AIRFOIL 3024 9.84% (s3024-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: AIRFOIL 3024 9.84% (s3024-il) Reynolds number: 100,000 Max Cl/Cd: 57.8 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s3024-il-100000.txt Download as CSV file: xf-s3024-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: AIRFOIL 3024 9.84%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2552 0.09723 0.09287 -0.0364 1.0000 0.0914
-9.000 -0.2646 0.09540 0.09112 -0.0354 1.0000 0.0940
-8.750 -0.2822 0.09427 0.09009 -0.0345 1.0000 0.0957
-8.500 -0.3039 0.09352 0.08944 -0.0335 1.0000 0.0965
-8.250 -0.3287 0.09284 0.08886 -0.0325 1.0000 0.0969
-8.000 -0.3518 0.09196 0.08808 -0.0311 1.0000 0.0971
-7.750 -0.3189 0.08574 0.08184 -0.0268 1.0000 0.1004
-7.500 -0.3235 0.08376 0.07991 -0.0244 1.0000 0.1028
-7.250 -0.3343 0.08199 0.07820 -0.0225 1.0000 0.1049
-7.000 -0.3496 0.08031 0.07660 -0.0208 1.0000 0.1068
-6.750 -0.4350 0.08603 0.08207 -0.0232 1.0000 0.1008
-6.500 -0.4397 0.08396 0.08003 -0.0215 1.0000 0.1029
-6.250 -0.4467 0.08156 0.07768 -0.0212 1.0000 0.1062
-6.000 -0.4645 0.07838 0.07421 -0.0331 1.0000 0.1117
-5.750 -0.4588 0.07444 0.07049 -0.0276 1.0000 0.1132
-5.500 -0.4547 0.07221 0.06832 -0.0242 1.0000 0.1157
-5.250 -0.4504 0.06955 0.06563 -0.0240 1.0000 0.1208
-5.000 -0.4234 0.06435 0.06029 -0.0313 0.9949 0.1300
-4.750 -0.3939 0.05984 0.05556 -0.0379 0.9887 0.1430
-4.500 -0.3628 0.05607 0.05160 -0.0429 0.9831 0.1574
-4.250 -0.3357 0.05276 0.04816 -0.0458 0.9768 0.1723
-4.000 -0.2780 0.03975 0.03355 -0.0548 0.9736 0.0814
-3.750 -0.2443 0.03446 0.02754 -0.0563 0.9689 0.0655
-3.500 -0.2096 0.03166 0.02418 -0.0579 0.9635 0.0647
-3.250 -0.1682 0.02927 0.02121 -0.0603 0.9594 0.0643
-3.000 -0.1383 0.02744 0.01894 -0.0603 0.9527 0.0640
-2.750 -0.0992 0.02596 0.01711 -0.0620 0.9476 0.0658
-2.500 -0.0654 0.02501 0.01589 -0.0628 0.9412 0.0696
-2.250 -0.0312 0.02396 0.01491 -0.0640 0.9348 0.0826
-2.000 0.0058 0.02263 0.01384 -0.0656 0.9296 0.1241
-1.750 0.0305 0.02085 0.01345 -0.0656 0.9222 0.3674
-1.500 0.0936 0.01916 0.01343 -0.0708 0.9202 1.0000
-1.250 0.1278 0.01945 0.01334 -0.0723 0.9122 1.0000
-1.000 0.1655 0.01969 0.01327 -0.0743 0.9045 1.0000
-0.750 0.1947 0.01996 0.01332 -0.0748 0.8953 1.0000
-0.500 0.2381 0.02014 0.01325 -0.0778 0.8888 1.0000
-0.250 0.2638 0.02040 0.01335 -0.0776 0.8784 1.0000
0.000 0.3131 0.02045 0.01323 -0.0814 0.8731 1.0000
0.250 0.3363 0.02071 0.01337 -0.0807 0.8620 1.0000
0.500 0.3877 0.02059 0.01312 -0.0846 0.8575 1.0000
0.750 0.4093 0.02082 0.01328 -0.0835 0.8456 1.0000
1.000 0.4373 0.02095 0.01334 -0.0833 0.8352 1.0000
1.250 0.4798 0.02071 0.01303 -0.0853 0.8292 1.0000
1.500 0.5039 0.02083 0.01311 -0.0844 0.8175 1.0000
1.750 0.5345 0.02076 0.01301 -0.0843 0.8080 1.0000
2.000 0.5712 0.02042 0.01264 -0.0850 0.8004 1.0000
2.250 0.5962 0.02043 0.01264 -0.0840 0.7887 1.0000
2.500 0.6260 0.02024 0.01245 -0.0836 0.7789 1.0000
2.750 0.6602 0.01979 0.01200 -0.0836 0.7707 1.0000
3.000 0.6856 0.01970 0.01192 -0.0826 0.7584 1.0000
3.250 0.7128 0.01951 0.01175 -0.0817 0.7468 1.0000
3.500 0.7433 0.01913 0.01140 -0.0811 0.7363 1.0000
3.750 0.7750 0.01864 0.01092 -0.0806 0.7259 1.0000
4.000 0.8016 0.01842 0.01073 -0.0796 0.7121 1.0000
4.250 0.8287 0.01817 0.01053 -0.0786 0.6979 1.0000
4.500 0.8562 0.01791 0.01028 -0.0776 0.6829 1.0000
4.750 0.8838 0.01766 0.01003 -0.0767 0.6670 1.0000
5.000 0.9072 0.01764 0.01006 -0.0753 0.6472 1.0000
5.250 0.9332 0.01752 0.00997 -0.0742 0.6277 1.0000
5.500 0.9594 0.01744 0.00986 -0.0732 0.6072 1.0000
5.750 0.9830 0.01752 0.00995 -0.0719 0.5835 1.0000
6.000 1.0064 0.01767 0.01010 -0.0706 0.5589 1.0000
6.250 1.0297 0.01788 0.01025 -0.0693 0.5334 1.0000
6.500 1.0520 0.01820 0.01049 -0.0680 0.5067 1.0000
6.750 1.0724 0.01861 0.01087 -0.0665 0.4785 1.0000
7.000 1.0922 0.01910 0.01134 -0.0649 0.4502 1.0000
7.250 1.1112 0.01965 0.01184 -0.0633 0.4219 1.0000
7.500 1.1295 0.02028 0.01242 -0.0616 0.3940 1.0000
7.750 1.1470 0.02098 0.01306 -0.0599 0.3664 1.0000
8.000 1.1636 0.02174 0.01378 -0.0581 0.3396 1.0000
8.250 1.1796 0.02258 0.01453 -0.0563 0.3138 1.0000
8.500 1.1944 0.02346 0.01535 -0.0544 0.2885 1.0000
8.750 1.2070 0.02435 0.01625 -0.0522 0.2632 1.0000
9.000 1.2170 0.02524 0.01711 -0.0498 0.2385 1.0000
9.250 1.2246 0.02616 0.01795 -0.0471 0.2152 1.0000
9.500 1.2294 0.02704 0.01890 -0.0441 0.1920 1.0000
9.750 1.2317 0.02806 0.01987 -0.0409 0.1713 1.0000
10.000 1.2347 0.02925 0.02110 -0.0380 0.1505 1.0000
10.250 1.2354 0.03078 0.02258 -0.0352 0.1295 1.0000
10.500 1.2309 0.03282 0.02451 -0.0321 0.1041 1.0000
10.750 1.2227 0.03547 0.02696 -0.0292 0.0817 1.0000
11.000 1.2197 0.03813 0.02959 -0.0266 0.0660 1.0000
11.250 1.2202 0.04075 0.03206 -0.0246 0.0585 1.0000
11.500 1.2297 0.04291 0.03445 -0.0229 0.0532 1.0000
11.750 1.2409 0.04512 0.03663 -0.0215 0.0495 1.0000
12.000 1.2614 0.04780 0.03941 -0.0205 0.0466 1.0000
12.250 1.2719 0.05024 0.04212 -0.0192 0.0443 1.0000
12.500 1.2782 0.05278 0.04487 -0.0180 0.0421 1.0000
12.750 1.2858 0.05544 0.04765 -0.0170 0.0404 1.0000
13.000 1.2965 0.05875 0.05113 -0.0162 0.0394 1.0000
13.250 1.3018 0.06301 0.05563 -0.0154 0.0387 1.0000
13.500 1.2967 0.06747 0.06038 -0.0145 0.0385 1.0000
13.750 1.2861 0.07171 0.06492 -0.0138 0.0385 1.0000
14.000 1.2719 0.07619 0.06968 -0.0135 0.0385 1.0000
14.250 1.2557 0.08071 0.07451 -0.0138 0.0386 1.0000
14.500 1.2367 0.08570 0.07977 -0.0149 0.0388 1.0000
14.750 1.2142 0.09135 0.08568 -0.0169 0.0391 1.0000
15.000 1.1862 0.09831 0.09293 -0.0204 0.0395 1.0000
15.250 1.1474 0.10817 0.10314 -0.0265 0.0404 1.0000
15.500 1.0777 0.12709 0.12243 -0.0404 0.0428 1.0000
15.750 1.0265 0.14630 0.14171 -0.0530 0.0460 1.0000
16.000 1.0112 0.15667 0.15202 -0.0585 0.0473 1.0000
16.250 1.0029 0.16532 0.16064 -0.0627 0.0479 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AIRFOIL 3024 9.84% (s3024-il)