NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 1,000,000 Max Cl/Cd: 109.06 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m16-il-1000000.txt Download as CSV file: xf-m16-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M16 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5169 0.10181 0.10047 0.0242 1.0000 0.0098
-7.750 -0.5178 0.09886 0.09734 0.0233 0.8886 0.0098
-7.500 -0.5144 0.09526 0.09360 0.0204 0.8511 0.0099
-7.250 -0.5051 0.09102 0.08926 0.0164 0.8259 0.0099
-7.000 -0.4933 0.08662 0.08476 0.0122 0.8050 0.0099
-6.750 -0.4876 0.08137 0.07944 0.0099 0.7880 0.0103
-6.500 -0.4731 0.07836 0.07634 0.0075 0.7712 0.0106
-6.250 -0.4564 0.07510 0.07300 0.0046 0.7559 0.0110
-6.000 -0.4380 0.07159 0.06941 0.0015 0.7422 0.0116
-5.750 -0.4092 0.06751 0.06519 -0.0038 0.7300 0.0137
-5.500 -0.3841 0.06323 0.06080 -0.0076 0.7187 0.0139
-5.250 -0.3610 0.05880 0.05625 -0.0104 0.7086 0.0139
-5.000 -0.3370 0.05405 0.05136 -0.0130 0.6997 0.0139
-4.750 -0.3120 0.04902 0.04617 -0.0153 0.6910 0.0140
-4.500 -0.2912 0.04350 0.04047 -0.0173 0.6835 0.0133
-4.250 -0.2615 0.03782 0.03454 -0.0190 0.6763 0.0138
-4.000 -0.2411 0.03958 0.03635 -0.0192 0.6672 0.0163
-3.250 -0.1508 0.02242 0.01806 -0.0208 0.6497 0.0202
-3.000 -0.1259 0.02451 0.02027 -0.0214 0.6422 0.0214
-2.750 -0.0970 0.02678 0.02262 -0.0214 0.6348 0.0267
-2.500 -0.0679 0.02436 0.01996 -0.0212 0.6292 0.0271
-2.250 -0.0397 0.02164 0.01696 -0.0212 0.6235 0.0272
-2.000 -0.0124 0.01579 0.01042 -0.0211 0.6189 0.0290
-1.750 0.0152 0.01511 0.00967 -0.0213 0.6130 0.0303
-1.500 0.0432 0.01455 0.00901 -0.0215 0.6069 0.0317
-1.250 0.0715 0.01388 0.00820 -0.0216 0.6012 0.0334
-1.000 0.1001 0.01308 0.00727 -0.0216 0.5954 0.0350
-0.750 0.1290 0.01159 0.00553 -0.0215 0.5898 0.0334
-0.500 0.1576 0.01080 0.00461 -0.0214 0.5842 0.0339
-0.250 0.1860 0.01026 0.00399 -0.0215 0.5781 0.0342
0.000 0.2144 0.00976 0.00342 -0.0215 0.5724 0.0339
0.250 0.2426 0.00933 0.00295 -0.0215 0.5660 0.0336
0.500 0.2708 0.00901 0.00258 -0.0216 0.5600 0.0336
0.750 0.2991 0.00873 0.00228 -0.0216 0.5535 0.0339
1.000 0.3274 0.00848 0.00198 -0.0217 0.5469 0.0335
1.250 0.3558 0.00827 0.00176 -0.0218 0.5404 0.0336
1.500 0.3843 0.00814 0.00159 -0.0220 0.5333 0.0344
1.750 0.4128 0.00804 0.00148 -0.0221 0.5262 0.0362
2.250 0.4698 0.00796 0.00138 -0.0225 0.5113 0.0392
2.500 0.4984 0.00794 0.00133 -0.0227 0.5036 0.0422
2.750 0.5269 0.00791 0.00133 -0.0229 0.4956 0.0502
3.000 0.5431 0.00590 0.00151 -0.0208 0.4888 0.9536
3.250 0.5818 0.00589 0.00151 -0.0233 0.4793 1.0000
3.500 0.6095 0.00599 0.00157 -0.0233 0.4652 1.0000
3.750 0.6373 0.00611 0.00163 -0.0234 0.4504 1.0000
4.000 0.6652 0.00625 0.00171 -0.0236 0.4343 1.0000
4.250 0.6931 0.00640 0.00180 -0.0237 0.4152 1.0000
4.500 0.7209 0.00661 0.00194 -0.0239 0.3922 1.0000
4.750 0.7485 0.00694 0.00211 -0.0242 0.3498 1.0000
5.000 0.7715 0.00975 0.00340 -0.0255 0.0365 1.0000
5.250 0.7987 0.01022 0.00381 -0.0256 0.0189 1.0000
5.500 0.8259 0.01072 0.00441 -0.0257 0.0150 1.0000
5.750 0.8532 0.01105 0.00478 -0.0258 0.0144 1.0000
6.000 0.8801 0.01145 0.00522 -0.0259 0.0131 1.0000
6.250 0.9069 0.01188 0.00568 -0.0260 0.0117 1.0000
6.500 0.9328 0.01254 0.00641 -0.0260 0.0107 1.0000
6.750 0.9555 0.01407 0.00812 -0.0258 0.0097 1.0000
7.000 0.9806 0.01471 0.00882 -0.0257 0.0094 1.0000
7.250 1.0055 0.01534 0.00949 -0.0256 0.0089 1.0000
7.500 1.0288 0.01626 0.01049 -0.0252 0.0084 1.0000
7.750 1.0514 0.01723 0.01152 -0.0248 0.0079 1.0000
8.000 1.0748 0.01797 0.01229 -0.0246 0.0074 1.0000
8.250 1.0969 0.01888 0.01327 -0.0242 0.0071 1.0000
8.500 1.1178 0.02002 0.01448 -0.0236 0.0068 1.0000
8.750 1.1377 0.02136 0.01587 -0.0229 0.0066 1.0000
9.000 1.1559 0.02321 0.01780 -0.0219 0.0064 1.0000
9.250 1.1716 0.02648 0.02124 -0.0206 0.0062 1.0000
9.500 1.1887 0.02860 0.02363 -0.0195 0.0059 1.0000
9.750 1.2067 0.03035 0.02559 -0.0185 0.0056 1.0000
10.000 1.2186 0.03381 0.02937 -0.0170 0.0055 1.0000
10.250 1.2159 0.04272 0.03897 -0.0131 0.0073 1.0000
10.500 1.2199 0.04562 0.04208 -0.0117 0.0072 1.0000
10.750 1.2200 0.04853 0.04519 -0.0101 0.0070 1.0000
11.000 1.2155 0.05129 0.04812 -0.0083 0.0069 1.0000
11.250 1.2038 0.05395 0.05091 -0.0063 0.0069 1.0000
11.500 1.1914 0.05716 0.05427 -0.0057 0.0068 1.0000
11.750 1.1783 0.06092 0.05817 -0.0060 0.0068 1.0000
12.000 1.1641 0.06524 0.06262 -0.0072 0.0067 1.0000
12.250 1.1482 0.07023 0.06775 -0.0091 0.0067 1.0000
12.500 1.1304 0.07591 0.07358 -0.0118 0.0067 1.0000
12.750 1.1115 0.08225 0.08006 -0.0151 0.0067 1.0000
13.000 1.0893 0.08974 0.08770 -0.0192 0.0068 1.0000
13.250 1.0652 0.09839 0.09649 -0.0244 0.0068 1.0000
13.500 1.0366 0.10910 0.10735 -0.0311 0.0070 1.0000
13.750 1.0045 0.12310 0.12148 -0.0399 0.0072 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M16 AIRFOIL (m16-il)