GOE 744 AIRFOIL (goe744-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 744 AIRFOIL (goe744-il) Reynolds number: 1,000,000 Max Cl/Cd: 128.37 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe744-il-1000000.txt Download as CSV file: xf-goe744-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 744 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.1903 0.09904 0.09596 0.0106 0.4917 0.0244
-8.750 -0.1925 0.09463 0.09155 0.0080 0.4906 0.0245
-8.500 -0.1909 0.09053 0.08745 0.0060 0.4891 0.0245
-8.250 -0.1849 0.08659 0.08350 0.0053 0.4873 0.0247
-8.000 -0.1733 0.08390 0.08080 0.0050 0.4849 0.0248
-7.750 -0.1655 0.08107 0.07795 0.0042 0.4824 0.0249
-7.500 -0.1585 0.07822 0.07511 0.0033 0.4812 0.0250
-7.250 -0.1525 0.07539 0.07229 0.0024 0.4801 0.0252
-7.000 -0.1489 0.07236 0.06928 0.0013 0.4789 0.0253
-6.750 -0.1463 0.06947 0.06640 0.0003 0.4774 0.0256
-6.500 -0.1481 0.06655 0.06351 -0.0008 0.4760 0.0258
-6.250 -0.1515 0.06365 0.06062 -0.0015 0.4746 0.0260
-6.000 -0.1469 0.06029 0.05725 -0.0032 0.4731 0.0264
-5.750 -0.1406 0.05669 0.05363 -0.0051 0.4716 0.0269
-5.500 -0.1393 0.04874 0.04557 -0.0117 0.4709 0.0281
-5.000 -0.1198 0.05065 0.04708 -0.0185 0.4702 0.0284
-4.750 -0.1006 0.04879 0.04516 -0.0188 0.4679 0.0286
-4.500 -0.0803 0.04704 0.04335 -0.0192 0.4649 0.0288
-4.250 -0.0593 0.04506 0.04131 -0.0196 0.4630 0.0290
-4.000 -0.0375 0.04303 0.03923 -0.0199 0.4616 0.0294
-3.750 -0.0237 0.03078 0.02642 -0.0207 0.4611 0.0324
-3.500 -0.0003 0.02983 0.02545 -0.0204 0.4592 0.0326
-3.250 0.0233 0.02860 0.02415 -0.0201 0.4571 0.0328
-3.000 0.0473 0.02741 0.02289 -0.0197 0.4549 0.0331
-2.750 0.0598 0.02039 0.01508 -0.0167 0.4534 0.0315
-2.500 0.0809 0.01767 0.01198 -0.0153 0.4513 0.0305
-2.250 0.1048 0.01554 0.00943 -0.0141 0.4488 0.0302
-2.000 0.1316 0.01459 0.00827 -0.0136 0.4458 0.0304
-1.750 0.1594 0.01382 0.00737 -0.0132 0.4442 0.0306
-1.500 0.1876 0.01318 0.00663 -0.0129 0.4420 0.0309
-1.250 0.2161 0.01266 0.00603 -0.0127 0.4397 0.0312
-1.000 0.2447 0.01225 0.00554 -0.0125 0.4372 0.0315
-0.750 0.2735 0.01195 0.00516 -0.0123 0.4344 0.0318
-0.500 0.3020 0.01166 0.00480 -0.0122 0.4312 0.0323
-0.250 0.3301 0.01125 0.00439 -0.0120 0.4283 0.0331
0.000 0.3588 0.01105 0.00421 -0.0120 0.4263 0.0337
0.250 0.3875 0.01085 0.00402 -0.0118 0.4236 0.0343
0.500 0.4161 0.01068 0.00384 -0.0117 0.4207 0.0350
0.750 0.4446 0.01053 0.00367 -0.0116 0.4175 0.0357
1.000 0.4730 0.01046 0.00355 -0.0115 0.4139 0.0363
1.250 0.5007 0.01024 0.00334 -0.0113 0.4111 0.0372
1.500 0.5291 0.01010 0.00324 -0.0112 0.4084 0.0382
1.750 0.5575 0.01002 0.00317 -0.0111 0.4053 0.0394
2.000 0.5860 0.00998 0.00312 -0.0111 0.4020 0.0409
2.250 0.6137 0.00990 0.00303 -0.0109 0.3986 0.0427
2.500 0.6418 0.00990 0.00302 -0.0108 0.3951 0.0447
2.750 0.6702 0.00985 0.00300 -0.0108 0.3923 0.0469
3.000 0.6980 0.00978 0.00295 -0.0106 0.3888 0.0506
3.250 0.7259 0.00976 0.00293 -0.0106 0.3851 0.0554
3.500 0.7533 0.00977 0.00295 -0.0104 0.3812 0.0657
3.750 0.7788 0.00958 0.00308 -0.0100 0.3782 0.1902
4.000 0.8066 0.00957 0.00315 -0.0100 0.3752 0.2224
4.250 0.8342 0.00958 0.00322 -0.0099 0.3716 0.2506
4.500 0.8594 0.00948 0.00336 -0.0096 0.3677 0.3694
4.750 0.9243 0.00848 0.00378 -0.0175 0.3625 0.9755
5.000 0.9805 0.00873 0.00400 -0.0233 0.3581 0.9848
5.250 1.0412 0.00900 0.00421 -0.0301 0.3529 0.9925
5.500 1.1051 0.00914 0.00428 -0.0378 0.3473 0.9973
5.750 1.1439 0.00920 0.00434 -0.0402 0.3439 0.9992
6.000 1.1776 0.00931 0.00443 -0.0416 0.3399 1.0000
6.250 1.2039 0.00949 0.00458 -0.0415 0.3357 1.0000
6.500 1.2300 0.00970 0.00476 -0.0415 0.3314 1.0000
6.750 1.2563 0.00982 0.00489 -0.0413 0.3282 1.0000
7.000 1.2824 0.00999 0.00505 -0.0412 0.3241 1.0000
7.250 1.3079 0.01023 0.00526 -0.0412 0.3196 1.0000
7.500 1.3333 0.01047 0.00548 -0.0410 0.3156 1.0000
7.750 1.3590 0.01063 0.00566 -0.0409 0.3124 1.0000
8.000 1.3842 0.01084 0.00588 -0.0408 0.3086 1.0000
8.250 1.4088 0.01113 0.00614 -0.0406 0.3045 1.0000
8.500 1.4329 0.01145 0.00645 -0.0404 0.3004 1.0000
8.750 1.4577 0.01164 0.00667 -0.0402 0.2975 1.0000
9.000 1.4818 0.01191 0.00694 -0.0400 0.2939 1.0000
9.250 1.5048 0.01226 0.00728 -0.0397 0.2899 1.0000
9.500 1.5265 0.01270 0.00771 -0.0393 0.2854 1.0000
9.750 1.5499 0.01294 0.00798 -0.0390 0.2828 1.0000
10.000 1.5720 0.01326 0.00833 -0.0386 0.2796 1.0000
10.250 1.5926 0.01369 0.00876 -0.0381 0.2762 1.0000
10.500 1.6107 0.01426 0.00932 -0.0374 0.2724 1.0000
10.750 1.6289 0.01476 0.00985 -0.0366 0.2695 1.0000
11.000 1.6474 0.01519 0.01032 -0.0359 0.2672 1.0000
11.250 1.6626 0.01583 0.01100 -0.0350 0.2639 1.0000
11.500 1.6689 0.01715 0.01235 -0.0341 0.2605 1.0000
11.750 1.6491 0.01908 0.01430 -0.0297 0.2572 1.0000
12.000 1.6502 0.02041 0.01567 -0.0277 0.2544 1.0000
12.250 1.6565 0.02157 0.01688 -0.0264 0.2515 1.0000
12.500 1.6591 0.02307 0.01841 -0.0251 0.2487 1.0000
12.750 1.6581 0.02493 0.02028 -0.0238 0.2456 1.0000
13.000 1.6524 0.02720 0.02257 -0.0224 0.2420 1.0000
13.250 1.6573 0.02876 0.02418 -0.0216 0.2394 1.0000
13.500 1.6626 0.03030 0.02577 -0.0209 0.2372 1.0000
13.750 1.6646 0.03214 0.02764 -0.0201 0.2342 1.0000
14.000 1.6632 0.03431 0.02985 -0.0194 0.2317 1.0000
14.250 1.6560 0.03705 0.03259 -0.0187 0.2281 1.0000
14.500 1.6595 0.03891 0.03449 -0.0182 0.2259 1.0000
14.750 1.6652 0.04067 0.03632 -0.0179 0.2241 1.0000
15.000 1.6702 0.04248 0.03817 -0.0177 0.2216 1.0000
15.250 1.6705 0.04481 0.04054 -0.0174 0.2190 1.0000
15.500 1.6696 0.04724 0.04299 -0.0172 0.2161 1.0000
15.750 1.6651 0.05007 0.04583 -0.0170 0.2129 1.0000
16.000 1.6717 0.05185 0.04768 -0.0170 0.2111 1.0000
16.250 1.6769 0.05380 0.04968 -0.0170 0.2087 1.0000
16.500 1.6775 0.05624 0.05216 -0.0170 0.2058 1.0000
16.750 1.6772 0.05877 0.05471 -0.0170 0.2029 1.0000
17.000 1.6701 0.06207 0.05802 -0.0172 0.1993 1.0000
17.250 1.6759 0.06400 0.06002 -0.0173 0.1972 1.0000
17.500 1.6805 0.06609 0.06217 -0.0174 0.1947 1.0000
17.750 1.6779 0.06904 0.06515 -0.0177 0.1914 1.0000
18.000 1.6755 0.07192 0.06805 -0.0180 0.1882 1.0000
18.250 1.6699 0.07524 0.07141 -0.0184 0.1849 1.0000
18.500 1.6738 0.07749 0.07372 -0.0187 0.1821 1.0000
18.750 1.6726 0.08033 0.07659 -0.0191 0.1782 1.0000
19.000 1.6613 0.08444 0.08072 -0.0198 0.1739 1.0000
19.250 1.6616 0.08713 0.08346 -0.0203 0.1709 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 744 AIRFOIL (goe744-il)