ULTIMATE (ultimate-il)
ULTIMATE - Carl Goldberg R/C Ultimate Bipe airfoil
| Details | Dat file | Parser | |
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(ultimate-il) ULTIMATE Carl Goldberg R/C Ultimate Bipe airfoil Max thickness 12.8% at 34.2% chord. Max camber 0% at 0% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
ULTIMATE 1.000000 0.002886 0.997540 0.002863 0.990700 0.003094 0.980370 0.004027 0.966980 0.005805 0.950440 0.008274 0.930640 0.011103 0.907750 0.013982 0.882020 0.016972 0.853700 0.020447 0.823090 0.024581 0.790480 0.029046 0.756160 0.033444 0.720430 0.037602 0.683590 0.041508 0.645940 0.045227 0.607780 0.048794 0.569370 0.052159 0.530990 0.055209 0.492650 0.057869 0.454350 0.060114 0.416380 0.061916 0.378870 0.063171 0.342040 0.063718 0.306090 0.063459 0.271200 0.062388 0.237600 0.060562 0.205490 0.058079 0.175040 0.055033 0.146480 0.051474 0.119990 0.047391 0.095760 0.042768 0.073950 0.037669 0.054680 0.032282 0.038110 0.026863 0.024330 0.021542 0.013380 0.016225 0.005480 0.010723 0.000980 0.004912 0.000000 0.000533 0.000980 -0.003898 0.005480 -0.009945 0.013380 -0.015812 0.024330 -0.021543 0.038110 -0.027258 0.054680 -0.032984 0.073950 -0.038571 0.095760 -0.043818 0.119990 -0.048597 0.146480 -0.052838 0.175040 -0.056512 0.205490 -0.059595 0.237600 -0.062030 0.271200 -0.063725 0.306090 -0.064621 0.342040 -0.064754 0.378870 -0.064245 0.416380 -0.063248 0.454350 -0.061870 0.492650 -0.060119 0.530990 -0.057952 0.569370 -0.055317 0.607780 -0.052206 0.645940 -0.048690 0.683590 -0.044885 0.720430 -0.040929 0.756160 -0.036957 0.790480 -0.033065 0.823090 -0.029257 0.853700 -0.025505 0.882020 -0.021977 0.907750 -0.018951 0.930640 -0.016575 0.950440 -0.014660 0.966980 -0.012681 0.980370 -0.010169 0.990700 -0.007092 0.997540 -0.004180 0.999983 -0.002918 |
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Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
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Polars for ULTIMATE (ultimate-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| ultimate-il | 50,000 | 9 | 28.4 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 50,000 | 5 | 27.5 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 100,000 | 9 | 39 at α=5.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 100,000 | 5 | 36.1 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 200,000 | 9 | 47.8 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 200,000 | 5 | 42 at α=4.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 500,000 | 9 | 55.4 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 500,000 | 5 | 53.5 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| ultimate-il | 1,000,000 | 9 | 63.4 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| ultimate-il | 1,000,000 | 5 | 67 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||