NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 200,000 Max Cl/Cd: 56.73 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08n1-il-200000-n5.txt Download as CSV file: xf-rc08n1-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(N)1 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5881 0.09854 0.09536 0.0175 1.0000 0.0139
-8.250 -0.5862 0.09421 0.09106 0.0144 1.0000 0.0142
-8.000 -0.5847 0.08967 0.08656 0.0105 1.0000 0.0145
-7.500 -0.5772 0.07930 0.07617 -0.0015 1.0000 0.0154
-7.000 -0.4943 0.05791 0.05476 -0.0147 1.0000 0.0161
-6.750 -0.4971 0.05112 0.04792 -0.0155 1.0000 0.0166
-6.500 -0.4911 0.04618 0.04292 -0.0162 0.9874 0.0169
-6.250 -0.4770 0.04104 0.03763 -0.0177 0.9299 0.0174
-6.000 -0.4659 0.03688 0.03327 -0.0177 0.9051 0.0178
-5.750 -0.4542 0.03300 0.02918 -0.0172 0.8885 0.0184
-5.250 -0.4249 0.02570 0.02135 -0.0158 0.8646 0.0199
-5.000 -0.4077 0.02235 0.01768 -0.0148 0.8553 0.0204
-4.500 -0.3712 0.01671 0.01137 -0.0129 0.8390 0.0266
-4.000 -0.3434 0.02818 0.02174 -0.0093 0.8410 0.0311
-3.750 -0.3178 0.02439 0.01745 -0.0075 0.8341 0.0212
-3.500 -0.2921 0.02208 0.01477 -0.0065 0.8268 0.0201
-3.250 -0.2660 0.02020 0.01253 -0.0055 0.8205 0.0198
-3.000 -0.2390 0.01875 0.01082 -0.0049 0.8136 0.0205
-2.750 -0.2124 0.01791 0.00973 -0.0043 0.8074 0.0227
-2.500 -0.1855 0.01666 0.00828 -0.0038 0.8011 0.0224
-2.250 -0.1593 0.01562 0.00713 -0.0032 0.7950 0.0222
-2.000 -0.1334 0.01475 0.00618 -0.0027 0.7892 0.0222
-1.750 -0.1077 0.01402 0.00539 -0.0021 0.7829 0.0224
-1.500 -0.0824 0.01343 0.00472 -0.0016 0.7775 0.0227
-1.250 -0.0563 0.01294 0.00414 -0.0012 0.7709 0.0232
-1.000 -0.0306 0.01250 0.00359 -0.0007 0.7654 0.0250
-0.750 -0.0037 0.01221 0.00325 -0.0006 0.7588 0.0282
-0.500 0.0232 0.01203 0.00298 -0.0003 0.7529 0.0292
-0.250 0.0505 0.01190 0.00278 -0.0001 0.7466 0.0307
0.250 0.1046 0.01161 0.00251 0.0002 0.7333 0.0705
0.500 0.1189 0.00958 0.00238 0.0022 0.7264 0.6512
1.000 0.2351 0.00901 0.00262 -0.0096 0.7110 1.0000
1.250 0.2615 0.00903 0.00260 -0.0093 0.7021 1.0000
1.500 0.2876 0.00906 0.00256 -0.0090 0.6939 1.0000
1.750 0.3139 0.00909 0.00257 -0.0087 0.6846 1.0000
2.000 0.3402 0.00913 0.00260 -0.0085 0.6754 1.0000
2.250 0.3663 0.00917 0.00262 -0.0081 0.6664 1.0000
2.500 0.3924 0.00921 0.00265 -0.0078 0.6551 1.0000
2.750 0.4182 0.00924 0.00266 -0.0074 0.6385 1.0000
3.000 0.4438 0.00928 0.00269 -0.0068 0.6164 1.0000
3.250 0.4694 0.00935 0.00272 -0.0064 0.5944 1.0000
3.500 0.4950 0.00946 0.00278 -0.0059 0.5713 1.0000
3.750 0.5206 0.00959 0.00287 -0.0055 0.5430 1.0000
4.000 0.5458 0.00978 0.00301 -0.0050 0.5061 1.0000
4.250 0.5707 0.01006 0.00316 -0.0045 0.4594 1.0000
4.500 0.5947 0.01051 0.00337 -0.0040 0.3960 1.0000
4.750 0.6179 0.01117 0.00371 -0.0036 0.3196 1.0000
5.000 0.6406 0.01195 0.00416 -0.0033 0.2421 1.0000
5.250 0.6634 0.01271 0.00469 -0.0030 0.1784 1.0000
5.500 0.6862 0.01344 0.00523 -0.0026 0.1339 1.0000
5.750 0.7093 0.01411 0.00578 -0.0022 0.1040 1.0000
6.000 0.7327 0.01469 0.00633 -0.0017 0.0825 1.0000
6.250 0.7555 0.01536 0.00693 -0.0013 0.0640 1.0000
6.500 0.7784 0.01603 0.00765 -0.0007 0.0516 1.0000
6.750 0.8010 0.01673 0.00841 -0.0002 0.0421 1.0000
7.000 0.8234 0.01745 0.00921 0.0004 0.0343 1.0000
7.250 0.8439 0.01850 0.01029 0.0012 0.0287 1.0000
7.500 0.8654 0.01938 0.01134 0.0020 0.0256 1.0000
7.750 0.8869 0.02021 0.01226 0.0026 0.0217 1.0000
8.000 0.9045 0.02172 0.01389 0.0037 0.0186 1.0000
8.250 0.9244 0.02295 0.01532 0.0046 0.0170 1.0000
8.500 0.9437 0.02431 0.01687 0.0056 0.0151 1.0000
8.750 0.9640 0.02516 0.01784 0.0062 0.0128 1.0000
9.000 0.9779 0.02744 0.02030 0.0074 0.0111 1.0000
9.250 0.9935 0.02960 0.02275 0.0086 0.0104 1.0000
9.500 1.0079 0.03194 0.02543 0.0099 0.0097 1.0000
9.750 1.0193 0.03460 0.02850 0.0112 0.0091 1.0000
10.000 1.0294 0.03689 0.03111 0.0124 0.0082 1.0000
10.250 1.0399 0.03847 0.03290 0.0133 0.0074 1.0000
10.500 1.0494 0.03971 0.03425 0.0140 0.0068 1.0000
10.750 1.0491 0.04199 0.03671 0.0154 0.0064 1.0000
11.000 1.0380 0.04555 0.04055 0.0166 0.0062 1.0000
11.500 1.0097 0.05447 0.05000 0.0155 0.0060 1.0000
11.750 0.9951 0.05961 0.05537 0.0134 0.0060 1.0000
12.000 0.9788 0.06555 0.06153 0.0102 0.0060 1.0000
12.500 0.7834 0.08067 0.07728 0.0091 0.0090 1.0000
12.750 0.7553 0.09093 0.08762 0.0042 0.0095 1.0000
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Polar data table (+)
Polar graphs
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