20-32C AIRFOIL (2032c-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: 20-32C AIRFOIL (2032c-il) Reynolds number: 200,000 Max Cl/Cd: 75.32 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-2032c-il-200000.txt Download as CSV file: xf-2032c-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: 20-32C 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 -0.1581 0.09761 0.09467 -0.0262 1.0000 0.0240
-8.500 -0.1536 0.09468 0.09178 -0.0262 1.0000 0.0244
-8.250 -0.2407 0.10671 0.10359 -0.0219 1.0000 0.0232
-8.000 -0.2303 0.10247 0.09935 -0.0213 1.0000 0.0237
-7.750 -0.2229 0.09947 0.09639 -0.0215 1.0000 0.0241
-7.500 -0.2174 0.09690 0.09388 -0.0218 1.0000 0.0245
-7.250 -0.2150 0.09475 0.09180 -0.0215 1.0000 0.0249
-7.000 -0.2194 0.09328 0.09042 -0.0198 1.0000 0.0252
-6.750 -0.2314 0.09255 0.08980 -0.0166 1.0000 0.0254
-6.500 -0.2233 0.09009 0.08738 -0.0186 0.9966 0.0259
-6.250 -0.1900 0.08553 0.08281 -0.0268 0.9890 0.0270
-6.000 -0.1523 0.08088 0.07814 -0.0361 0.9817 0.0284
-5.750 -0.0972 0.07684 0.07404 -0.0523 0.9713 0.0304
-5.500 -0.0417 0.07273 0.06982 -0.0675 0.9588 0.0309
-5.250 -0.0078 0.06691 0.06397 -0.0749 0.9463 0.0313
-5.000 -0.0001 0.06252 0.05960 -0.0718 0.9327 0.0327
-4.750 0.0224 0.05945 0.05649 -0.0735 0.9160 0.0347
-4.500 0.0512 0.05621 0.05317 -0.0777 0.8984 0.0368
-4.250 0.0843 0.05296 0.04980 -0.0831 0.8792 0.0395
-4.000 0.1492 0.05014 0.04662 -0.0960 0.8598 0.0421
-3.750 0.1656 0.04499 0.04141 -0.0971 0.8423 0.0435
-3.500 0.1832 0.04264 0.03899 -0.0966 0.8235 0.0454
-3.250 0.2121 0.04019 0.03639 -0.0987 0.8040 0.0491
-3.000 0.2608 0.03631 0.03213 -0.1058 0.7860 0.0569
-2.500 0.3282 0.03126 0.02652 -0.1108 0.7482 0.0704
-2.250 0.3508 0.02964 0.02483 -0.1106 0.7284 0.0748
-2.000 0.3864 0.02736 0.02219 -0.1130 0.7101 0.0866
-1.750 0.4186 0.02566 0.02020 -0.1145 0.6908 0.1006
-1.500 0.4502 0.02414 0.01836 -0.1156 0.6714 0.1147
-1.250 0.4799 0.02282 0.01678 -0.1162 0.6523 0.1294
-1.000 0.5252 0.01961 0.01249 -0.1175 0.6356 0.0830
-0.750 0.5588 0.01688 0.00921 -0.1181 0.6175 0.0678
-0.500 0.5882 0.01607 0.00803 -0.1180 0.5987 0.0682
-0.250 0.6169 0.01548 0.00711 -0.1178 0.5804 0.0690
0.000 0.6454 0.01479 0.00618 -0.1177 0.5617 0.0692
0.250 0.6735 0.01417 0.00542 -0.1176 0.5430 0.0709
0.500 0.7011 0.01388 0.00503 -0.1174 0.5249 0.0738
0.750 0.7285 0.01381 0.00484 -0.1172 0.5078 0.0802
1.000 0.7562 0.01368 0.00461 -0.1170 0.4917 0.0855
1.250 0.7840 0.01363 0.00449 -0.1169 0.4758 0.0987
1.500 0.8117 0.01296 0.00461 -0.1172 0.4609 0.4068
1.750 0.8328 0.01219 0.00455 -0.1155 0.4480 1.0000
2.000 0.8599 0.01248 0.00463 -0.1153 0.4349 1.0000
2.250 0.8869 0.01276 0.00475 -0.1151 0.4228 1.0000
2.500 0.9139 0.01307 0.00492 -0.1149 0.4121 1.0000
2.750 0.9407 0.01341 0.00511 -0.1147 0.4027 1.0000
3.000 0.9675 0.01371 0.00529 -0.1145 0.3934 1.0000
3.250 0.9943 0.01403 0.00553 -0.1143 0.3848 1.0000
3.500 1.0209 0.01440 0.00576 -0.1141 0.3776 1.0000
3.750 1.0478 0.01471 0.00607 -0.1139 0.3703 1.0000
4.000 1.0743 0.01507 0.00634 -0.1137 0.3642 1.0000
4.250 1.1009 0.01544 0.00668 -0.1136 0.3582 1.0000
4.500 1.1274 0.01577 0.00699 -0.1134 0.3521 1.0000
4.750 1.1538 0.01622 0.00737 -0.1132 0.3470 1.0000
5.000 1.1801 0.01654 0.00776 -0.1130 0.3414 1.0000
5.250 1.2065 0.01693 0.00813 -0.1128 0.3368 1.0000
5.500 1.2328 0.01743 0.00859 -0.1127 0.3326 1.0000
5.750 1.2583 0.01767 0.00896 -0.1123 0.3257 1.0000
6.000 1.2831 0.01796 0.00919 -0.1120 0.3174 1.0000
6.250 1.3078 0.01812 0.00942 -0.1115 0.3087 1.0000
6.500 1.3324 0.01841 0.00974 -0.1111 0.3009 1.0000
6.750 1.3561 0.01862 0.00993 -0.1106 0.2907 1.0000
7.000 1.3795 0.01867 0.01008 -0.1100 0.2783 1.0000
7.250 1.4031 0.01880 0.01030 -0.1095 0.2656 1.0000
7.500 1.4266 0.01899 0.01056 -0.1089 0.2513 1.0000
7.750 1.4492 0.01924 0.01077 -0.1083 0.2235 1.0000
8.000 1.4701 0.01987 0.01126 -0.1074 0.1848 1.0000
8.250 1.4649 0.02361 0.01395 -0.1040 0.0406 1.0000
8.500 1.4793 0.02496 0.01535 -0.1023 0.0295 1.0000
8.750 1.4948 0.02610 0.01665 -0.1006 0.0266 1.0000
9.000 1.5071 0.02746 0.01825 -0.0986 0.0241 1.0000
9.250 1.5181 0.02881 0.01982 -0.0965 0.0232 1.0000
9.500 1.5253 0.03028 0.02155 -0.0941 0.0227 1.0000
9.750 1.5265 0.03194 0.02343 -0.0909 0.0224 1.0000
10.000 1.5248 0.03397 0.02567 -0.0881 0.0222 1.0000
10.250 1.5203 0.03647 0.02838 -0.0858 0.0220 1.0000
10.500 1.5136 0.03951 0.03161 -0.0843 0.0219 1.0000
10.750 1.5051 0.04315 0.03543 -0.0835 0.0218 1.0000
11.000 1.4954 0.04733 0.03979 -0.0835 0.0218 1.0000
11.250 1.4848 0.05197 0.04459 -0.0840 0.0218 1.0000
11.500 1.4738 0.05687 0.04964 -0.0847 0.0218 1.0000
11.750 1.4626 0.06186 0.05476 -0.0855 0.0218 1.0000
12.000 1.4523 0.06673 0.05976 -0.0862 0.0217 1.0000
12.250 1.4431 0.07139 0.06452 -0.0867 0.0216 1.0000
12.500 1.4354 0.07576 0.06898 -0.0870 0.0213 1.0000
12.750 1.4295 0.07975 0.07305 -0.0871 0.0209 1.0000
13.000 1.4266 0.08325 0.07661 -0.0869 0.0206 1.0000
13.250 1.4286 0.08579 0.07922 -0.0858 0.0205 1.0000
13.500 1.4369 0.08708 0.08054 -0.0835 0.0208 1.0000
13.750 1.4509 0.08742 0.08092 -0.0801 0.0215 1.0000
|
Polar data table (+)
Polar graphs
<< Back to 20-32C AIRFOIL (2032c-il)