NREL's S804 Airfoil (s804-nr) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NREL's S804 Airfoil (s804-nr) Reynolds number: 200,000 Max Cl/Cd: 58.23 at α=11.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s804-nr-200000.txt Download as CSV file: xf-s804-nr-200000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S804 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5127 0.06327 0.05946 -0.0742 1.0000 0.0464
-10.250 -0.5363 0.06117 0.05741 -0.0718 1.0000 0.0463
-10.000 -0.5344 0.05464 0.05068 -0.0817 0.9893 0.0460
-9.750 -0.5223 0.04801 0.04373 -0.0919 0.9745 0.0453
-9.500 -0.5099 0.04172 0.03695 -0.0992 0.9594 0.0445
-9.250 -0.4902 0.03685 0.03156 -0.1038 0.9459 0.0443
-9.000 -0.4640 0.03324 0.02753 -0.1070 0.9334 0.0446
-8.750 -0.4318 0.03021 0.02406 -0.1101 0.9225 0.0454
-8.500 -0.3902 0.02763 0.02096 -0.1145 0.9135 0.0468
-8.250 -0.3493 0.02605 0.01947 -0.1174 0.9002 0.0487
-8.000 -0.3063 0.02444 0.01766 -0.1208 0.8841 0.0506
-7.750 -0.2637 0.02297 0.01580 -0.1240 0.8644 0.0523
-7.500 -0.2246 0.02159 0.01445 -0.1263 0.8435 0.0544
-7.250 -0.1910 0.02082 0.01348 -0.1275 0.8217 0.0571
-7.000 -0.1645 0.02004 0.01257 -0.1273 0.7998 0.0596
-6.750 -0.1395 0.01943 0.01192 -0.1268 0.7793 0.0621
-6.500 -0.1145 0.01892 0.01122 -0.1263 0.7607 0.0650
-6.250 -0.0913 0.01827 0.01057 -0.1257 0.7430 0.0685
-6.000 -0.0668 0.01783 0.01002 -0.1252 0.7263 0.0730
-5.750 -0.0428 0.01720 0.00940 -0.1249 0.7111 0.0779
-5.500 -0.0174 0.01665 0.00878 -0.1248 0.6971 0.0854
-5.250 0.0081 0.01606 0.00819 -0.1248 0.6826 0.0981
-5.000 0.0346 0.01524 0.00749 -0.1253 0.6689 0.1287
-4.750 0.0625 0.01430 0.00684 -0.1265 0.6567 0.2040
-4.500 0.0904 0.01356 0.00646 -0.1274 0.6436 0.2965
-4.250 0.1187 0.01336 0.00650 -0.1277 0.6323 0.3801
-4.000 0.1470 0.01350 0.00661 -0.1276 0.6209 0.4255
-3.750 0.1754 0.01373 0.00673 -0.1274 0.6102 0.4538
-3.500 0.2032 0.01401 0.00692 -0.1270 0.5999 0.4761
-3.250 0.2306 0.01437 0.00722 -0.1264 0.5900 0.4968
-3.000 0.2582 0.01470 0.00742 -0.1260 0.5803 0.5144
-2.750 0.2846 0.01498 0.00768 -0.1251 0.5714 0.5253
-2.500 0.3125 0.01515 0.00774 -0.1249 0.5623 0.5372
-2.250 0.3394 0.01546 0.00796 -0.1242 0.5548 0.5470
-2.000 0.3674 0.01558 0.00802 -0.1241 0.5465 0.5578
-1.750 0.3937 0.01578 0.00818 -0.1233 0.5394 0.5654
-1.500 0.4226 0.01599 0.00825 -0.1234 0.5327 0.5752
-1.250 0.4493 0.01606 0.00833 -0.1229 0.5257 0.5820
-1.000 0.4766 0.01623 0.00843 -0.1225 0.5196 0.5892
-0.750 0.5064 0.01642 0.00846 -0.1230 0.5138 0.5976
-0.500 0.5327 0.01646 0.00854 -0.1225 0.5075 0.6029
-0.250 0.5601 0.01657 0.00860 -0.1222 0.5018 0.6088
0.000 0.5899 0.01677 0.00863 -0.1226 0.4968 0.6155
0.250 0.6182 0.01684 0.00869 -0.1227 0.4918 0.6211
0.500 0.6451 0.01692 0.00880 -0.1224 0.4869 0.6255
0.750 0.6733 0.01704 0.00887 -0.1224 0.4824 0.6308
1.000 0.7033 0.01722 0.00892 -0.1229 0.4785 0.6366
1.250 0.7325 0.01741 0.00906 -0.1233 0.4746 0.6414
1.500 0.7589 0.01751 0.00923 -0.1230 0.4705 0.6457
1.750 0.7866 0.01765 0.00937 -0.1230 0.4663 0.6507
2.000 0.8158 0.01779 0.00944 -0.1233 0.4626 0.6561
2.250 0.8458 0.01800 0.00955 -0.1238 0.4593 0.6614
2.500 0.8732 0.01828 0.00985 -0.1237 0.4562 0.6661
2.750 0.8997 0.01847 0.01012 -0.1235 0.4530 0.6717
3.000 0.9278 0.01868 0.01035 -0.1237 0.4497 0.6778
3.250 0.9558 0.01888 0.01056 -0.1239 0.4467 0.6831
3.500 0.9832 0.01909 0.01079 -0.1238 0.4439 0.6884
3.750 1.0120 0.01935 0.01101 -0.1241 0.4413 0.6954
4.000 1.0419 0.01980 0.01138 -0.1246 0.4385 0.7024
4.250 1.0663 0.02002 0.01175 -0.1241 0.4359 0.7085
4.500 1.0923 0.02032 0.01214 -0.1239 0.4332 0.7160
4.750 1.1187 0.02062 0.01252 -0.1238 0.4306 0.7232
5.000 1.1446 0.02092 0.01290 -0.1235 0.4281 0.7307
5.250 1.1723 0.02125 0.01323 -0.1237 0.4258 0.7403
5.500 1.1988 0.02152 0.01355 -0.1235 0.4235 0.7487
5.750 1.2280 0.02192 0.01391 -0.1239 0.4211 0.7594
6.000 1.2531 0.02244 0.01451 -0.1236 0.4189 0.7685
6.250 1.2764 0.02284 0.01508 -0.1230 0.4168 0.7799
6.500 1.2987 0.02325 0.01563 -0.1222 0.4143 0.7917
6.750 1.3212 0.02361 0.01613 -0.1214 0.4116 0.8047
7.000 1.3442 0.02395 0.01659 -0.1207 0.4090 0.8196
7.250 1.3675 0.02426 0.01698 -0.1199 0.4067 0.8359
7.500 1.3916 0.02454 0.01729 -0.1193 0.4045 0.8544
7.750 1.4177 0.02494 0.01769 -0.1190 0.4023 0.8786
8.000 1.4357 0.02540 0.01830 -0.1175 0.4000 0.9112
8.250 1.4535 0.02575 0.01887 -0.1161 0.3971 1.0000
8.500 1.4788 0.02635 0.01955 -0.1164 0.3939 1.0000
8.750 1.5047 0.02679 0.02004 -0.1166 0.3906 1.0000
9.000 1.5327 0.02708 0.02031 -0.1170 0.3877 1.0000
9.250 1.5647 0.02732 0.02047 -0.1179 0.3848 1.0000
9.500 1.5889 0.02798 0.02117 -0.1177 0.3816 1.0000
9.750 1.6037 0.02859 0.02196 -0.1161 0.3778 1.0000
10.000 1.6229 0.02902 0.02249 -0.1151 0.3739 1.0000
10.250 1.6480 0.02915 0.02262 -0.1148 0.3703 1.0000
10.500 1.6814 0.02910 0.02245 -0.1157 0.3668 1.0000
10.750 1.7000 0.02969 0.02314 -0.1146 0.3630 1.0000
11.000 1.7090 0.03031 0.02395 -0.1121 0.3588 1.0000
11.250 1.7263 0.03058 0.02430 -0.1107 0.3546 1.0000
11.500 1.7551 0.03045 0.02412 -0.1109 0.3508 1.0000
11.750 1.7842 0.03064 0.02425 -0.1112 0.3469 1.0000
12.000 1.7807 0.03147 0.02533 -0.1070 0.3426 1.0000
12.250 1.7876 0.03188 0.02585 -0.1040 0.3382 1.0000
12.500 1.8106 0.03173 0.02567 -0.1033 0.3340 1.0000
12.750 1.8320 0.03188 0.02579 -0.1024 0.3297 1.0000
13.000 1.8151 0.03302 0.02717 -0.0966 0.3255 1.0000
13.250 1.8158 0.03368 0.02793 -0.0934 0.3207 1.0000
13.500 1.8362 0.03359 0.02779 -0.0924 0.3160 1.0000
13.750 1.8315 0.03479 0.02912 -0.0891 0.3114 1.0000
14.000 1.8170 0.03659 0.03111 -0.0854 0.3062 1.0000
14.250 1.8232 0.03736 0.03191 -0.0837 0.3009 1.0000
14.500 1.8208 0.03892 0.03354 -0.0817 0.2955 1.0000
14.750 1.8011 0.04193 0.03676 -0.0793 0.2893 1.0000
15.000 1.8105 0.04283 0.03761 -0.0784 0.2831 1.0000
15.250 1.7803 0.04753 0.04256 -0.0769 0.2759 1.0000
15.500 1.7780 0.04984 0.04488 -0.0763 0.2686 1.0000
15.750 1.7483 0.05548 0.05071 -0.0762 0.2600 1.0000
16.000 1.7417 0.05875 0.05397 -0.0762 0.2513 1.0000
16.250 1.7171 0.06454 0.05986 -0.0770 0.2412 1.0000
16.500 1.6950 0.07036 0.06574 -0.0780 0.2308 1.0000
16.750 1.6806 0.07530 0.07066 -0.0789 0.2205 1.0000
17.000 1.6713 0.07958 0.07484 -0.0797 0.2100 1.0000
17.250 1.6504 0.08582 0.08112 -0.0814 0.1995 1.0000
17.500 1.6356 0.09127 0.08658 -0.0828 0.1897 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NREL's S804 Airfoil (s804-nr)