GOE 81 AIRFOIL (goe81-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 81 AIRFOIL (goe81-il) Reynolds number: 500,000 Max Cl/Cd: 114.08 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe81-il-500000.txt Download as CSV file: xf-goe81-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 81 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.2506 0.10456 0.10234 -0.0316 1.0000 0.0170
-8.750 -0.2482 0.10238 0.10020 -0.0314 1.0000 0.0174
-8.500 -0.2478 0.10041 0.09828 -0.0308 0.9995 0.0177
-8.250 -0.2287 0.09660 0.09446 -0.0355 0.9953 0.0185
-8.000 -0.2074 0.09290 0.09076 -0.0424 0.9869 0.0191
-7.750 -0.1855 0.08920 0.08705 -0.0517 0.9726 0.0193
-7.500 -0.0781 0.06777 0.06560 -0.0598 0.9217 0.0198
-7.250 -0.0713 0.06503 0.06280 -0.0604 0.9078 0.0201
-7.000 -0.0648 0.06235 0.06007 -0.0611 0.8953 0.0204
-6.750 -0.0578 0.05958 0.05725 -0.0624 0.8837 0.0209
-6.500 -0.0471 0.05649 0.05413 -0.0648 0.8722 0.0215
-6.250 -0.0323 0.05319 0.05078 -0.0689 0.8615 0.0226
-6.000 -0.0538 0.06550 0.06301 -0.0756 0.8865 0.0221
-5.750 -0.0111 0.06164 0.05901 -0.0879 0.8745 0.0229
-5.500 0.0228 0.05787 0.05508 -0.0950 0.8639 0.0230
-5.250 0.0385 0.05418 0.05135 -0.0960 0.8535 0.0232
-5.000 0.0546 0.05160 0.04875 -0.0961 0.8429 0.0234
-4.750 0.0755 0.04922 0.04630 -0.0975 0.8328 0.0238
-4.500 0.1008 0.04676 0.04376 -0.1000 0.8217 0.0244
-4.250 0.1293 0.04414 0.04104 -0.1031 0.8096 0.0252
-4.000 0.1613 0.04142 0.03818 -0.1065 0.7981 0.0265
-3.750 0.2121 0.03841 0.03481 -0.1126 0.7876 0.0275
-3.500 0.2377 0.03495 0.03122 -0.1146 0.7761 0.0278
-3.250 0.2599 0.03297 0.02917 -0.1154 0.7643 0.0281
-3.000 0.2857 0.03122 0.02732 -0.1164 0.7537 0.0286
-2.750 0.3139 0.02947 0.02543 -0.1176 0.7441 0.0294
-2.500 0.3440 0.02766 0.02348 -0.1189 0.7350 0.0306
-2.250 0.3848 0.02603 0.02141 -0.1199 0.7272 0.0329
-2.000 0.4125 0.02309 0.01832 -0.1213 0.7182 0.0335
-1.750 0.4386 0.02181 0.01695 -0.1219 0.7090 0.0341
-1.500 0.4657 0.02078 0.01581 -0.1223 0.6997 0.0351
-1.250 0.4946 0.01965 0.01454 -0.1227 0.6911 0.0368
-1.000 0.5275 0.01944 0.01393 -0.1223 0.6831 0.0398
-0.750 0.5555 0.01695 0.01132 -0.1233 0.6760 0.0413
-0.500 0.5830 0.01618 0.01048 -0.1235 0.6684 0.0428
-0.250 0.6117 0.01562 0.00976 -0.1235 0.6609 0.0465
0.000 0.6406 0.01463 0.00855 -0.1237 0.6523 0.0509
0.250 0.6682 0.01411 0.00798 -0.1238 0.6432 0.0542
1.750 0.8372 0.01108 0.00416 -0.1225 0.5860 0.0477
2.000 0.8649 0.01053 0.00358 -0.1223 0.5739 0.0447
2.250 0.8922 0.01029 0.00329 -0.1220 0.5595 0.0438
2.500 0.9191 0.01018 0.00315 -0.1218 0.5422 0.0448
2.750 0.9457 0.01014 0.00305 -0.1215 0.5216 0.0458
3.000 0.9715 0.01019 0.00298 -0.1211 0.4939 0.0457
3.250 0.9968 0.01035 0.00299 -0.1206 0.4671 0.0464
3.500 1.0223 0.01054 0.00307 -0.1202 0.4482 0.0477
4.000 1.0743 0.01087 0.00329 -0.1196 0.4221 0.0585
4.250 1.0963 0.00961 0.00363 -0.1188 0.4127 1.0000
4.500 1.1220 0.00987 0.00380 -0.1184 0.4032 1.0000
4.750 1.1482 0.01007 0.00397 -0.1182 0.3933 1.0000
5.000 1.1739 0.01032 0.00416 -0.1178 0.3824 1.0000
5.250 1.1993 0.01057 0.00436 -0.1175 0.3720 1.0000
5.500 1.2252 0.01078 0.00456 -0.1172 0.3609 1.0000
5.750 1.2506 0.01102 0.00476 -0.1168 0.3495 1.0000
6.000 1.2757 0.01128 0.00498 -0.1164 0.3374 1.0000
6.250 1.3005 0.01156 0.00522 -0.1160 0.3248 1.0000
6.500 1.3250 0.01186 0.00548 -0.1155 0.3082 1.0000
6.750 1.3485 0.01223 0.00577 -0.1149 0.2865 1.0000
7.000 1.3709 0.01270 0.00612 -0.1142 0.2642 1.0000
7.250 1.3928 0.01321 0.00651 -0.1134 0.2395 1.0000
7.500 1.4134 0.01381 0.00699 -0.1124 0.2125 1.0000
7.750 1.4326 0.01452 0.00753 -0.1112 0.1821 1.0000
8.000 1.4499 0.01538 0.00818 -0.1098 0.1501 1.0000
8.250 1.4690 0.01604 0.00876 -0.1086 0.1356 1.0000
8.500 1.4884 0.01665 0.00934 -0.1074 0.1263 1.0000
8.750 1.5085 0.01716 0.00987 -0.1063 0.1192 1.0000
9.000 1.5273 0.01775 0.01045 -0.1051 0.1126 1.0000
9.250 1.5469 0.01824 0.01097 -0.1040 0.1063 1.0000
9.500 1.5639 0.01889 0.01160 -0.1025 0.0964 1.0000
9.750 1.5785 0.01965 0.01223 -0.1007 0.0700 1.0000
10.000 1.5852 0.02074 0.01321 -0.0976 0.0562 1.0000
10.250 1.5962 0.02159 0.01409 -0.0952 0.0489 1.0000
10.500 1.6052 0.02260 0.01511 -0.0926 0.0400 1.0000
11.000 1.6129 0.02549 0.01789 -0.0869 0.0200 1.0000
11.250 1.6185 0.02694 0.01941 -0.0846 0.0181 1.0000
11.500 1.6255 0.02835 0.02092 -0.0826 0.0172 1.0000
11.750 1.6310 0.02994 0.02262 -0.0807 0.0165 1.0000
12.000 1.6347 0.03175 0.02455 -0.0788 0.0159 1.0000
12.250 1.6362 0.03385 0.02676 -0.0770 0.0154 1.0000
12.500 1.6348 0.03632 0.02934 -0.0752 0.0149 1.0000
12.750 1.6302 0.03923 0.03238 -0.0736 0.0145 1.0000
13.000 1.6227 0.04260 0.03589 -0.0723 0.0142 1.0000
13.250 1.6222 0.04535 0.03875 -0.0714 0.0140 1.0000
13.500 1.6197 0.04839 0.04192 -0.0706 0.0138 1.0000
13.750 1.6159 0.05168 0.04533 -0.0701 0.0136 1.0000
14.000 1.6109 0.05522 0.04900 -0.0697 0.0134 1.0000
14.250 1.6050 0.05896 0.05285 -0.0694 0.0132 1.0000
14.500 1.5984 0.06291 0.05693 -0.0694 0.0129 1.0000
14.750 1.5910 0.06708 0.06122 -0.0696 0.0127 1.0000
15.000 1.5831 0.07141 0.06566 -0.0699 0.0125 1.0000
15.250 1.5746 0.07592 0.07028 -0.0704 0.0123 1.0000
15.500 1.5659 0.08059 0.07505 -0.0711 0.0121 1.0000
15.750 1.5571 0.08534 0.07991 -0.0719 0.0119 1.0000
16.000 1.5480 0.09018 0.08484 -0.0728 0.0118 1.0000
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Polar data table (+)
Polar graphs
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