USA 35 AIRFOIL (usa35-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 35 AIRFOIL (usa35-il) Reynolds number: 100,000 Max Cl/Cd: 46.36 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa35-il-100000-n5.txt Download as CSV file: xf-usa35-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 35 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
------ -------- --------- --------- -------- -------- --------
-10.500 0.0103 0.11465 0.10916 -0.0864 0.9087 0.0808
-10.250 0.0114 0.11144 0.10592 -0.0895 0.8988 0.0820
-10.000 0.0087 0.10825 0.10273 -0.0923 0.8866 0.0823
-9.250 0.0505 0.10017 0.09459 -0.0924 0.8549 0.0858
-8.750 0.0374 0.08860 0.08285 -0.0989 0.8349 0.0627
-8.500 0.0459 0.08599 0.08023 -0.0993 0.8238 0.0620
-8.250 0.0511 0.08286 0.07706 -0.1004 0.8140 0.0613
-8.000 0.0532 0.07944 0.07363 -0.1019 0.8029 0.0606
-7.750 0.0538 0.07573 0.06987 -0.1039 0.7933 0.0604
-7.500 0.0519 0.07197 0.06610 -0.1061 0.7818 0.0607
-7.250 0.0458 0.06777 0.06185 -0.1090 0.7722 0.0608
-7.000 0.0402 0.06212 0.05614 -0.1147 0.7609 0.0607
-6.750 0.0320 0.05382 0.04756 -0.1221 0.7530 0.0602
-6.500 0.0225 0.04616 0.03949 -0.1264 0.7422 0.0601
-6.250 0.0278 0.04109 0.03389 -0.1281 0.7338 0.0605
-6.000 0.0383 0.03736 0.02964 -0.1286 0.7243 0.0615
-5.750 0.0529 0.03405 0.02564 -0.1287 0.7154 0.0632
-5.500 0.0775 0.03325 0.02479 -0.1285 0.7075 0.0642
-5.250 0.0993 0.03207 0.02343 -0.1281 0.6980 0.0655
-5.000 0.1234 0.03048 0.02145 -0.1280 0.6909 0.0668
-4.750 0.1461 0.02899 0.01956 -0.1276 0.6822 0.0682
-4.500 0.1710 0.02799 0.01833 -0.1273 0.6741 0.0697
-4.250 0.1980 0.02736 0.01759 -0.1272 0.6674 0.0716
-4.000 0.2225 0.02663 0.01666 -0.1268 0.6593 0.0739
-3.750 0.2492 0.02582 0.01560 -0.1265 0.6528 0.0759
-3.500 0.2767 0.02525 0.01494 -0.1264 0.6469 0.0777
-3.250 0.3018 0.02479 0.01439 -0.1260 0.6394 0.0804
-3.000 0.3288 0.02428 0.01372 -0.1257 0.6331 0.0833
-2.500 0.3822 0.02351 0.01280 -0.1252 0.6211 0.0892
-2.250 0.4083 0.02315 0.01243 -0.1248 0.6144 0.0925
-2.000 0.4366 0.02279 0.01196 -0.1247 0.6086 0.0964
-1.750 0.4610 0.02254 0.01172 -0.1241 0.6011 0.1002
-1.500 0.4868 0.02230 0.01147 -0.1236 0.5943 0.1048
-1.250 0.5148 0.02203 0.01112 -0.1235 0.5890 0.1098
-1.000 0.5401 0.02190 0.01101 -0.1230 0.5831 0.1159
-0.750 0.5653 0.02177 0.01092 -0.1225 0.5772 0.1226
-0.500 0.5922 0.02163 0.01076 -0.1223 0.5721 0.1316
-0.250 0.6208 0.02148 0.01058 -0.1223 0.5677 0.1447
0.000 0.6451 0.02143 0.01068 -0.1217 0.5622 0.1642
0.250 0.6703 0.02134 0.01074 -0.1213 0.5567 0.2019
0.500 0.6967 0.02117 0.01076 -0.1211 0.5518 0.2663
0.750 0.7245 0.02091 0.01077 -0.1212 0.5474 0.3701
1.000 0.7468 0.02065 0.01104 -0.1203 0.5422 0.5214
1.250 0.7626 0.02023 0.01135 -0.1173 0.5369 0.7426
1.500 0.8033 0.02009 0.01136 -0.1191 0.5315 1.0000
1.750 0.8311 0.02030 0.01137 -0.1190 0.5272 1.0000
2.000 0.8554 0.02061 0.01156 -0.1184 0.5223 1.0000
2.250 0.8773 0.02095 0.01184 -0.1175 0.5164 1.0000
2.500 0.9020 0.02122 0.01199 -0.1169 0.5111 1.0000
2.750 0.9293 0.02145 0.01205 -0.1168 0.5064 1.0000
3.000 0.9529 0.02179 0.01231 -0.1161 0.5014 1.0000
3.250 0.9734 0.02219 0.01269 -0.1150 0.4956 1.0000
3.500 0.9972 0.02250 0.01293 -0.1144 0.4904 1.0000
3.750 1.0238 0.02276 0.01306 -0.1141 0.4859 1.0000
4.000 1.0460 0.02314 0.01339 -0.1133 0.4807 1.0000
4.250 1.0643 0.02358 0.01385 -0.1119 0.4744 1.0000
4.500 1.0871 0.02390 0.01411 -0.1111 0.4690 1.0000
4.750 1.1138 0.02414 0.01421 -0.1109 0.4646 1.0000
5.000 1.1299 0.02468 0.01480 -0.1092 0.4588 1.0000
5.250 1.1480 0.02515 0.01529 -0.1078 0.4531 1.0000
5.500 1.1706 0.02548 0.01556 -0.1071 0.4483 1.0000
5.750 1.1952 0.02578 0.01576 -0.1066 0.4441 1.0000
6.000 1.2054 0.02647 0.01655 -0.1041 0.4376 1.0000
6.250 1.2221 0.02693 0.01700 -0.1025 0.4323 1.0000
6.500 1.2445 0.02722 0.01721 -0.1017 0.4280 1.0000
6.750 1.2563 0.02787 0.01791 -0.0995 0.4228 1.0000
7.000 1.2664 0.02860 0.01868 -0.0971 0.4170 1.0000
7.250 1.2841 0.02901 0.01905 -0.0958 0.4119 1.0000
7.500 1.2993 0.02956 0.01957 -0.0942 0.4066 1.0000
7.750 1.3044 0.03055 0.02063 -0.0915 0.4000 1.0000
8.000 1.3201 0.03105 0.02109 -0.0901 0.3945 1.0000
8.250 1.3320 0.03181 0.02185 -0.0883 0.3889 1.0000
8.500 1.3368 0.03299 0.02309 -0.0860 0.3825 1.0000
8.750 1.3521 0.03362 0.02368 -0.0847 0.3774 1.0000
9.000 1.3658 0.03445 0.02450 -0.0834 0.3728 1.0000
9.250 1.3687 0.03599 0.02615 -0.0813 0.3673 1.0000
9.500 1.3797 0.03705 0.02721 -0.0800 0.3627 1.0000
9.750 1.4003 0.03748 0.02757 -0.0793 0.3589 1.0000
10.000 1.4017 0.03928 0.02948 -0.0774 0.3539 1.0000
10.250 1.4039 0.04109 0.03137 -0.0758 0.3487 1.0000
10.500 1.4166 0.04211 0.03238 -0.0747 0.3443 1.0000
10.750 1.4397 0.04238 0.03256 -0.0743 0.3407 1.0000
11.000 1.4268 0.04555 0.03594 -0.0721 0.3352 1.0000
11.250 1.4286 0.04759 0.03805 -0.0708 0.3304 1.0000
11.500 1.4442 0.04843 0.03887 -0.0701 0.3264 1.0000
11.750 1.4570 0.04957 0.04000 -0.0693 0.3223 1.0000
12.000 1.4364 0.05385 0.04448 -0.0676 0.3163 1.0000
12.250 1.4412 0.05578 0.04646 -0.0667 0.3115 1.0000
12.500 1.4651 0.05578 0.04638 -0.0662 0.3078 1.0000
12.750 1.4387 0.06111 0.05193 -0.0651 0.3016 1.0000
13.000 1.4289 0.06483 0.05575 -0.0644 0.2957 1.0000
13.250 1.4488 0.06519 0.05607 -0.0639 0.2920 1.0000
13.500 1.4274 0.07048 0.06152 -0.0635 0.2859 1.0000
13.750 1.4036 0.07629 0.06749 -0.0634 0.2789 1.0000
14.000 1.4264 0.07622 0.06739 -0.0629 0.2756 1.0000
14.250 1.3593 0.08819 0.07963 -0.0640 0.2650 1.0000
14.500 1.3751 0.08900 0.08045 -0.0636 0.2612 1.0000
14.750 1.4057 0.08773 0.07912 -0.0630 0.2589 1.0000
15.250 1.3558 0.10100 0.09263 -0.0647 0.2446 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 35 AIRFOIL (usa35-il)