GOE 370 AIRFOIL (goe370-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 370 AIRFOIL (goe370-il) Reynolds number: 100,000 Max Cl/Cd: 67.47 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe370-il-100000-n5.txt Download as CSV file: xf-goe370-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 370 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
------ -------- --------- --------- -------- -------- --------
-7.250 -0.3040 0.09758 0.09316 -0.0274 1.0000 0.0272
-7.000 -0.3117 0.09616 0.09184 -0.0251 1.0000 0.0274
-6.750 -0.3194 0.09479 0.09056 -0.0229 1.0000 0.0275
-6.500 -0.3022 0.09110 0.08688 -0.0269 0.9947 0.0282
-6.250 -0.2782 0.08693 0.08269 -0.0328 0.9875 0.0290
-6.000 -0.2541 0.08293 0.07868 -0.0387 0.9797 0.0297
-5.750 -0.2289 0.07912 0.07481 -0.0445 0.9715 0.0315
-5.500 -0.1917 0.07558 0.07119 -0.0553 0.9623 0.0353
-5.250 -0.1427 0.07153 0.06699 -0.0687 0.9552 0.0361
-5.000 -0.1109 0.06734 0.06271 -0.0748 0.9464 0.0362
-4.750 -0.0885 0.06187 0.05725 -0.0774 0.9417 0.0370
-4.250 -0.0087 0.05143 0.04645 -0.0909 0.9278 0.0257
-4.000 0.0239 0.04695 0.04184 -0.0955 0.9187 0.0260
-3.750 0.0502 0.04478 0.03959 -0.0976 0.9141 0.0315
-3.500 0.0851 0.04084 0.03543 -0.1018 0.9057 0.0317
-3.250 0.1275 0.03633 0.03060 -0.1071 0.9011 0.0319
-3.000 0.1664 0.03161 0.02543 -0.1109 0.8941 0.0350
-2.750 0.2020 0.02795 0.02135 -0.1136 0.8887 0.0385
-2.500 0.2352 0.02512 0.01805 -0.1151 0.8828 0.0427
-2.250 0.2680 0.02220 0.01444 -0.1163 0.8766 0.0493
-2.000 0.3008 0.02075 0.01257 -0.1173 0.8717 0.0536
-1.750 0.3295 0.01960 0.01091 -0.1173 0.8641 0.0598
-1.500 0.3622 0.01827 0.00915 -0.1180 0.8593 0.0603
-1.250 0.3893 0.01740 0.00805 -0.1178 0.8516 0.0612
-1.000 0.4200 0.01665 0.00715 -0.1182 0.8461 0.0633
-0.750 0.4471 0.01620 0.00661 -0.1180 0.8383 0.0673
-0.500 0.4774 0.01571 0.00598 -0.1182 0.8324 0.0751
-0.250 0.5046 0.01545 0.00576 -0.1180 0.8247 0.0927
0.000 0.5349 0.01529 0.00556 -0.1182 0.8186 0.1241
0.250 0.5613 0.01530 0.00555 -0.1179 0.8103 0.1592
0.500 0.5905 0.01513 0.00548 -0.1182 0.8040 0.2016
0.750 0.6160 0.01506 0.00544 -0.1177 0.7955 0.2246
1.000 0.6449 0.01491 0.00535 -0.1178 0.7892 0.2569
1.250 0.6695 0.01478 0.00544 -0.1173 0.7802 0.3338
1.500 0.7101 0.01363 0.00534 -0.1199 0.7740 1.0000
1.750 0.7351 0.01380 0.00543 -0.1192 0.7650 1.0000
2.000 0.7611 0.01395 0.00551 -0.1187 0.7567 1.0000
2.250 0.7876 0.01408 0.00558 -0.1182 0.7480 1.0000
2.500 0.8120 0.01421 0.00572 -0.1174 0.7360 1.0000
2.750 0.8367 0.01430 0.00579 -0.1164 0.7228 1.0000
3.000 0.8612 0.01441 0.00591 -0.1155 0.7097 1.0000
3.250 0.8860 0.01455 0.00609 -0.1147 0.6979 1.0000
3.500 0.9111 0.01470 0.00633 -0.1140 0.6868 1.0000
3.750 0.9363 0.01484 0.00655 -0.1133 0.6757 1.0000
4.000 0.9602 0.01503 0.00687 -0.1124 0.6632 1.0000
4.250 0.9843 0.01523 0.00725 -0.1115 0.6508 1.0000
4.500 1.0085 0.01543 0.00762 -0.1107 0.6381 1.0000
4.750 1.0325 0.01560 0.00798 -0.1097 0.6232 1.0000
5.000 1.0471 0.01552 0.00752 -0.1058 0.5446 1.0000
5.250 1.0590 0.01622 0.00776 -0.1022 0.4418 1.0000
5.500 1.0527 0.01886 0.00878 -0.0967 0.2054 1.0000
5.750 1.0508 0.02192 0.01070 -0.0925 0.0383 1.0000
6.000 1.0656 0.02320 0.01211 -0.0902 0.0286 1.0000
6.250 1.0793 0.02454 0.01366 -0.0879 0.0252 1.0000
6.500 1.0927 0.02582 0.01518 -0.0854 0.0230 1.0000
6.750 1.1039 0.02719 0.01671 -0.0828 0.0202 1.0000
7.000 1.1099 0.02902 0.01862 -0.0797 0.0181 1.0000
7.250 1.1182 0.03112 0.02075 -0.0768 0.0174 1.0000
7.500 1.1364 0.03298 0.02270 -0.0752 0.0169 1.0000
7.750 1.1619 0.03524 0.02517 -0.0744 0.0164 1.0000
8.000 1.1876 0.03743 0.02761 -0.0737 0.0151 1.0000
8.250 1.2120 0.03988 0.03035 -0.0728 0.0138 1.0000
8.500 1.2350 0.04290 0.03370 -0.0718 0.0136 1.0000
8.750 1.2528 0.04612 0.03732 -0.0701 0.0135 1.0000
9.000 1.2653 0.04946 0.04109 -0.0679 0.0136 1.0000
9.250 1.2732 0.05291 0.04495 -0.0653 0.0137 1.0000
9.500 1.2765 0.05641 0.04884 -0.0624 0.0139 1.0000
9.750 1.2752 0.05991 0.05272 -0.0594 0.0141 1.0000
10.000 1.2694 0.06320 0.05632 -0.0560 0.0143 1.0000
10.250 1.2590 0.06643 0.05984 -0.0525 0.0145 1.0000
10.500 1.2462 0.06975 0.06342 -0.0495 0.0147 1.0000
10.750 1.2312 0.07339 0.06731 -0.0472 0.0148 1.0000
11.000 1.2154 0.07729 0.07143 -0.0457 0.0150 1.0000
11.250 1.1993 0.08146 0.07581 -0.0450 0.0151 1.0000
11.500 1.1816 0.08616 0.08071 -0.0452 0.0152 1.0000
11.750 1.1638 0.09126 0.08601 -0.0463 0.0153 1.0000
12.000 1.1454 0.09691 0.09182 -0.0482 0.0154 1.0000
12.250 1.1276 0.10301 0.09808 -0.0511 0.0155 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 370 AIRFOIL (goe370-il)