REPUBLIC S-3 AIRFOIL (s3-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: REPUBLIC S-3 AIRFOIL (s3-il) Reynolds number: 200,000 Max Cl/Cd: 56.39 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s3-il-200000-n5.txt Download as CSV file: xf-s3-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: REPUBLIC S-3 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6094 0.06154 0.05806 -0.0295 1.0000 0.0120
-9.000 -0.6276 0.05740 0.05379 -0.0292 1.0000 0.0120
-8.750 -0.6391 0.05278 0.04897 -0.0287 1.0000 0.0120
-8.500 -0.6461 0.04839 0.04434 -0.0278 1.0000 0.0121
-8.250 -0.6483 0.04374 0.03937 -0.0265 1.0000 0.0120
-8.000 -0.6462 0.03948 0.03474 -0.0249 1.0000 0.0121
-7.750 -0.6408 0.03548 0.03035 -0.0233 1.0000 0.0122
-7.500 -0.6305 0.03221 0.02671 -0.0217 1.0000 0.0123
-7.250 -0.6165 0.02962 0.02381 -0.0202 1.0000 0.0126
-7.000 -0.5998 0.02762 0.02156 -0.0189 1.0000 0.0130
-6.750 -0.5821 0.02571 0.01937 -0.0174 1.0000 0.0134
-6.500 -0.5635 0.02414 0.01757 -0.0160 1.0000 0.0141
-6.250 -0.5445 0.02287 0.01611 -0.0146 1.0000 0.0150
-6.000 -0.5187 0.02142 0.01443 -0.0144 0.9944 0.0163
-5.750 -0.4838 0.01977 0.01252 -0.0158 0.9814 0.0172
-5.250 -0.4163 0.01711 0.00966 -0.0187 0.9520 0.0202
-5.000 -0.3824 0.01642 0.00889 -0.0200 0.9365 0.0232
-4.750 -0.3508 0.01563 0.00795 -0.0206 0.9206 0.0253
-4.500 -0.3233 0.01482 0.00714 -0.0207 0.9025 0.0289
-4.250 -0.2958 0.01441 0.00663 -0.0205 0.8844 0.0340
-4.000 -0.2710 0.01382 0.00599 -0.0198 0.8664 0.0398
-3.750 -0.2458 0.01347 0.00553 -0.0191 0.8482 0.0476
-3.500 -0.2213 0.01311 0.00509 -0.0182 0.8294 0.0565
-3.250 -0.1967 0.01277 0.00469 -0.0174 0.8110 0.0684
-3.000 -0.1720 0.01244 0.00430 -0.0167 0.7941 0.0856
-2.750 -0.1479 0.01198 0.00391 -0.0159 0.7786 0.1255
-2.500 -0.1314 0.01050 0.00349 -0.0145 0.7646 0.3942
-2.250 -0.1097 0.01001 0.00338 -0.0132 0.7515 0.5173
-2.000 -0.0856 0.00979 0.00330 -0.0121 0.7403 0.5835
-1.750 -0.0610 0.00964 0.00326 -0.0111 0.7286 0.6343
-1.500 -0.0364 0.00955 0.00323 -0.0101 0.7150 0.6796
-1.250 -0.0119 0.00949 0.00323 -0.0089 0.7000 0.7217
-1.000 0.0127 0.00948 0.00322 -0.0077 0.6833 0.7567
-0.750 0.0382 0.00947 0.00316 -0.0068 0.6654 0.7795
-0.500 0.0640 0.00946 0.00310 -0.0060 0.6430 0.7985
0.000 0.1157 0.00949 0.00294 -0.0046 0.5886 0.8300
0.250 0.1417 0.00955 0.00287 -0.0039 0.5554 0.8439
0.500 0.1673 0.00966 0.00281 -0.0032 0.5169 0.8581
0.750 0.1933 0.00982 0.00278 -0.0026 0.4772 0.8722
1.000 0.2201 0.01003 0.00279 -0.0023 0.4435 0.8861
1.250 0.2483 0.01023 0.00284 -0.0023 0.4185 0.9003
1.500 0.2783 0.01043 0.00292 -0.0027 0.3973 0.9145
1.750 0.3101 0.01062 0.00301 -0.0035 0.3781 0.9280
2.000 0.3436 0.01079 0.00310 -0.0047 0.3588 0.9403
2.250 0.3779 0.01097 0.00319 -0.0061 0.3393 0.9522
2.500 0.4126 0.01117 0.00330 -0.0077 0.3206 0.9638
2.750 0.4475 0.01138 0.00341 -0.0094 0.3040 0.9748
3.000 0.4827 0.01161 0.00355 -0.0111 0.2892 0.9851
3.250 0.5186 0.01184 0.00372 -0.0131 0.2765 0.9949
3.500 0.5473 0.01206 0.00389 -0.0135 0.2664 1.0000
3.750 0.5693 0.01232 0.00409 -0.0126 0.2582 1.0000
4.000 0.5924 0.01253 0.00429 -0.0118 0.2500 1.0000
4.250 0.6153 0.01282 0.00452 -0.0110 0.2425 1.0000
4.500 0.6390 0.01304 0.00474 -0.0104 0.2336 1.0000
4.750 0.6627 0.01331 0.00499 -0.0097 0.2261 1.0000
5.000 0.6869 0.01355 0.00524 -0.0092 0.2188 1.0000
5.250 0.7109 0.01385 0.00550 -0.0086 0.2127 1.0000
5.500 0.7356 0.01409 0.00580 -0.0081 0.2070 1.0000
5.750 0.7599 0.01438 0.00609 -0.0076 0.2015 1.0000
6.000 0.7843 0.01468 0.00641 -0.0071 0.1971 1.0000
6.250 0.8090 0.01496 0.00674 -0.0067 0.1924 1.0000
6.500 0.8332 0.01527 0.00708 -0.0062 0.1866 1.0000
6.750 0.8578 0.01556 0.00741 -0.0058 0.1785 1.0000
7.000 0.8815 0.01590 0.00773 -0.0053 0.1683 1.0000
7.250 0.9056 0.01622 0.00808 -0.0049 0.1581 1.0000
7.500 0.9299 0.01654 0.00844 -0.0045 0.1478 1.0000
7.750 0.9536 0.01691 0.00882 -0.0041 0.1301 1.0000
8.000 0.9748 0.01756 0.00926 -0.0035 0.0920 1.0000
8.250 0.9919 0.01874 0.01017 -0.0025 0.0642 1.0000
8.500 1.0116 0.01959 0.01097 -0.0017 0.0509 1.0000
8.750 1.0281 0.02080 0.01204 -0.0006 0.0208 1.0000
9.000 1.0455 0.02185 0.01313 0.0006 0.0164 1.0000
9.250 1.0645 0.02266 0.01412 0.0016 0.0149 1.0000
9.500 1.0824 0.02352 0.01516 0.0026 0.0141 1.0000
9.750 1.0983 0.02453 0.01635 0.0039 0.0132 1.0000
10.000 1.1111 0.02573 0.01776 0.0053 0.0122 1.0000
10.250 1.1201 0.02715 0.01938 0.0071 0.0114 1.0000
10.500 1.1227 0.02882 0.02126 0.0095 0.0108 1.0000
10.750 1.1222 0.03044 0.02305 0.0121 0.0105 1.0000
11.000 1.1244 0.03201 0.02477 0.0139 0.0103 1.0000
11.250 1.1240 0.03394 0.02686 0.0152 0.0101 1.0000
11.500 1.1221 0.03623 0.02930 0.0159 0.0101 1.0000
11.750 1.1178 0.03906 0.03229 0.0159 0.0099 1.0000
12.000 1.1118 0.04239 0.03577 0.0153 0.0097 1.0000
12.250 1.1042 0.04621 0.03975 0.0141 0.0096 1.0000
12.500 1.0957 0.05054 0.04423 0.0123 0.0096 1.0000
12.750 1.0857 0.05547 0.04930 0.0099 0.0095 1.0000
13.000 1.0736 0.06105 0.05507 0.0070 0.0094 1.0000
13.250 1.0630 0.06658 0.06073 0.0041 0.0094 1.0000
13.500 1.0514 0.07230 0.06658 0.0011 0.0092 1.0000
13.750 1.0401 0.07793 0.07231 -0.0016 0.0093 1.0000
14.000 1.0297 0.08347 0.07796 -0.0043 0.0092 1.0000
14.250 1.0204 0.08893 0.08351 -0.0069 0.0091 1.0000
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Polar data table (+)
Polar graphs
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