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| Updated with Smoothed Pressure Coefficient Curves |
| Copyright © 2000- 2009 by Richard W. Fraser All Rights Reserved |
| Click on Graph below to Enlarge. An additional click on a viewer that generally shows up in the lower right part of the widow will greatly enlarge and sharpen the image. |
| Click on Graph below to Enlarge. An additional click on a viewer that generally shows up in the lower right part of the widow will greatly enlarge and sharpen the image. |
1.00000000 0.00000000 Trailing Edge (Starting on top airfoil surface)
0.99931477 0.00013733
0.99726095 0.00054754
0.99384417 0.00122524
0.98907380 0.00216137
0.98296291 0.00334296
0.97552826 0.00475297
0.96679021 0.00636997
0.95677273 0.00816790
0.94550326 0.01009786
0.93301270 0.01208833
0.90797428 0.01633209
0.88293587 0.02066729
0.85574811 0.02533268
0.82468880 0.03088362
0.78391318 0.03818513
0.73672200 0.04669967
0.69561525 0.05435195
0.65450850 0.06192413
0.62940952 0.06643887
0.60395585 0.07099621
0.57821723 0.07559404
0.55226423 0.08022180
0.52616798 0.08487237
0.50000000 0.08955391
0.46807074 0.09519783
0.43576126 0.10045917
0.41143408 0.10411721
0.39109807 0.10696538
0.37076206 0.10962554
0.34896891 0.11217650
0.32754149 0.11425011
0.30574361 0.11590032
0.28685302 0.11696838
0.26845381 0.11755155
0.24765233 0.11754128
0.22799738 0.11703685
0.20593541 0.11545316
0.18518853 0.11317932
0.16534368 0.11005294
0.14874617 0.10665485
0.13151723 0.10238931
0.11230381 0.09648138
0.09480426 0.08970078
0.08001082 0.08301706
0.06639004 0.07583083
0.05250233 0.06745359
0.03878766 0.05754542
0.02845287 0.04864890
0.01923802 0.03932753
0.01185536 0.03002385
0.00603546 0.02080346
0.00273905 0.01326182
0.00171214 0.01031333
0.00068523 0.00643130
0.00034262 0.00452274
0.00000000 0.00000000 Leading edge
0.00034262 -0.00352975
0.00068523 -0.00497204
0.00171214 -0.00785303
0.00273905 -0.00963475
0.00603546 -0.01330622
0.01185536 -0.01714639
0.01923802 -0.02033901
0.02845287 -0.02295219
0.03878766 -0.02462798
0.05250233 -0.02577501
0.06639004 -0.02638832
0.08001082 -0.02677841
0.09480426 -0.02697420
0.13151723 -0.02733398
0.14874617 -0.02750706
0.16534368 -0.02771527
0.18518853 -0.02817871
0.20593541 -0.02888552
0.22799738 -0.02989879
0.24765233 -0.03089756
0.26845381 -0.03201640
0.28685302 -0.03302760
0.30574361 -0.03409968
0.32754149 -0.03537136
0.34896891 -0.03660603
0.37076206 -0.03774686
0.39109807 -0.03871979
0.41143408 -0.03961421
0.43576126 -0.04056140
0.46807074 -0.04151628
0.50000000 -0.04205002
0.52616798 -0.04223773
0.55226423 -0.04215111
0.57821723 -0.04184985
0.60395585 -0.04143548
0.62940952 -0.04085247
0.65450850 -0.03999924
0.69561525 -0.03819881
0.73672200 -0.03569570
0.78391318 -0.03172289
0.82468880 -0.02767588
0.85574811 -0.02415565
0.88293587 -0.02071138
0.90797428 -0.01712793
0.93301270 -0.01315808
0.94550326 -0.01100354
0.95677273 -0.00899249
0.96679021 -0.00707660
0.97552826 -0.00531762
0.98296291 -0.00376053
0.98907380 -0.00244137
0.99384417 -0.00138812
0.99726095 -0.00062159
0.99931477 -0.00015608
1.00000000 0.00000000 (Trailing edge ending at lower airfoil surface)
| F5Fras15 Airfoil Coordinates (For General Release December, 2007) |
The F5 Prefix now supersedes the previous F3 and F4 series
prefixes due to the smoothing of the Pressure Coefficient (Cp)
curve for cruise and landing/stall speeds. There are enough
numbers to go around. You do not have to plot all of them to make
a smooth airfoil curve, unless you want to.
Multiply the Percent values (X,Y) given below by the chord length
you desire.
X% (Chord) Y% (±Vertical)
prefixes due to the smoothing of the Pressure Coefficient (Cp)
curve for cruise and landing/stall speeds. There are enough
numbers to go around. You do not have to plot all of them to make
a smooth airfoil curve, unless you want to.
Multiply the Percent values (X,Y) given below by the chord length
you desire.
X% (Chord) Y% (±Vertical)
The F5 Prefix now supersedes the previous F3 and F4 series
prefixes due to the smoothing of the Pressure Coefficient (Cp)
curve for cruise and landing/stall speeds. There are enough
numbers to go around. You do not have to plot all of them to make
a smooth airfoil curve, unless you want to.
Multiply the Percent values (X,Y) given below by the chord length
you desire.
X% (Chord) Y% (±Vertical)
1.00000000 0.00000000 Trailing Edge (Starting on top airfoil surface)
0.99931477 0.00011773
0.99726095 0.00046942
0.99384417 0.00105063
0.98907380 0.00185384
0.98296291 0.00286834
0.97552826 0.00408010
0.96679021 0.00547153
0.95677273 0.00702132
0.94550326 0.00868796
0.93301270 0.01040148
0.90797428 0.01409645
0.88293587 0.01790256
0.85574811 0.02202610
0.82468880 0.02697095
0.78391318 0.03351420
0.73672200 0.04119440
0.69561525 0.04816814
0.65450850 0.05511409
0.62940952 0.05927017
0.60395585 0.06348405
0.57821723 0.06774699
0.55226423 0.07204542
0.52616798 0.07637947
0.50000000 0.08076075
0.46807074 0.08606324
0.43576126 0.09103684
0.41143408 0.09451375
0.39109807 0.09723138
0.37076206 0.09977881
0.34896891 0.10223555
0.32754149 0.10425311
0.30574361 0.10587802
0.28685302 0.10694635
0.26845381 0.10755812
0.24765233 0.10762329
0.22799738 0.10721930
0.20593541 0.10580913
0.18518853 0.10373444
0.16534368 0.10084792
0.14874617 0.09769078
0.13151723 0.09372181
0.11230381 0.08821938
0.09480426 0.08190511
0.08001082 0.07568104
0.06639004 0.06900103
0.05250233 0.06122449
0.03878766 0.05205498
0.02845287 0.04386485
0.01923802 0.03534089
0.01185536 0.02687216
0.00603546 0.01852441
0.00273905 0.01173198
0.00171214 0.00909954
0.00068523 0.00566938
0.00034262 0.00398471
0.00000000 0.00000000 Leading edge
0.00034262 -0.00299172
0.00068523 -0.00421012
0.00171214 -0.00663924
0.00273905 -0.00810491
0.00617992 -0.01112386
0.01130043 -0.01377240
0.01915776 -0.01633266
0.02885704 -0.01822030
0.03878766 -0.01913754
0.05258885 -0.01954609
0.06639004 -0.01955852
0.08022130 -0.01943995
0.09489446 -0.01917703
0.11216705 -0.01891281
0.12882187 -0.01869441
0.14547669 -0.01855429
0.16543470 -0.01851024
0.18533980 -0.01873618
0.20594016 -0.01924163
0.22799738 -0.02008124
0.24822560 -0.02100703
0.26845381 -0.02202297
0.28685302 -0.02300557
0.30574361 -0.02407739
0.32561695 -0.02525884
0.34896891 -0.02666508
0.37076206 -0.02790013
0.39109807 -0.02898579
0.41143408 -0.03001075
0.43576126 -0.03113907
0.46788063 -0.03237555
0.49702431 -0.03318960
0.52616798 -0.03374483
0.55226423 -0.03397473
0.57821723 -0.03400280
0.60312424 -0.03392726
0.62940952 -0.03368377
0.65768348 -0.03311366
0.68595745 -0.03229143
0.71198438 -0.03148645
0.76083005 -0.02863288
0.78667318 -0.02684891
0.80837777 -0.02512088
0.82399613 -0.02382386
0.85523283 -0.02090057
0.88293587 -0.01794665
0.90390685 -0.01540877
0.92487783 -0.01264436
0.94233865 -0.01007167
0.95677273 -0.00784591
0.96679021 -0.00617816
0.97552826 -0.00464475
0.98296291 -0.00328591
0.98907380 -0.00213384
0.99384417 -0.00121351
0.99726095 -0.00054347
0.99931477 -0.00013648
1.00000000 0.00000000 Trailing Edge (Ending at lower airfoil surface)
prefixes due to the smoothing of the Pressure Coefficient (Cp)
curve for cruise and landing/stall speeds. There are enough
numbers to go around. You do not have to plot all of them to make
a smooth airfoil curve, unless you want to.
Multiply the Percent values (X,Y) given below by the chord length
you desire.
X% (Chord) Y% (±Vertical)
1.00000000 0.00000000 Trailing Edge (Starting on top airfoil surface)
0.99931477 0.00011773
0.99726095 0.00046942
0.99384417 0.00105063
0.98907380 0.00185384
0.98296291 0.00286834
0.97552826 0.00408010
0.96679021 0.00547153
0.95677273 0.00702132
0.94550326 0.00868796
0.93301270 0.01040148
0.90797428 0.01409645
0.88293587 0.01790256
0.85574811 0.02202610
0.82468880 0.02697095
0.78391318 0.03351420
0.73672200 0.04119440
0.69561525 0.04816814
0.65450850 0.05511409
0.62940952 0.05927017
0.60395585 0.06348405
0.57821723 0.06774699
0.55226423 0.07204542
0.52616798 0.07637947
0.50000000 0.08076075
0.46807074 0.08606324
0.43576126 0.09103684
0.41143408 0.09451375
0.39109807 0.09723138
0.37076206 0.09977881
0.34896891 0.10223555
0.32754149 0.10425311
0.30574361 0.10587802
0.28685302 0.10694635
0.26845381 0.10755812
0.24765233 0.10762329
0.22799738 0.10721930
0.20593541 0.10580913
0.18518853 0.10373444
0.16534368 0.10084792
0.14874617 0.09769078
0.13151723 0.09372181
0.11230381 0.08821938
0.09480426 0.08190511
0.08001082 0.07568104
0.06639004 0.06900103
0.05250233 0.06122449
0.03878766 0.05205498
0.02845287 0.04386485
0.01923802 0.03534089
0.01185536 0.02687216
0.00603546 0.01852441
0.00273905 0.01173198
0.00171214 0.00909954
0.00068523 0.00566938
0.00034262 0.00398471
0.00000000 0.00000000 Leading edge
0.00034262 -0.00299172
0.00068523 -0.00421012
0.00171214 -0.00663924
0.00273905 -0.00810491
0.00617992 -0.01112386
0.01130043 -0.01377240
0.01915776 -0.01633266
0.02885704 -0.01822030
0.03878766 -0.01913754
0.05258885 -0.01954609
0.06639004 -0.01955852
0.08022130 -0.01943995
0.09489446 -0.01917703
0.11216705 -0.01891281
0.12882187 -0.01869441
0.14547669 -0.01855429
0.16543470 -0.01851024
0.18533980 -0.01873618
0.20594016 -0.01924163
0.22799738 -0.02008124
0.24822560 -0.02100703
0.26845381 -0.02202297
0.28685302 -0.02300557
0.30574361 -0.02407739
0.32561695 -0.02525884
0.34896891 -0.02666508
0.37076206 -0.02790013
0.39109807 -0.02898579
0.41143408 -0.03001075
0.43576126 -0.03113907
0.46788063 -0.03237555
0.49702431 -0.03318960
0.52616798 -0.03374483
0.55226423 -0.03397473
0.57821723 -0.03400280
0.60312424 -0.03392726
0.62940952 -0.03368377
0.65768348 -0.03311366
0.68595745 -0.03229143
0.71198438 -0.03148645
0.76083005 -0.02863288
0.78667318 -0.02684891
0.80837777 -0.02512088
0.82399613 -0.02382386
0.85523283 -0.02090057
0.88293587 -0.01794665
0.90390685 -0.01540877
0.92487783 -0.01264436
0.94233865 -0.01007167
0.95677273 -0.00784591
0.96679021 -0.00617816
0.97552826 -0.00464475
0.98296291 -0.00328591
0.98907380 -0.00213384
0.99384417 -0.00121351
0.99726095 -0.00054347
0.99931477 -0.00013648
1.00000000 0.00000000 Trailing Edge (Ending at lower airfoil surface)
F5Fras15
| Copyright © 2000- 2011 by Richard W. Fraser All Rights Reserved |
| F3Fras13 and 15 Airfoil |
| The F5Fras13 & 15 airfoil coordinates are free for you to copy |
| ||||||||||
FRASER AERO TECHNOLOGY COMPANY |
Technical Page