OAF128 AIRFOIL (oaf128-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: OAF128 AIRFOIL (oaf128-il) Reynolds number: 50,000 Max Cl/Cd: 31.08 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-oaf128-il-50000-n5.txt Download as CSV file: xf-oaf128-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: OAF128 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.7864 0.08688 0.07931 0.0045 1.0000 0.0756
-10.500 -0.8701 0.06649 0.05852 -0.0132 1.0000 0.0732
-10.250 -0.8916 0.05996 0.05171 -0.0177 1.0000 0.0737
-10.000 -0.8965 0.05521 0.04664 -0.0200 1.0000 0.0753
-9.750 -0.8955 0.05072 0.04174 -0.0217 1.0000 0.0773
-9.500 -0.8858 0.04723 0.03792 -0.0226 1.0000 0.0797
-9.250 -0.8683 0.04535 0.03603 -0.0227 1.0000 0.0830
-9.000 -0.8529 0.04247 0.03276 -0.0234 1.0000 0.0875
-8.750 -0.8331 0.04058 0.03085 -0.0236 1.0000 0.0917
-8.500 -0.8130 0.03836 0.02836 -0.0240 1.0000 0.0980
-8.250 -0.7913 0.03676 0.02678 -0.0242 1.0000 0.1045
-8.000 -0.7688 0.03492 0.02479 -0.0245 1.0000 0.1132
-7.750 -0.7453 0.03327 0.02300 -0.0249 1.0000 0.1240
-7.500 -0.7209 0.03189 0.02156 -0.0253 1.0000 0.1365
-7.250 -0.6958 0.03071 0.02036 -0.0257 1.0000 0.1505
-7.000 -0.6699 0.02971 0.01936 -0.0262 1.0000 0.1659
-6.750 -0.6435 0.02882 0.01847 -0.0266 1.0000 0.1826
-6.500 -0.6167 0.02798 0.01762 -0.0270 1.0000 0.2002
-6.250 -0.5895 0.02722 0.01688 -0.0274 1.0000 0.2178
-6.000 -0.5620 0.02654 0.01623 -0.0277 1.0000 0.2349
-5.750 -0.5340 0.02588 0.01556 -0.0280 1.0000 0.2517
-5.500 -0.5057 0.02524 0.01491 -0.0282 1.0000 0.2685
-5.250 -0.4772 0.02465 0.01431 -0.0285 1.0000 0.2853
-5.000 -0.4486 0.02409 0.01374 -0.0288 1.0000 0.3022
-4.750 -0.4198 0.02359 0.01324 -0.0291 1.0000 0.3195
-4.500 -0.3909 0.02311 0.01277 -0.0294 1.0000 0.3369
-4.250 -0.3621 0.02268 0.01236 -0.0297 1.0000 0.3548
-4.000 -0.3332 0.02227 0.01198 -0.0301 1.0000 0.3728
-3.750 -0.3047 0.02191 0.01165 -0.0303 1.0000 0.3911
-3.500 -0.2762 0.02158 0.01138 -0.0306 1.0000 0.4097
-3.250 -0.2481 0.02128 0.01116 -0.0310 1.0000 0.4287
-3.000 -0.2208 0.02103 0.01099 -0.0312 1.0000 0.4478
-2.750 -0.1953 0.02082 0.01089 -0.0314 1.0000 0.4662
-2.500 -0.1524 0.02069 0.01078 -0.0345 0.9617 0.4879
-2.250 -0.1106 0.02055 0.01063 -0.0370 0.9287 0.5082
-2.000 -0.0751 0.02047 0.01053 -0.0380 0.8970 0.5272
-1.750 -0.0450 0.02044 0.01046 -0.0378 0.8672 0.5455
-1.500 -0.0180 0.02043 0.01041 -0.0370 0.8395 0.5634
-1.250 0.0076 0.02042 0.01036 -0.0358 0.8140 0.5820
-1.000 0.0327 0.02040 0.01034 -0.0345 0.7897 0.6021
-0.750 0.0574 0.02039 0.01033 -0.0331 0.7675 0.6238
-0.500 0.0820 0.02037 0.01032 -0.0317 0.7467 0.6474
-0.250 0.1071 0.02036 0.01030 -0.0305 0.7271 0.6723
0.000 0.1312 0.02033 0.01032 -0.0290 0.7086 0.6963
0.250 0.1556 0.02031 0.01033 -0.0275 0.6908 0.7226
0.500 0.1795 0.02028 0.01034 -0.0260 0.6738 0.7520
0.750 0.2027 0.02024 0.01034 -0.0242 0.6575 0.7859
1.000 0.2245 0.02015 0.01030 -0.0221 0.6416 0.8250
1.250 0.2474 0.02001 0.01017 -0.0202 0.6260 0.8746
1.500 0.2794 0.01986 0.01002 -0.0203 0.6086 0.9481
1.750 0.3143 0.01997 0.01000 -0.0219 0.5909 1.0000
2.000 0.3455 0.02027 0.01016 -0.0229 0.5746 1.0000
2.250 0.3761 0.02059 0.01036 -0.0235 0.5592 1.0000
2.500 0.4061 0.02092 0.01058 -0.0239 0.5445 1.0000
2.750 0.4355 0.02125 0.01078 -0.0241 0.5311 1.0000
3.000 0.4651 0.02162 0.01110 -0.0245 0.5161 1.0000
3.250 0.4947 0.02202 0.01147 -0.0248 0.5012 1.0000
3.500 0.5239 0.02242 0.01183 -0.0251 0.4871 1.0000
3.750 0.5527 0.02281 0.01215 -0.0251 0.4734 1.0000
4.000 0.5815 0.02321 0.01252 -0.0251 0.4591 1.0000
4.250 0.6102 0.02366 0.01298 -0.0253 0.4435 1.0000
4.500 0.6386 0.02409 0.01342 -0.0254 0.4277 1.0000
4.750 0.6665 0.02449 0.01380 -0.0253 0.4109 1.0000
5.000 0.6942 0.02486 0.01414 -0.0252 0.3932 1.0000
5.250 0.7216 0.02523 0.01450 -0.0251 0.3750 1.0000
5.500 0.7488 0.02562 0.01489 -0.0250 0.3566 1.0000
5.750 0.7758 0.02606 0.01532 -0.0249 0.3384 1.0000
6.000 0.8025 0.02654 0.01580 -0.0247 0.3202 1.0000
6.250 0.8288 0.02706 0.01631 -0.0246 0.3019 1.0000
6.500 0.8546 0.02765 0.01689 -0.0245 0.2835 1.0000
6.750 0.8799 0.02832 0.01752 -0.0243 0.2652 1.0000
7.000 0.9047 0.02911 0.01829 -0.0241 0.2471 1.0000
7.250 0.9288 0.03003 0.01920 -0.0240 0.2297 1.0000
7.500 0.9522 0.03105 0.02017 -0.0237 0.2142 1.0000
7.750 0.9750 0.03215 0.02124 -0.0235 0.2001 1.0000
8.000 0.9974 0.03341 0.02259 -0.0232 0.1870 1.0000
8.250 1.0192 0.03473 0.02400 -0.0230 0.1759 1.0000
8.500 1.0405 0.03600 0.02522 -0.0226 0.1671 1.0000
8.750 1.0610 0.03744 0.02679 -0.0222 0.1586 1.0000
9.000 1.0810 0.03891 0.02831 -0.0218 0.1517 1.0000
9.250 1.1000 0.04049 0.03002 -0.0213 0.1451 1.0000
9.500 1.1194 0.04197 0.03146 -0.0208 0.1399 1.0000
9.750 1.1349 0.04403 0.03382 -0.0203 0.1343 1.0000
10.000 1.1523 0.04563 0.03545 -0.0198 0.1298 1.0000
10.250 1.1674 0.04767 0.03760 -0.0192 0.1259 1.0000
10.500 1.1762 0.05030 0.04056 -0.0187 0.1219 1.0000
10.750 1.1873 0.05245 0.04284 -0.0180 0.1184 1.0000
11.000 1.2054 0.05400 0.04429 -0.0174 0.1154 1.0000
11.250 1.1993 0.05783 0.04854 -0.0169 0.1130 1.0000
11.500 1.1853 0.06194 0.05299 -0.0164 0.1111 1.0000
11.750 1.1660 0.06693 0.05827 -0.0174 0.1097 1.0000
12.000 1.1386 0.07377 0.06539 -0.0205 0.1086 1.0000
12.250 1.0564 0.09130 0.08334 -0.0325 0.1095 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF128 AIRFOIL (oaf128-il)