ONERA NACA CAMBRE AIRFOIL (ncambre-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: ONERA NACA CAMBRE AIRFOIL (ncambre-il) Reynolds number: 100,000 Max Cl/Cd: 39.61 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ncambre-il-100000.txt Download as CSV file: xf-ncambre-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: ONERA NACA CAMBRE 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5441 0.10922 0.10405 -0.0037 1.0000 0.1181
-9.750 -0.5693 0.10405 0.09897 -0.0116 1.0000 0.1234
-9.500 -0.6182 0.09709 0.09208 -0.0231 1.0000 0.1242
-9.250 -0.5502 0.09654 0.09151 -0.0095 1.0000 0.1302
-9.000 -0.5559 0.09229 0.08731 -0.0125 1.0000 0.1361
-8.750 -0.6101 0.08519 0.08024 -0.0229 1.0000 0.1390
-8.500 -0.5823 0.08152 0.07666 -0.0201 1.0000 0.1440
-8.250 -0.5834 0.07829 0.07345 -0.0201 1.0000 0.1500
-7.750 -0.5976 0.07029 0.06542 -0.0213 1.0000 0.1632
-7.500 -0.6198 0.06602 0.06099 -0.0221 1.0000 0.1736
-7.000 -0.5997 0.06060 0.05565 -0.0190 1.0000 0.1959
-6.750 -0.5962 0.05804 0.05308 -0.0172 1.0000 0.2123
-5.250 -0.5357 0.03383 0.02578 -0.0073 1.0000 0.0999
-5.000 -0.5147 0.03144 0.02294 -0.0051 1.0000 0.0894
-4.750 -0.4943 0.02949 0.02032 -0.0025 1.0000 0.0825
-4.500 -0.4734 0.02819 0.01876 -0.0009 1.0000 0.0818
-4.250 -0.4523 0.02631 0.01677 0.0003 1.0000 0.0829
-4.000 -0.4302 0.02490 0.01525 0.0015 1.0000 0.0833
-3.750 -0.4073 0.02364 0.01390 0.0027 1.0000 0.0833
-3.500 -0.3844 0.02247 0.01270 0.0038 1.0000 0.0842
-3.250 -0.3622 0.02149 0.01174 0.0048 1.0000 0.0860
-3.000 -0.3408 0.02074 0.01103 0.0059 1.0000 0.0901
-2.750 -0.3194 0.02015 0.01041 0.0071 1.0000 0.0940
-2.500 -0.2991 0.01924 0.00964 0.0082 1.0000 0.0978
-2.250 -0.2795 0.01867 0.00915 0.0092 1.0000 0.1040
-2.000 -0.2598 0.01817 0.00870 0.0102 1.0000 0.1145
-1.750 -0.2393 0.01769 0.00830 0.0109 1.0000 0.1335
-1.500 -0.2239 0.01459 0.00844 0.0145 0.9984 0.7774
-1.250 -0.1487 0.01584 0.00983 0.0108 0.9993 0.9716
-1.000 -0.0207 0.01635 0.00998 -0.0074 1.0000 1.0000
-0.750 0.0260 0.01632 0.00981 -0.0129 0.9891 1.0000
-0.500 0.0937 0.01615 0.00952 -0.0214 0.9744 1.0000
-0.250 0.1596 0.01588 0.00916 -0.0293 0.9598 1.0000
0.000 0.2197 0.01549 0.00872 -0.0357 0.9427 1.0000
0.250 0.2665 0.01508 0.00826 -0.0393 0.9185 1.0000
0.500 0.3048 0.01466 0.00778 -0.0408 0.8933 1.0000
0.750 0.3303 0.01439 0.00743 -0.0397 0.8632 1.0000
1.000 0.3523 0.01420 0.00715 -0.0380 0.8334 1.0000
1.250 0.3736 0.01408 0.00693 -0.0362 0.8042 1.0000
1.500 0.3947 0.01402 0.00676 -0.0343 0.7751 1.0000
1.750 0.4157 0.01401 0.00661 -0.0325 0.7466 1.0000
2.000 0.4369 0.01406 0.00652 -0.0308 0.7171 1.0000
2.250 0.4581 0.01416 0.00648 -0.0291 0.6868 1.0000
2.500 0.4795 0.01430 0.00648 -0.0276 0.6564 1.0000
2.750 0.5009 0.01448 0.00650 -0.0260 0.6260 1.0000
3.000 0.5223 0.01469 0.00656 -0.0246 0.5965 1.0000
3.250 0.5440 0.01494 0.00664 -0.0231 0.5681 1.0000
3.500 0.5654 0.01523 0.00679 -0.0218 0.5389 1.0000
3.750 0.5868 0.01555 0.00698 -0.0204 0.5103 1.0000
4.000 0.6081 0.01591 0.00720 -0.0191 0.4832 1.0000
4.250 0.6295 0.01631 0.00742 -0.0178 0.4574 1.0000
4.500 0.6506 0.01670 0.00774 -0.0165 0.4311 1.0000
4.750 0.6719 0.01712 0.00800 -0.0152 0.4084 1.0000
5.000 0.6931 0.01755 0.00838 -0.0140 0.3851 1.0000
5.250 0.7145 0.01804 0.00875 -0.0128 0.3645 1.0000
5.500 0.7361 0.01860 0.00917 -0.0117 0.3456 1.0000
5.750 0.7574 0.01917 0.00970 -0.0105 0.3269 1.0000
6.000 0.7788 0.01977 0.01027 -0.0094 0.3096 1.0000
6.250 0.8005 0.02040 0.01086 -0.0084 0.2938 1.0000
6.500 0.8224 0.02109 0.01149 -0.0074 0.2791 1.0000
6.750 0.8447 0.02183 0.01214 -0.0065 0.2657 1.0000
7.000 0.8666 0.02256 0.01291 -0.0055 0.2530 1.0000
7.250 0.8885 0.02341 0.01386 -0.0045 0.2416 1.0000
7.500 0.9111 0.02431 0.01476 -0.0038 0.2318 1.0000
7.750 0.9337 0.02512 0.01557 -0.0030 0.2223 1.0000
8.000 0.9552 0.02617 0.01678 -0.0021 0.2139 1.0000
8.250 0.9788 0.02714 0.01774 -0.0015 0.2067 1.0000
8.500 0.9991 0.02834 0.01917 -0.0005 0.1994 1.0000
8.750 1.0215 0.02928 0.02012 0.0002 0.1923 1.0000
9.000 1.0407 0.03051 0.02157 0.0013 0.1850 1.0000
9.250 1.0624 0.03120 0.02224 0.0021 0.1768 1.0000
9.500 1.0789 0.03228 0.02359 0.0035 0.1686 1.0000
9.750 1.1021 0.03297 0.02408 0.0038 0.1604 1.0000
10.000 1.1149 0.03381 0.02531 0.0058 0.1518 1.0000
10.250 1.1330 0.03465 0.02618 0.0068 0.1435 1.0000
10.500 1.1500 0.03523 0.02679 0.0081 0.1345 1.0000
10.750 1.1604 0.03656 0.02841 0.0100 0.1256 1.0000
11.000 1.1754 0.03758 0.02939 0.0113 0.1166 1.0000
11.250 1.1874 0.03852 0.03039 0.0130 0.1074 1.0000
11.500 1.1903 0.04046 0.03265 0.0156 0.0995 1.0000
11.750 1.2036 0.04181 0.03388 0.0169 0.0923 1.0000
12.000 1.2006 0.04409 0.03656 0.0198 0.0871 1.0000
12.250 1.2191 0.04575 0.03796 0.0203 0.0816 1.0000
12.500 1.2044 0.04842 0.04113 0.0240 0.0793 1.0000
12.750 1.1915 0.05117 0.04420 0.0269 0.0770 1.0000
13.000 1.1835 0.05375 0.04697 0.0287 0.0746 1.0000
13.250 1.2042 0.05526 0.04824 0.0290 0.0707 1.0000
13.500 0.8341 0.12185 0.11634 -0.0085 0.1029 1.0000
13.750 0.8047 0.13381 0.12822 -0.0145 0.1028 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA NACA CAMBRE AIRFOIL (ncambre-il)