CURTISS CR-1 AIRFOIL (cr1-il)
CURTISS CR-1 AIRFOIL - Curtiss CR-1 general aviation airfoil
| Details | Dat file | Parser | |
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(cr1-il) CURTISS CR-1 AIRFOIL Curtiss CR-1 general aviation airfoil Max thickness 12.2% at 24% chord. Max camber 4.7% at 42% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
CURTISS CR-1 AIRFOIL
61.0 61.0
0.0000000 0.0000000
0.0005000 0.0028877
0.0010000 0.0045082
0.0020000 0.0069026
0.0040000 0.0104752
0.0080000 0.0159130
0.0120000 0.0204608
0.0200000 0.0284739
0.0300000 0.0372190
0.0400000 0.0449223
0.0500000 0.0515424
0.0600000 0.0571529
0.0800000 0.0667125
0.1000000 0.0748305
0.1200000 0.0811676
0.1400000 0.0861657
0.1600000 0.0903037
0.1800000 0.0938086
0.2000000 0.0967598
0.2200000 0.0992238
0.2400000 0.1012466
0.2600000 0.1028692
0.2800000 0.1041314
0.3000000 0.1050716
0.3200000 0.1057211
0.3400000 0.1060956
0.3600000 0.1062082
0.3800000 0.1060724
0.4000000 0.1057010
0.4200000 0.1051022
0.4400000 0.1042715
0.4600000 0.1032027
0.4800000 0.1018895
0.5000000 0.1003252
0.5200000 0.0985024
0.5400000 0.0964112
0.5600000 0.0940411
0.5800000 0.0913811
0.6000000 0.0884199
0.6200000 0.0851617
0.6400000 0.0816512
0.6600000 0.0779406
0.6800000 0.0740830
0.7000000 0.0701324
0.7200000 0.0661238
0.7400000 0.0620373
0.7600000 0.0578435
0.7800000 0.0535129
0.8000000 0.0490156
0.8200000 0.0443404
0.8400000 0.0395362
0.8600000 0.0346645
0.8800000 0.0297881
0.9000000 0.0249688
0.9200000 0.0202277
0.9400000 0.0154453
0.9600000 0.0104761
0.9700000 0.0079079
0.9800000 0.0052997
0.9900000 0.0026651
1.0000000 0.0000173
0.0000000 0.0000000
0.0005000 -.0049013
0.0010000 -.0063669
0.0020000 -.0084885
0.0040000 -.0115467
0.0080000 -.0158525
0.0120000 -.0190076
0.0200000 -.0232770
0.0300000 -.0265473
0.0400000 -.0288161
0.0500000 -.0303860
0.0600000 -.0314227
0.0800000 -.0322357
0.1000000 -.0319040
0.1200000 -.0310798
0.1400000 -.0298930
0.1600000 -.0283204
0.1800000 -.0264458
0.2000000 -.0244737
0.2200000 -.0225857
0.2400000 -.0208260
0.2600000 -.0192004
0.2800000 -.0177150
0.3000000 -.0163755
0.3200000 -.0151843
0.3400000 -.0141278
0.3600000 -.0131887
0.3800000 -.0123493
0.4000000 -.0115922
0.4200000 -.0109030
0.4400000 -.0102799
0.4600000 -.0097239
0.4800000 -.0092366
0.5000000 -.0088192
0.5200000 -.0084708
0.5400000 -.0081817
0.5600000 -.0079399
0.5800000 -.0077333
0.6000000 -.0075498
0.6200000 -.0073791
0.6400000 -.0072182
0.6600000 -.0070659
0.6800000 -.0069210
0.7000000 -.0067822
0.7200000 -.0066495
0.7400000 -.0065270
0.7600000 -.0064199
0.7800000 -.0063335
0.8000000 -.0062730
0.8200000 -.0062355
0.8400000 -.0061851
0.8600000 -.0060768
0.8800000 -.0058660
0.9000000 -.0055081
0.9200000 -.0049546
0.9400000 -.0041417
0.9600000 -.0030058
0.9700000 -.0023222
0.9800000 -.0015842
0.9900000 -.0008099
1.0000000 -.0000173
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Polars for CURTISS CR-1 AIRFOIL (cr1-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| cr1-il | 50,000 | 9 | 22.7 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| cr1-il | 50,000 | 5 | 32.4 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| cr1-il | 100,000 | 9 | 54.7 at α=6° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| cr1-il | 100,000 | 5 | 53 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| cr1-il | 200,000 | 9 | 79.1 at α=4.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| cr1-il | 200,000 | 5 | 72.9 at α=3.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| cr1-il | 500,000 | 9 | 109.5 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| cr1-il | 500,000 | 5 | 89.8 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| cr1-il | 1,000,000 | 9 | 120.9 at α=3.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| cr1-il | 1,000,000 | 5 | 104.3 at α=5.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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