NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 500,000 Max Cl/Cd: 63.46 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64110-il-500000.txt Download as CSV file: xf-n64110-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-110 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.7101 0.05867 0.05586 -0.0310 1.0000 0.0232
-8.750 -0.7148 0.05490 0.05187 -0.0303 1.0000 0.0232
-8.500 -0.7159 0.05101 0.04772 -0.0294 1.0000 0.0232
-8.250 -0.7353 0.04110 0.03742 -0.0285 1.0000 0.0241
-8.000 -0.7432 0.02971 0.02507 -0.0251 1.0000 0.0174
-7.750 -0.7271 0.02671 0.02174 -0.0239 1.0000 0.0174
-7.500 -0.7101 0.02380 0.01853 -0.0230 1.0000 0.0178
-7.250 -0.6884 0.02303 0.01776 -0.0227 1.0000 0.0188
-7.000 -0.6667 0.02200 0.01665 -0.0220 1.0000 0.0198
-6.750 -0.6466 0.02010 0.01452 -0.0208 1.0000 0.0201
-6.500 -0.6274 0.01865 0.01290 -0.0194 1.0000 0.0206
-6.250 -0.6100 0.01751 0.01164 -0.0176 0.9998 0.0212
-6.000 -0.5747 0.01624 0.01022 -0.0194 0.9949 0.0220
-5.750 -0.5386 0.01583 0.00971 -0.0213 0.9893 0.0231
-5.500 -0.5062 0.01364 0.00743 -0.0228 0.9841 0.0248
-5.250 -0.4729 0.01278 0.00654 -0.0244 0.9763 0.0263
-5.000 -0.4383 0.01210 0.00582 -0.0261 0.9692 0.0280
-4.750 -0.4074 0.01158 0.00525 -0.0269 0.9586 0.0303
-4.500 -0.3800 0.01104 0.00463 -0.0269 0.9464 0.0328
-4.250 -0.3554 0.01045 0.00399 -0.0264 0.9338 0.0372
-4.000 -0.3304 0.01015 0.00361 -0.0257 0.9220 0.0415
-3.750 -0.3060 0.00973 0.00317 -0.0250 0.9107 0.0509
-3.500 -0.2818 0.00922 0.00279 -0.0244 0.8996 0.0868
-3.250 -0.2621 0.00770 0.00222 -0.0239 0.8878 0.3271
-3.000 -0.2398 0.00684 0.00203 -0.0233 0.8771 0.5136
-2.750 -0.2140 0.00667 0.00196 -0.0228 0.8676 0.5672
-2.500 -0.1875 0.00657 0.00190 -0.0224 0.8576 0.6002
-2.250 -0.1604 0.00651 0.00184 -0.0222 0.8477 0.6237
-2.000 -0.1334 0.00647 0.00179 -0.0219 0.8384 0.6432
-1.750 -0.1068 0.00644 0.00180 -0.0215 0.8286 0.6743
-1.500 -0.0802 0.00643 0.00186 -0.0210 0.8190 0.7034
-1.250 -0.0529 0.00644 0.00183 -0.0208 0.8103 0.7169
-1.000 -0.0250 0.00643 0.00179 -0.0207 0.8007 0.7279
-0.750 0.0027 0.00642 0.00177 -0.0206 0.7915 0.7373
-0.500 0.0303 0.00642 0.00174 -0.0205 0.7827 0.7458
0.000 0.0861 0.00642 0.00173 -0.0204 0.7646 0.7643
0.250 0.1140 0.00642 0.00173 -0.0203 0.7557 0.7735
0.500 0.1419 0.00644 0.00176 -0.0203 0.7467 0.7831
0.750 0.1694 0.00646 0.00178 -0.0202 0.7385 0.7922
1.000 0.1973 0.00647 0.00182 -0.0201 0.7293 0.8020
1.250 0.2250 0.00650 0.00187 -0.0200 0.7198 0.8121
1.500 0.2521 0.00653 0.00191 -0.0198 0.7099 0.8213
1.750 0.2794 0.00655 0.00194 -0.0196 0.6970 0.8311
2.000 0.3059 0.00656 0.00193 -0.0191 0.6725 0.8416
2.250 0.3316 0.00659 0.00194 -0.0185 0.6444 0.8515
2.500 0.3580 0.00665 0.00200 -0.0181 0.6248 0.8619
2.750 0.3845 0.00672 0.00207 -0.0177 0.6043 0.8728
3.000 0.4101 0.00681 0.00214 -0.0171 0.5791 0.8834
3.250 0.4351 0.00693 0.00225 -0.0164 0.5458 0.8944
3.500 0.4582 0.00722 0.00235 -0.0154 0.4791 0.9064
3.750 0.4742 0.00845 0.00273 -0.0137 0.2709 0.9204
4.000 0.4898 0.00980 0.00329 -0.0121 0.0905 0.9363
4.250 0.5119 0.01035 0.00370 -0.0111 0.0516 0.9532
4.500 0.5405 0.01078 0.00414 -0.0114 0.0412 0.9697
4.750 0.5744 0.01151 0.00489 -0.0130 0.0332 0.9854
5.000 0.6069 0.01194 0.00536 -0.0143 0.0304 1.0000
5.250 0.6318 0.01247 0.00592 -0.0141 0.0279 1.0000
5.500 0.6555 0.01324 0.00670 -0.0137 0.0260 1.0000
5.750 0.6763 0.01456 0.00809 -0.0128 0.0242 1.0000
6.000 0.7022 0.01501 0.00858 -0.0126 0.0231 1.0000
6.250 0.7265 0.01580 0.00945 -0.0122 0.0222 1.0000
6.500 0.7507 0.01672 0.01042 -0.0117 0.0213 1.0000
6.750 0.7750 0.01770 0.01147 -0.0113 0.0206 1.0000
7.000 0.7993 0.01874 0.01258 -0.0109 0.0198 1.0000
7.250 0.8232 0.01993 0.01383 -0.0105 0.0192 1.0000
7.500 0.8457 0.02188 0.01589 -0.0100 0.0183 1.0000
7.750 0.8655 0.02507 0.01940 -0.0092 0.0176 1.0000
8.000 0.8868 0.02672 0.02131 -0.0084 0.0174 1.0000
8.250 0.9051 0.02927 0.02417 -0.0073 0.0172 1.0000
8.500 0.9196 0.03247 0.02773 -0.0059 0.0172 1.0000
8.750 0.9306 0.03727 0.03278 -0.0047 0.0181 1.0000
11.500 0.7453 0.11065 0.10848 -0.0210 0.0228 1.0000
11.750 0.7405 0.11759 0.11541 -0.0250 0.0221 1.0000
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