NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il)
NASA/LANGLEY NLF(2)-0415 AIRFOIL - NASA/Langley-DFVLR/Somers(NASA)-Horstmann(DFVLR) NLF(2)-0415 natural laminar flow airfoil
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(nlf415-il) NASA/LANGLEY NLF(2)-0415 AIRFOIL NASA/Langley-DFVLR/Somers(NASA)-Horstmann(DFVLR) NLF(2)-0415 natural laminar flow airfoil Max thickness 15% at 47.1% chord. Max camber 3% at 71.4% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
NASA/LANGLEY NLF(2)-0415 AIRFOIL
67.0 65.0
0.0000283 0.0001977
0.0001580 0.0018782
0.0005781 0.0034513
0.0012354 0.0049151
0.0020767 0.0062942
0.0030483 0.0076132
0.0048657 0.0097598
0.0072362 0.0121647
0.0097802 0.0143864
0.0126772 0.0165900
0.0161151 0.0188727
0.0196734 0.0209524
0.0233271 0.0228708
0.0284762 0.0253213
0.0344332 0.0278754
0.0404768 0.0302213
0.0481623 0.0329347
0.0567129 0.0356738
0.0653312 0.0381994
0.0770819 0.0413574
0.0888958 0.0442664
0.1034498 0.0475683
0.1194057 0.0509036
0.1366659 0.0542431
0.1545967 0.0574695
0.1762957 0.0610990
0.1999052 0.0647528
0.2252574 0.0683730
0.2511119 0.0717686
0.2779291 0.0749936
0.3054984 0.0780026
0.3334588 0.0807457
0.3622074 0.0832480
0.3911308 0.0854336
0.4203734 0.0872973
0.4464551 0.0886623
0.4709578 0.0896782
0.4953701 0.0904177
0.5197378 0.0908721
0.5438408 0.0910307
0.5678078 0.0908768
0.5891203 0.0904519
0.6090986 0.0897875
0.6280169 0.0889052
0.6448200 0.0878954
0.6616064 0.0866422
0.6757030 0.0853765
0.6884416 0.0840321
0.7011539 0.0824596
0.7141083 0.0805784
0.7275568 0.0783021
0.7409475 0.0757130
0.7541982 0.0728412
0.7672220 0.0697441
0.7801957 0.0664444
0.7964010 0.0621125
0.8190836 0.0557941
0.8505346 0.0467394
0.8955132 0.0334861
0.9249468 0.0247040
0.9399937 0.0201262
0.9540721 0.0156569
0.9662560 0.0115595
0.9781089 0.0073526
0.9879752 0.0038258
0.9941875 0.0017348
1.0000000 0.0000000
0.0000283 0.0001977
0.0002160 -.0014055
0.0007030 -.0029037
0.0014684 -.0043042
0.0025414 -.0056681
0.0039070 -.0069317
0.0054110 -.0079697
0.0070198 -.0088373
0.0086994 -.0095897
0.0104158 -.0102821
0.0165490 -.0126481
0.0227369 -.0148649
0.0291896 -.0169172
0.0358008 -.0187691
0.0424562 -.0204598
0.0546129 -.0233186
0.0695675 -.0265412
0.0842934 -.0294302
0.0990596 -.0321051
0.1192222 -.0354972
0.1401296 -.0387463
0.1624827 -.0419577
0.1860574 -.0450866
0.2102588 -.0480371
0.2359435 -.0508779
0.2620312 -.0534583
0.2888914 -.0557890
0.3132250 -.0576002
0.3363004 -.0590362
0.3574037 -.0600921
0.3775230 -.0608539
0.3968363 -.0613459
0.4145228 -.0615713
0.4322107 -.0615507
0.4474670 -.0613098
0.4614996 -.0608672
0.4755211 -.0601643
0.4895594 -.0591512
0.5036273 -.0578040
0.5176603 -.0561485
0.5316613 -.0542105
0.5478030 -.0516619
0.5639025 -.0488660
0.5799760 -.0459166
0.6175163 -.0388157
0.6503252 -.0326355
0.6807982 -.0270767
0.7059723 -.0226781
0.7311841 -.0185007
0.7556095 -.0147144
0.7800767 -.0112075
0.8032020 -.0081807
0.8263655 -.0054618
0.8476593 -.0032672
0.8689834 -.0013869
0.8879449 -.0000039
0.9069262 0.0010752
0.9231043 0.0017266
0.9392914 0.0021036
0.9522953 0.0021864
0.9652988 0.0020429
0.9767429 0.0016978
0.9867086 0.0011870
0.9937244 0.0006571
1.0000000 0.0000000
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Polars for NASA/LANGLEY NLF(2)-0415 AIRFOIL (nlf415-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| nlf415-il | 50,000 | 9 | 23.4 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf415-il | 50,000 | 5 | 21.1 at α=9.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf415-il | 100,000 | 9 | 27.2 at α=8.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf415-il | 100,000 | 5 | 28.6 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf415-il | 200,000 | 9 | 36.2 at α=6.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf415-il | 200,000 | 5 | 45.7 at α=5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf415-il | 500,000 | 9 | 85.3 at α=4.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf415-il | 500,000 | 5 | 82.7 at α=3.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| nlf415-il | 1,000,000 | 9 | 118.4 at α=3.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| nlf415-il | 1,000,000 | 5 | 107.5 at α=2.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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