 0.6616 0.0315
-7.500 -0.3874 0.09186 0.08801 0.0114 0.6566 0.0323
-7.250 -0.3840 0.08898 0.08514 0.0104 0.6509 0.0333
-7.000 -0.3802 0.08608 0.08220 0.0089 0.6464 0.0342
-6.750 -0.3735 0.08293 0.07900 0.0067 0.6424 0.0354
-6.500 -0.3641 0.07962 0.07568 0.0037 0.6375 0.0366
-6.250 -0.3478 0.07675 0.07267 -0.0030 0.6335 0.0383
-6.000 -0.3282 0.07367 0.06932 -0.0078 0.6301 0.0388
-5.750 -0.3223 0.06780 0.06350 -0.0074 0.6265 0.0396
-5.500 -0.3100 0.06459 0.06032 -0.0068 0.6214 0.0405
-5.250 -0.2947 0.06173 0.05738 -0.0071 0.6171 0.0418
-5.000 -0.2770 0.05889 0.05440 -0.0080 0.6137 0.0442
-4.750 -0.2392 0.05715 0.05210 -0.0124 0.6097 0.0493
-4.500 -0.2284 0.05129 0.04632 -0.0127 0.6058 0.0505
-4.250 -0.2120 0.04871 0.04372 -0.0123 0.6020 0.0520
-4.000 -0.1923 0.04643 0.04131 -0.0122 0.5987 0.0546
-3.750 -0.1562 0.04639 0.04053 -0.0128 0.5945 0.0617
-3.500 -0.1429 0.04064 0.03500 -0.0132 0.5905 0.0639
-3.250 -0.1225 0.03878 0.03311 -0.0130 0.5868 0.0674
-3.000 -0.0944 0.03705 0.03074 -0.0125 0.5840 0.0765
-2.750 -0.0738 0.03440 0.02820 -0.0125 0.5804 0.0790
-2.500 -0.0491 0.03293 0.02662 -0.0123 0.5760 0.0856
-2.250 -0.0250 0.03092 0.02435 -0.0119 0.5723 0.0938
-2.000 -0.0004 0.02966 0.02280 -0.0114 0.5692 0.1072
-1.750 0.0241 0.02854 0.02153 -0.0111 0.5660 0.1219
-1.500 0.0491 0.02734 0.02026 -0.0109 0.5619 0.1372
-1.250 0.0735 0.02606 0.01894 -0.0106 0.5581 0.1536
-1.000 0.0979 0.02500 0.01782 -0.0101 0.5548 0.1735
-0.750 0.1216 0.02407 0.01682 -0.0096 0.5519 0.2033
-0.500 0.1641 0.02159 0.01333 -0.0083 0.5482 0.1084
-0.250 0.1977 0.02026 0.01152 -0.0073 0.5445 0.0731
0.000 0.2264 0.01920 0.01030 -0.0069 0.5411 0.0688
0.250 0.2549 0.01853 0.00941 -0.0065 0.5382 0.0660
0.500 0.2833 0.01807 0.00894 -0.0064 0.5345 0.0652
0.750 0.3111 0.01767 0.00856 -0.0062 0.5303 0.0653
1.000 0.3381 0.01734 0.00825 -0.0059 0.5267 0.0684
1.250 0.3646 0.01704 0.00791 -0.0054 0.5237 0.0698
1.500 0.3910 0.01689 0.00778 -0.0050 0.5202 0.0709
1.750 0.4175 0.01678 0.00774 -0.0048 0.5157 0.0729
2.000 0.4440 0.01667 0.00759 -0.0044 0.5119 0.0761
2.250 0.4703 0.01652 0.00738 -0.0038 0.5088 0.0854
2.500 0.5929 0.01453 0.00750 -0.0234 0.5007 1.0000
2.750 0.6183 0.01460 0.00748 -0.0228 0.4963 1.0000
3.000 0.6435 0.01468 0.00741 -0.0221 0.4929 1.0000
3.250 0.6691 0.01490 0.00766 -0.0219 0.4879 1.0000
3.500 0.6946 0.01504 0.00780 -0.0215 0.4831 1.0000
3.750 0.7199 0.01511 0.00779 -0.0209 0.4795 1.0000
4.000 0.7451 0.01528 0.00790 -0.0204 0.4759 1.0000
4.250 0.7705 0.01549 0.00820 -0.0202 0.4702 1.0000
4.500 0.7958 0.01557 0.00827 -0.0196 0.4659 1.0000
4.750 0.8210 0.01563 0.00823 -0.0189 0.4626 1.0000
5.000 0.8461 0.01589 0.00862 -0.0188 0.4566 1.0000
5.250 0.8713 0.01598 0.00872 -0.0183 0.4518 1.0000
5.500 0.8966 0.01601 0.00870 -0.0176 0.4482 1.0000
5.750 0.9214 0.01626 0.00908 -0.0173 0.4421 1.0000
6.000 0.9465 0.01631 0.00916 -0.0168 0.4368 1.0000
6.250 0.9719 0.01631 0.00909 -0.0162 0.4329 1.0000
6.500 0.9962 0.01657 0.00955 -0.0159 0.4257 1.0000
6.750 1.0215 0.01656 0.00953 -0.0153 0.4205 1.0000
7.000 1.0460 0.01658 0.00964 -0.0149 0.4125 1.0000
7.250 1.0714 0.01622 0.00921 -0.0141 0.4039 1.0000
7.500 1.0957 0.01610 0.00922 -0.0136 0.3922 1.0000
7.750 1.1200 0.01603 0.00922 -0.0131 0.3803 1.0000
8.000 1.1441 0.01598 0.00924 -0.0126 0.3666 1.0000
8.250 1.1678 0.01604 0.00936 -0.0121 0.3512 1.0000
8.500 1.1906 0.01621 0.00951 -0.0116 0.3325 1.0000
8.750 1.2122 0.01658 0.00990 -0.0111 0.3067 1.0000
9.000 1.2307 0.01734 0.01053 -0.0105 0.2698 1.0000
9.250 1.2412 0.01897 0.01184 -0.0098 0.2157 1.0000
9.500 1.2458 0.02108 0.01362 -0.0088 0.1609 1.0000
9.750 1.2282 0.02488 0.01689 -0.0075 0.0848 1.0000
10.000 1.2078 0.02825 0.02012 -0.0058 0.0543 1.0000
10.250 1.1938 0.03146 0.02324 -0.0046 0.0425 1.0000
10.500 1.1880 0.03426 0.02606 -0.0040 0.0385 1.0000
10.750 1.1836 0.03707 0.02896 -0.0035 0.0357 1.0000
11.000 1.1818 0.03972 0.03172 -0.0032 0.0343 1.0000
11.250 1.1787 0.04257 0.03470 -0.0030 0.0332 1.0000
11.500 1.1737 0.04567 0.03793 -0.0030 0.0321 1.0000
11.750 1.1668 0.04906 0.04143 -0.0032 0.0309 1.0000
12.000 1.1608 0.05241 0.04490 -0.0034 0.0303 1.0000
12.250 1.1511 0.05630 0.04889 -0.0038 0.0293 1.0000
12.500 1.1438 0.06010 0.05281 -0.0044 0.0290 1.0000
12.750 1.1348 0.06419 0.05698 -0.0051 0.0284 1.0000
13.000 1.1292 0.06788 0.06074 -0.0056 0.0281 1.0000
13.250 1.1231 0.07144 0.06434 -0.0058 0.0274 1.0000
13.500 1.1257 0.07401 0.06696 -0.0058 0.0271 1.0000
13.750 1.1300 0.07624 0.06925 -0.0055 0.0268 1.0000
14.000 1.1370 0.07800 0.07105 -0.0048 0.0265 1.0000
14.250 1.1455 0.07957 0.07267 -0.0041 0.0259 1.0000
14.500 1.1541 0.08118 0.07433 -0.0034 0.0249 1.0000
14.750 1.1665 0.08214 0.07534 -0.0021 0.0246 1.0000
15.000 1.1784 0.08325 0.07648 -0.0009 0.0237 1.0000
15.250 1.1930 0.08405 0.07736 0.0007 0.0235 1.0000
15.500 1.2057 0.08536 0.07879 0.0021 0.0235 1.0000
15.750 1.2156 0.08726 0.08083 0.0031 0.0234 1.0000
16.000 1.2227 0.08974 0.08351 0.0036 0.0239 1.0000
16.250 1.2256 0.09294 0.08693 0.0037 0.0243 1.0000
16.500 1.2249 0.09678 0.09100 0.0033 0.0249 1.0000
16.750 1.2204 0.10126 0.09570 0.0023 0.0254 1.0000
17.000 1.2130 0.10624 0.10089 0.0008 0.0258 1.0000
17.250 1.2035 0.11159 0.10642 -0.0012 0.0261 1.0000
17.500 1.1906 0.11753 0.11256 -0.0041 0.0260 1.0000
17.750 1.1797 0.12357 0.11877 -0.0064 0.0267 1.0000
18.000 0.9743 0.12392 0.11973 -0.0066 0.0265 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 746 AIRFOIL (goe746-il)