RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL (r1082-il)
RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL - Roncz 1082 Voyager root outer aft wing airfoil
| Details | Dat file | Parser | |
|
(r1082-il) RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL Roncz 1082 Voyager root outer aft wing airfoil Max thickness 18% at 40% chord. Max camber 3.4% at 40% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL
60.0 60.0
0.0000000 0.0000000
0.0004520 0.0039110
0.0009510 0.0056950
0.0019510 0.0082470
0.0039490 0.0120530
0.0079480 0.0180080
0.0119460 0.0228400
0.0199440 0.0307030
0.0299420 0.0386930
0.0399400 0.0455120
0.0499390 0.0515320
0.0599370 0.0569570
0.0799350 0.0664820
0.0999330 0.0746830
0.1199320 0.0818710
0.1399310 0.0882370
0.1599300 0.0938990
0.1799290 0.0989470
0.1999290 0.1034390
0.2199280 0.1074190
0.2399280 0.1109260
0.2599280 0.1139830
0.2799280 0.1166120
0.2999280 0.1188270
0.3199290 0.1206340
0.3399290 0.1220380
0.3599300 0.1230420
0.3799310 0.1236370
0.3999310 0.1238070
0.4199330 0.1235150
0.4399340 0.1226940
0.4599350 0.1212780
0.4799370 0.1192710
0.4999380 0.1167020
0.5199401 0.1136460
0.5399430 0.1102040
0.5599450 0.1064680
0.5799470 0.1024930
0.5999490 0.0983210
0.6199520 0.0939820
0.6399540 0.0895040
0.6599560 0.0849060
0.6799590 0.0802120
0.6999610 0.0754460
0.7199640 0.0706200
0.7399660 0.0657480
0.7599690 0.0608440
0.7799710 0.0559240
0.7999740 0.0509970
0.8199770 0.0460700
0.8399790 0.0411800
0.8599820 0.0362530
0.8799840 0.0313740
0.8999870 0.0265100
0.9199890 0.0216500
0.9399920 0.0167760
0.9599940 0.0117840
0.9799970 0.0061970
0.9899990 0.0030520
1.0000000 0.0000000
0.0000000 0.0000000
0.0004540 -.0037650
0.0009550 -.0053450
0.0019560 -.0073440
0.0039570 -.0098130
0.0079580 -.0129920
0.0119590 -.0154700
0.0199600 -.0195830
0.0299620 -.0236230
0.0399640 -.0268850
0.0499650 -.0296500
0.0599660 -.0320470
0.0799690 -.0360490
0.0999710 -.0392940
0.1199730 -.0420120
0.1399740 -.0443340
0.1599760 -.0463420
0.1799770 -.0480900
0.1999790 -.0496150
0.2199800 -.0509430
0.2399810 -.0520930
0.2599830 -.0530820
0.2799840 -.0539250
0.2999850 -.0546250
0.3199860 -.0551910
0.3399870 -.0556270
0.3599880 -.0559420
0.3799890 -.0561320
0.3999900 -.0562000
0.4199910 -.0561430
0.4399920 -.0559630
0.4599930 -.0556610
0.4799940 -.0552340
0.4999940 -.0546790
0.5199950 -.0539890
0.5399960 -.0531610
0.5599960 -.0521910
0.5799970 -.0510690
0.5999980 -.0497720
0.6199980 -.0482750
0.6399980 -.0465240
0.6599990 -.0444570
0.6799990 -.0420410
0.6999990 -.0392750
0.7199990 -.0361780
0.7399990 -.0328390
0.7599980 -.0293590
0.7799980 -.0258180
0.7999980 -.0222890
0.8199980 -.0188310
0.8399980 -.0154910
0.8599970 -.0123150
0.8799970 -.0093480
0.8999980 -.0066370
0.9199980 -.0042310
0.9399980 -.0021880
0.9599980 -.0005910
0.9799990 0.0003410
0.9899990 0.0004080
1.0000000 0.0000000
|
No parser warnings |
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for RONCZ 1082 VOYAGER ROOT OUTER AFT WING AIRFOIL (r1082-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| r1082-il | 50,000 | 9 | 4.2 at α=12° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| r1082-il | 50,000 | 5 | 12.1 at α=14.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| r1082-il | 100,000 | 9 | 25.4 at α=13.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| r1082-il | 100,000 | 5 | 21.5 at α=12.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| r1082-il | 200,000 | 9 | 60.9 at α=9.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| r1082-il | 200,000 | 5 | 61.2 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| r1082-il | 500,000 | 9 | 105.1 at α=9° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| r1082-il | 500,000 | 5 | 102.2 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| r1082-il | 1,000,000 | 9 | 136.8 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| r1082-il | 1,000,000 | 5 | 127.1 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||