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)
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