NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Reynolds number: 500,000 Max Cl/Cd: 67.82 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08b3-il-500000.txt Download as CSV file: xf-rc08b3-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(B)3 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6041 0.10499 0.10271 0.0000 1.0000 0.0193
-9.750 -0.6038 0.10005 0.09778 -0.0030 1.0000 0.0193
-9.500 -0.6168 0.09333 0.09110 -0.0050 1.0000 0.0199
-9.250 -0.6149 0.09001 0.08779 -0.0062 1.0000 0.0202
-9.000 -0.6139 0.08624 0.08404 -0.0085 1.0000 0.0205
-8.750 -0.6170 0.08138 0.07920 -0.0131 1.0000 0.0207
-8.500 -0.6226 0.07657 0.07437 -0.0167 1.0000 0.0209
-8.250 -0.6229 0.07213 0.06989 -0.0192 1.0000 0.0213
-8.000 -0.6212 0.06757 0.06526 -0.0213 1.0000 0.0217
-7.750 -0.6178 0.06294 0.06054 -0.0228 1.0000 0.0223
-7.500 -0.6128 0.05824 0.05571 -0.0237 1.0000 0.0232
-7.250 -0.5988 0.05295 0.05009 -0.0241 1.0000 0.0257
-7.000 -0.5907 0.04918 0.04604 -0.0228 1.0000 0.0259
-6.750 -0.5749 0.02835 0.02521 -0.0214 1.0000 0.0270
-6.500 -0.5705 0.02648 0.02332 -0.0186 1.0000 0.0275
-6.250 -0.5679 0.02489 0.02167 -0.0153 1.0000 0.0280
-6.000 -0.5596 0.02310 0.01978 -0.0131 0.9994 0.0288
-5.750 -0.5349 0.02017 0.01663 -0.0140 0.9967 0.0306
-5.500 -0.4986 0.01990 0.01594 -0.0147 0.9943 0.0346
-5.250 -0.4879 0.01339 0.00902 -0.0143 0.9903 0.0367
-5.000 -0.4979 0.02176 0.01630 -0.0074 0.9934 0.0265
-4.750 -0.4697 0.01807 0.01230 -0.0075 0.9907 0.0237
-4.500 -0.4376 0.01615 0.01010 -0.0083 0.9882 0.0235
-4.250 -0.4040 0.01474 0.00852 -0.0094 0.9862 0.0236
-4.000 -0.3690 0.01380 0.00748 -0.0109 0.9844 0.0244
-3.750 -0.3340 0.01297 0.00655 -0.0125 0.9827 0.0252
-3.500 -0.3026 0.01218 0.00571 -0.0132 0.9787 0.0256
-3.250 -0.2700 0.01129 0.00476 -0.0143 0.9755 0.0265
-3.000 -0.2374 0.01050 0.00393 -0.0154 0.9724 0.0286
-2.750 -0.2039 0.01008 0.00350 -0.0167 0.9692 0.0310
-2.500 -0.1756 0.00981 0.00321 -0.0167 0.9625 0.0343
-2.250 -0.1458 0.00936 0.00278 -0.0170 0.9574 0.0463
-2.000 -0.1216 0.00875 0.00249 -0.0162 0.9498 0.1224
-1.750 -0.0984 0.00809 0.00230 -0.0155 0.9426 0.2505
-1.500 -0.0831 0.00696 0.00212 -0.0132 0.9330 0.5175
-1.250 -0.0736 0.00569 0.00214 -0.0087 0.9248 0.8492
-1.000 -0.0350 0.00593 0.00251 -0.0099 0.9190 0.9379
-0.750 0.0024 0.00623 0.00275 -0.0111 0.9130 0.9609
-0.500 0.0393 0.00638 0.00284 -0.0127 0.9047 0.9711
-0.250 0.0883 0.00655 0.00294 -0.0170 0.8990 0.9807
0.000 0.1543 0.00668 0.00304 -0.0250 0.8932 0.9951
0.250 0.1987 0.00661 0.00291 -0.0286 0.8838 1.0000
0.500 0.2235 0.00655 0.00280 -0.0280 0.8711 1.0000
0.750 0.2486 0.00650 0.00271 -0.0274 0.8581 1.0000
1.000 0.2736 0.00646 0.00262 -0.0267 0.8430 1.0000
1.250 0.2988 0.00643 0.00254 -0.0261 0.8269 1.0000
1.500 0.3245 0.00642 0.00248 -0.0256 0.8111 1.0000
1.750 0.3502 0.00641 0.00243 -0.0252 0.7914 1.0000
2.000 0.3758 0.00643 0.00238 -0.0246 0.7683 1.0000
2.250 0.4014 0.00649 0.00234 -0.0241 0.7350 1.0000
2.500 0.4269 0.00660 0.00230 -0.0236 0.6929 1.0000
2.750 0.4526 0.00677 0.00230 -0.0231 0.6448 1.0000
3.000 0.4781 0.00705 0.00234 -0.0227 0.5791 1.0000
3.250 0.5031 0.00749 0.00243 -0.0224 0.4926 1.0000
3.500 0.5278 0.00804 0.00259 -0.0222 0.3982 1.0000
3.750 0.5523 0.00860 0.00278 -0.0219 0.3117 1.0000
4.250 0.5999 0.00976 0.00328 -0.0212 0.1597 1.0000
4.500 0.6226 0.01041 0.00361 -0.0206 0.0954 1.0000
4.750 0.6454 0.01097 0.00399 -0.0198 0.0618 1.0000
5.000 0.6684 0.01142 0.00442 -0.0190 0.0490 1.0000
5.250 0.6904 0.01202 0.00501 -0.0180 0.0411 1.0000
5.500 0.7134 0.01237 0.00541 -0.0171 0.0379 1.0000
5.750 0.7356 0.01281 0.00586 -0.0161 0.0345 1.0000
6.000 0.7546 0.01375 0.00685 -0.0146 0.0313 1.0000
6.250 0.7765 0.01417 0.00735 -0.0135 0.0301 1.0000
6.500 0.7975 0.01473 0.00797 -0.0122 0.0286 1.0000
6.750 0.8184 0.01526 0.00855 -0.0110 0.0269 1.0000
7.000 0.8387 0.01584 0.00914 -0.0098 0.0253 1.0000
7.250 0.8536 0.01766 0.01107 -0.0076 0.0236 1.0000
7.500 0.8742 0.01831 0.01182 -0.0063 0.0229 1.0000
7.750 0.8946 0.01895 0.01256 -0.0050 0.0219 1.0000
8.000 0.9146 0.01958 0.01327 -0.0037 0.0207 1.0000
8.250 0.9341 0.02022 0.01399 -0.0024 0.0195 1.0000
8.500 0.9527 0.02105 0.01488 -0.0009 0.0186 1.0000
8.750 0.9669 0.02334 0.01732 0.0010 0.0176 1.0000
9.000 0.9814 0.02529 0.01954 0.0030 0.0170 1.0000
9.250 0.9980 0.02632 0.02078 0.0048 0.0163 1.0000
9.500 1.0132 0.02753 0.02219 0.0067 0.0155 1.0000
9.750 1.0281 0.02856 0.02336 0.0084 0.0147 1.0000
10.000 1.0419 0.02971 0.02463 0.0102 0.0141 1.0000
10.250 1.0541 0.03097 0.02604 0.0121 0.0137 1.0000
10.500 1.0627 0.03286 0.02808 0.0142 0.0132 1.0000
10.750 1.0574 0.03681 0.03238 0.0175 0.0129 1.0000
11.000 1.0404 0.04166 0.03766 0.0215 0.0128 1.0000
11.250 1.0212 0.04534 0.04163 0.0256 0.0127 1.0000
11.500 1.0060 0.04826 0.04479 0.0283 0.0127 1.0000
11.750 0.9869 0.05218 0.04894 0.0295 0.0126 1.0000
12.000 0.9076 0.04219 0.03905 0.0359 0.0129 1.0000
12.250 0.8862 0.04720 0.04423 0.0352 0.0130 1.0000
12.500 0.8639 0.05316 0.05035 0.0332 0.0130 1.0000
12.750 0.8388 0.06059 0.05795 0.0298 0.0131 1.0000
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Polar data table (+)
Polar graphs
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