GOE 375 AIRFOIL (goe375-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 375 AIRFOIL (goe375-il) Reynolds number: 1,000,000 Max Cl/Cd: 109.89 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe375-il-1000000.txt Download as CSV file: xf-goe375-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 375 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3747 0.10029 0.09870 -0.0076 1.0000 0.0081
-8.500 -0.3704 0.09730 0.09573 -0.0091 1.0000 0.0082
-8.250 -0.3651 0.09450 0.09293 -0.0099 1.0000 0.0082
-8.000 -0.2805 0.07725 0.07582 -0.0157 1.0000 0.0084
-7.750 -0.2779 0.07426 0.07284 -0.0158 1.0000 0.0085
-7.500 -0.2770 0.07128 0.06988 -0.0161 1.0000 0.0086
-7.250 -0.2774 0.06848 0.06710 -0.0161 1.0000 0.0087
-7.000 -0.2807 0.06599 0.06465 -0.0156 1.0000 0.0088
-6.750 -0.2848 0.06346 0.06216 -0.0153 1.0000 0.0089
-6.250 -0.2558 0.05484 0.05352 -0.0248 0.9892 0.0093
-6.000 -0.2318 0.04980 0.04846 -0.0313 0.9802 0.0096
-5.750 -0.2014 0.04439 0.04298 -0.0390 0.9643 0.0101
-5.500 -0.1654 0.03896 0.03743 -0.0471 0.9358 0.0108
-5.250 -0.1427 0.03481 0.03304 -0.0508 0.8903 0.0109
-5.000 -0.1237 0.03109 0.02912 -0.0525 0.8622 0.0109
-4.750 -0.1091 0.02661 0.02449 -0.0535 0.8440 0.0111
-4.500 -0.0926 0.02395 0.02172 -0.0540 0.8284 0.0113
-4.250 -0.0726 0.02168 0.01935 -0.0546 0.8158 0.0119
-4.000 -0.0499 0.01925 0.01678 -0.0553 0.8052 0.0127
-3.750 -0.0199 0.01701 0.01432 -0.0558 0.7958 0.0140
-3.500 0.0077 0.01476 0.01184 -0.0558 0.7865 0.0141
-3.250 0.0319 0.01260 0.00949 -0.0557 0.7768 0.0141
-3.000 0.0530 0.00969 0.00640 -0.0560 0.7666 0.0144
-2.750 0.0751 0.00842 0.00502 -0.0561 0.7528 0.0147
-2.500 0.0992 0.00741 0.00387 -0.0559 0.7368 0.0155
-2.250 0.1299 0.00693 0.00305 -0.0546 0.7187 0.0177
0.000 0.3659 0.01201 0.00574 -0.0522 0.4635 0.0249
0.250 0.3928 0.01234 0.00585 -0.0516 0.4379 0.0276
0.500 0.4196 0.01245 0.00576 -0.0511 0.4158 0.0281
3.250 0.7035 0.00960 0.00221 -0.0460 0.2415 0.0454
3.500 0.7296 0.00962 0.00220 -0.0456 0.2337 0.0411
3.750 0.7560 0.00970 0.00223 -0.0452 0.2267 0.0368
4.000 0.7822 0.00980 0.00232 -0.0449 0.2200 0.0358
4.250 0.8085 0.00983 0.00234 -0.0445 0.2144 0.0358
4.500 0.8344 0.00992 0.00239 -0.0441 0.2077 0.0373
4.750 0.8609 0.00998 0.00246 -0.0438 0.2030 0.0390
5.000 0.8870 0.01011 0.00257 -0.0435 0.1970 0.0407
5.250 0.9131 0.01024 0.00270 -0.0432 0.1910 0.0430
5.500 0.9393 0.01037 0.00282 -0.0429 0.1847 0.0457
5.750 0.9848 0.00901 0.00323 -0.0477 0.1764 1.0000
6.000 1.0098 0.00922 0.00340 -0.0472 0.1679 1.0000
6.250 1.0349 0.00942 0.00357 -0.0467 0.1598 1.0000
6.500 1.0595 0.00968 0.00378 -0.0461 0.1515 1.0000
6.750 1.0846 0.00987 0.00398 -0.0457 0.1455 1.0000
7.000 1.1089 0.01016 0.00423 -0.0451 0.1376 1.0000
7.250 1.1338 0.01037 0.00445 -0.0446 0.1320 1.0000
7.500 1.1579 0.01068 0.00472 -0.0441 0.1242 1.0000
7.750 1.1825 0.01091 0.00495 -0.0436 0.1182 1.0000
8.000 1.2064 0.01123 0.00526 -0.0430 0.1103 1.0000
8.250 1.2297 0.01161 0.00558 -0.0424 0.0985 1.0000
8.500 1.2428 0.01319 0.00668 -0.0404 0.0346 1.0000
8.750 1.2604 0.01427 0.00768 -0.0389 0.0138 1.0000
9.000 1.2827 0.01475 0.00825 -0.0381 0.0124 1.0000
9.250 1.3038 0.01537 0.00894 -0.0372 0.0111 1.0000
9.500 1.3220 0.01632 0.01004 -0.0357 0.0096 1.0000
9.750 1.3434 0.01685 0.01062 -0.0349 0.0093 1.0000
10.000 1.3639 0.01743 0.01128 -0.0339 0.0087 1.0000
10.250 1.3832 0.01813 0.01207 -0.0327 0.0082 1.0000
10.500 1.4022 0.01880 0.01279 -0.0317 0.0076 1.0000
10.750 1.4180 0.01976 0.01386 -0.0301 0.0071 1.0000
11.000 1.4247 0.02146 0.01575 -0.0274 0.0066 1.0000
11.250 1.4380 0.02243 0.01681 -0.0256 0.0064 1.0000
11.500 1.4519 0.02327 0.01775 -0.0239 0.0062 1.0000
11.750 1.4602 0.02439 0.01897 -0.0215 0.0060 1.0000
12.000 1.4644 0.02543 0.02012 -0.0184 0.0059 1.0000
12.250 1.4656 0.02667 0.02146 -0.0152 0.0057 1.0000
12.500 1.4681 0.02797 0.02285 -0.0126 0.0055 1.0000
12.750 1.4652 0.02980 0.02480 -0.0102 0.0054 1.0000
13.000 1.4618 0.03196 0.02708 -0.0086 0.0054 1.0000
13.250 1.4579 0.03451 0.02975 -0.0078 0.0052 1.0000
13.500 1.4564 0.03714 0.03249 -0.0076 0.0051 1.0000
13.750 1.4551 0.04004 0.03548 -0.0080 0.0049 1.0000
14.000 1.4459 0.04431 0.03989 -0.0093 0.0048 1.0000
14.250 1.4363 0.04918 0.04489 -0.0114 0.0048 1.0000
14.500 1.4233 0.05500 0.05085 -0.0144 0.0048 1.0000
14.750 1.4103 0.06095 0.05693 -0.0173 0.0047 1.0000
15.000 1.3894 0.06811 0.06423 -0.0206 0.0047 1.0000
15.250 1.3734 0.07453 0.07075 -0.0235 0.0046 1.0000
15.500 1.3522 0.08160 0.07795 -0.0265 0.0047 1.0000
15.750 1.3342 0.08839 0.08484 -0.0294 0.0047 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 375 AIRFOIL (goe375-il)