S9000 (9%) (s9000-il)
S9000 (9%) - Selig S9000 (9%) low Reynolds number airfoil used on the Blackhawk R/C sailplane
| Details | Dat file | Parser | |
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(s9000-il) S9000 (9%) Selig S9000 (9%) low Reynolds number airfoil used on the Blackhawk R/C sailplane Max thickness 9% at 28.2% chord. Max camber 2.1% at 44.8% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
S9000 (9%)
1.000000 0.000000
0.999170 0.000110
0.996710 0.000500
0.992700 0.001210
0.987230 0.002250
0.980400 0.003610
0.972290 0.005230
0.962940 0.007050
0.952400 0.009000
0.940650 0.011030
0.927680 0.013140
0.913520 0.015360
0.898220 0.017680
0.881840 0.020100
0.864430 0.022610
0.846060 0.025200
0.826770 0.027850
0.806640 0.030560
0.785740 0.033310
0.764120 0.036080
0.741850 0.038850
0.719020 0.041600
0.695660 0.044280
0.671850 0.046910
0.647650 0.049460
0.623130 0.051910
0.598360 0.054220
0.573380 0.056410
0.548280 0.058450
0.523120 0.060330
0.497960 0.062020
0.472860 0.063510
0.447880 0.064810
0.423110 0.065900
0.398590 0.066740
0.374360 0.067350
0.350500 0.067720
0.327090 0.067850
0.304150 0.067710
0.281730 0.067300
0.259890 0.066650
0.238710 0.065730
0.218200 0.064540
0.198410 0.063080
0.179390 0.061380
0.161200 0.059420
0.143860 0.057190
0.127380 0.054720
0.111800 0.052040
0.097190 0.049130
0.083550 0.045990
0.070870 0.042650
0.059190 0.039160
0.048560 0.035530
0.038960 0.031740
0.030390 0.027850
0.022870 0.023920
0.016430 0.019970
0.011080 0.015970
0.006750 0.012020
0.003450 0.008200
0.001230 0.004630
0.000130 0.001360
0.000200 -0.001390
0.002030 -0.003680
0.005370 -0.006000
0.010060 -0.008270
0.016060 -0.010420
0.023360 -0.012430
0.031940 -0.014310
0.041770 -0.016030
0.052840 -0.017580
0.065120 -0.018970
0.078590 -0.020180
0.093220 -0.021220
0.108960 -0.022080
0.125800 -0.022760
0.143700 -0.023260
0.162600 -0.023600
0.182460 -0.023780
0.203250 -0.023790
0.224890 -0.023670
0.247350 -0.023400
0.270550 -0.023010
0.294450 -0.022490
0.318960 -0.021880
0.344030 -0.021170
0.369580 -0.020370
0.395530 -0.019510
0.421830 -0.018580
0.448380 -0.017590
0.475100 -0.016570
0.501930 -0.015510
0.528770 -0.014420
0.555530 -0.013310
0.582150 -0.012170
0.608530 -0.010990
0.634590 -0.009770
0.660300 -0.008440
0.685670 -0.007070
0.710620 -0.005700
0.735070 -0.004380
0.758930 -0.003130
0.782120 -0.001970
0.804570 -0.000920
0.826180 0.000000
0.846870 0.000790
0.866560 0.001420
0.885180 0.001910
0.902650 0.002240
0.918890 0.002420
0.933820 0.002450
0.947380 0.002340
0.959500 0.002120
0.970110 0.001810
0.979170 0.001430
0.986630 0.001040
0.992470 0.000650
0.996650 0.000320
0.999160 0.000090
1.000000 0.000000
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Similar airfoils
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Polars for S9000 (9%) (s9000-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| s9000-il | 50,000 | 9 | 34.7 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s9000-il | 50,000 | 5 | 37.9 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s9000-il | 100,000 | 9 | 53.3 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s9000-il | 100,000 | 5 | 52.7 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s9000-il | 200,000 | 9 | 71 at α=4.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s9000-il | 200,000 | 5 | 66.5 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s9000-il | 500,000 | 9 | 93 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s9000-il | 500,000 | 5 | 84.4 at α=4.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| s9000-il | 1,000,000 | 9 | 108.2 at α=4° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| s9000-il | 1,000,000 | 5 | 97.3 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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