United States Patent (19) 11) Patent Number: 4,519,746

Wainauski et al. (45) Date of Patent: May 28, 1985

54 AIRFOIL BLADE 4,121,787 10/1978 Wilby ................................ 244/35 R
4,123, 198 10/1978 Harbord .............................. 416/243
75 Inventors: Harry S. Wainauski, Simsbury; Carl 4,147,437 4/1979 Jonqueres ........................... 366/343
Rohrbach, Manchester, both of 4,240,597 12/1980 Ellis et al...... ... 244/35 R
Conn. 4,240,598 12/1980 Espin et al. ....................... 244/35 R
73) Assignee: United Technologies Corporation, FOREIGN PATENT DOCUMENTS
Hartford, Conn. 0647159 12/1950 United Kingdom................ 46/242
(21) Appl. No.: 286,484 Primary Examiner-Philip R. Coe
22 Filed: Jul. 24, 1981 Assistant Examiner-Christine A. Peterson
(51) Int. Cl. .............................................. B64C 11/18 Attorney, Agent, or Firm-John Swiatocha
(52) 416/223 R; 416/237 (57) ABSTRACT
(58) Field of Search ................... 416/237, 242, 223 R, An improved airfoil blade for aircraft propellers and the
416/DIG. 2 like which exhibits high aerodynamic performance and
(56) References Cited low far and near field noise levels at a minimum weight
U.S. PATENT DOCUMENTS is provided with a novel high lift and high lift to drag
2,709,052 5/1955 Berg ...................................... 244/35
cross-sectional airfoil shape characterized by a blunt,
3,494,424 2/1970 Stanley .. ... 416/223
generally parabolic leading edge portion (10) which
3,625,459 12/1971 Brown ................................... 244/35 fairs into a pressure surface (15) characterized by a
3,706,430 12/1972 Kline et al. ........................... 244/35 leading convex portion (20). For thickness ratios less
3,854,845 12/1974 Van DeWater . ... 416/228 than approximately 0.15, the pressure surface includes a
3,890,062 6/1975 Hendrix et al. ..................... 416/234 concave trailing portion (25) at one end thereof, fairing
3,915, 106 10/1975 DeWitt ................ ... 14/66.5 H into the leading convex portion (20) and at the other
3,946,688 3/1976 Gornstein et al. ... 114/66.5 H end thereof, terminating at a blunt trailing edge (35).
3,952,971 4/1976 Whitcomb ........................ 244/35 R The leading edge portion also fairs into a suction sur
4,046,489 9/1977 Fairchild et al. . ... 416/223 R
4,050,651 9/1977 Neal et al. ............................. 244/15 face (30) which is convex along substantially the entire
4,063,852 12/1977 O'Connor ..... ... 416/228 length thereof to the blade trailing edge.
4,072,282 2/1978 Fulker et al. . 244/35 R
4,120,609 10/1978 Chou et al. ......................... 416/223 3 Claims, 21 Drawing Figures

2%%%2

& 430
U.S. Patent May 28, 1985 Sheet 1 of 15 4,519,746
A/G./
T.O. LOW
Ya 2-N Speed
/C
NHs
airfoil
C
High
/ Speed
M LOW
Speed

High speed
HS airfoil

s LOW speed
U.S. Patent May 28, 1985 Sheet 2 of 15 4,519,746
U.S. Patent May 28, 1985 Sheet 3 of is 4,519,746
o

s
s :

O
CO
g g g(O St (N
g o rf) V r O
A gagwvo O/L SS3Nylot HL
U.S. Patent May 28, 1985 Sheet 4 of 15 4,519,746

A/G. 5

30

10-
(1-S
A. 25
/5

A/G. 6
U.S. Patent May 28, 1985
U.S. Patent May 28, 1985 Sheet 6 of 15 4,519,746
A/G. 9

2.5 -4.5- --Mn max = 2.88

2.0 -4.0
-3.5
5 Airfoil design condition 1 - take-off
Mn -3.0 R no. 0.55
HS1.606 NACA 16.7O6
-2.5 C1 592 592
Cod 0.0247 0.043
-2.0 Alpha 7.87 8.59

1.0 Cp -15 HS1 airfoil

Series 16 airfoil
-0.5

O.O

0.5

O
-
U.S. Patent May 28, 1985 Sheet 7 of 15 4,519,746

Airfoil design condition 2 - takeoff

Mach no. = 0.58
14 HS1.606 NACA 6-7O6
- 2.5 C .95 1,198
Cod 0.0118 O.O131
Alpha 4,639 4.750
1.2 - 2.0
Mn max F f4
M max F 25
O HS1 airfoil

Series 16 airfoil

O XIC 1.0
U.S. Patent May 28, 1985 Sheet 8 of 15 4,519,746

A/G. //

Airfoil design condition 3 - climb

Mach no. = 0.62
2.0 HS-606 NACA 16-706
2. C 0.848 0.849
Cod 0.0096 0.0101
Alpha 1829 1.760
- 1.5
Mn max F 0.958
Mn 1.0-
- 1.0 HS1 airfoil

- 0.5
Cp
0.0

0.5
Series 16 airfoil
1.0

XIC
U.S. Patent May 28, 1985 Sheet 9 of 15 4,519,746

A/G. /2 Airfoil design condition 4 - cruise

Mach no. = 0.75
HS1.606 NACA 16.7O6
C1 0.372 0.371
- 2.5 Cod 0.0114 O.O119
Alpha - 2.05 - 1.58

2.0

1.5
- 2.0

- 1.5
N Mn max = 2.2
Mn max 1.39
Mn
- O HS1 airfoil
Cp Series 16 airfoil
1-0 los
O.O

0.5

O
O XIC 1.O
U.S. Patent May 28, 1985 Sheet 10 of 15 4,519,746

O
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U.S. Patent May 28, 1985
U.S. Patent May 28, 1985 Sheet 12 of is 4,519,746

-?19S09H-

ow
8
ož?,
91.
o

10

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of7
0
o
U.S. Patent May 28, 1985 Sheet 13 of 15 4,519,746

dHs(diº)

KLH!ONBOISVH U=tIy™O"BION
HTEIOd-V?H
do “ueogeoo ueMod
U.S. Patent May 28, 1985 Sheet 14 of 15 4,519,746

WL8ONEH.|V-OE. S9|T8-IEHO dIVHS/UI=t™yOB"I0N py*aT‘ou?enu

t
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U.S. Patent May 28, 1985 Sheet 15 of 15 4,519,746
 4,519,746
1. 2
having high lift-to-drag coefficients during aircraft
AIRFOIL BILADE cruise modes of operation.
It is another object of the present invention to pro
TECHNICAL FIELD vide an airfoil blade characterized by airfoil sections
This invention relates in general to airfoil blades for 5 having high critical Mach numbers over an extensive
rotors and particularly to a family of airfoils particularly operation range.
for high-performance, low-noise aircraft propeller It is another object of the present invention to pro
blades. vide an airfoil blade characterized by airfoil shapes
consistent with known propeller manufacturing tech
BACKGROUND ART O niques.
Current and expected trends in air travel have led to It is another object of the present invention to pro
the design of a new generation of turboprop commuter vide an airfoil blade having a shape which exhibits en
aircraft which are expected to enter service in the mid hanced resistance to damage from both handling and
1980's. These are short haul aircraft, expected to service impact with foreign objects.
small airports located relatively close to populated ar 15 In accordance with the present invention, an im
eas. Accordingly, far-field noise restrictions for the proved airfoil blade for aircraft propellers and the like is
aircraft will be quite demanding. The commuter aircraft provided with novel airfoil sections along the length of
will be used extensively by travelers in the initial and the blade, the sections being characterized by rounded,
final legs of trips wherein the major portion of the dis generally parabolic leading edge portions fairing into
tance traveled will be by modern, comfortable, wide
20 pressure surfaces which, for thickness ratios less than
bodied turbofan aircraft. Accordingly, the demands on approximately 0.15, are characterized by leading, con
the commuter aircraft for safety, comfort, reliability vex protuberant portions which in turn fair into a con
and low cabin noise levels will be stringent. cave trailing portions. For thickness ratios greater than
To meet such stringent far-field and cabin noise re approximately 0.15, the trailing pressure surfaces are
strictions, propeller tip speed must be kept to a mini
25 slightly convex along their length. The leading edge
mum. However, since the new commuter aircraft are portions also fair into convex suction surfaces which
designed to operate from short runways, such low tip merge with the trailing pressure surfaces into slightly
speeds must impart high blade thrust levels (lift coeffici blunt trailing edges. The rounded leading edge portions
ents) to the propellers at low blade weights (low blade 30 at relatively high angles of attack and low Mach num
bers and protuberant leading pressure surface portions
solidities) for takeoff and climb modes of operation. at relatively low angles of attack and high Mach num
Even with a minimization of tip speed, air speed over bers tend to reduce the extent of turning of airflow over
the propellerblade surfaces is necessarily quite high. To the airfoil surfaces thereby reducing local surface Mach
avoid pronounced shock waves and the attendant flow numbers and maintaining lower pressure gradients than
separation and sacrifice in performance resulting there 35 encountered in state of the art airfoil shapes. The
from, it is necessary to maximize the critical Mach num slightly blunt trailing edge portion defines a trailing
bers associated with the aircraft propeller blade airfoil suction surface portion which exhibits a gradual pres
sections. For enhanced efficiency, high lift-to-drag co sure recovery thereby minimizing separation of flow
efficients at cruise conditions are also required. from the suction surface. Enhanced aerodynamic per
In addition to satisfying the above-noted aerody 40 formance is achieved at Mach numbers characteristic of
namic performance and noise requirements, the propel tip speeds sufficiently low to achieve enhanced far-field
ler blades must be capable of being manufactured with and cabin noise minimization.
known production techniques and should exhibit a mini
mum risk of damage from both normal handling and BRIEF DESCRIPTION OF THE DRAWINGS
impact with foreign objects. 45 FIG. 1 is a graphical representation of takeoff and
State of the art airfoil families which define propellers climb lift and drag performance of typical low and
and the like include the NACA Series 6 and Series 16 high-speed airfoils at various angles of attack;
airfoils which heretofore have exhibited adequate aero FIG. 2 is a graphical representation of the cruise
dynamic and noise performance. However, for the new performance (lift/drag) of typical high and low-speed
generation commuter aircraft noted hereinabove, the 50 airfoils at various values of lift coefficient;
performance characteristics of propellers defined by FIG. 3 is a series of cross-sectional elevations of the
such airfoil shapes are marginal at best. Newer airfoils airfoil of the present invention and a plan view of a
such as the Lieback, Wortmann, Whitcomb supercriti blade showing exemplary locations, along the blade axis
cal, and GAW airfoils have been designed for special of those sections, various of the airfoils being enlarged
wing configurations and as such, are not suitable for 55 to show details of the shape thereof;
general propeller use in that for the most part, these FIG. 4 is a graphical representation of the camber
airfoils incorporate shapes undesirable for propeller lines and thicknesses of a family of airfoil shapes within
manufacture from structural and fabrication stand which the airfoils of FIG. 3 are included;
points. FIG. 5 is a cross-sectional elevation of one of the
DISCLOSURE OF INVENTION
60 airfoils illustrated in FIG. 3;
FIG. 6 is a cross-sectional elevation of a prior art
Accordingly, it is an object of the present invention NACA Series 16 airfoil;
to provide an improved airfoil blade characterized by FIG. 7 is an elevation of the NACA airfoil of FIG. 5
airfoil sections having associated therewith, high lift in takeoff, climb, and cruise modes of operation;
coefficients, especially in aircraft takeoff and climb 65 FIG. 8 is an elevation of one of the airfoils of FIG. 3
modes of operation. in takeoff, climb, and cruise modes of operation;
It is another object of the present invention to pro FIGS. 9, 10, 11 and 12 are plots of pressure coeffici
vide an airfoil blade characterized by airfoil sections ent and Mach number along the pressure and suction
 4,519,746
3 4.
surfaces of one of the airfoil sections of the blade of the performance characteristics almost equal to the classic
present invention and a corresponding NACA Series 16 low speed airfoil and the cruise characteristics of the
airfoil; high speed airfoil, all in a single airfoil of a novel cross
FIGS. 13 and 14 are plots of lift and drag coefficients sectional airfoil shape as shown in FIG. 3.
respectively, for one of the airfoil sections of the blade Referring to FIG. 3, a series of cross sections of the
of the present invention for various angles of attack; airfoil blade of the present invention are shown. Each
FIGS. 15 and 16 are plots of lift and drag coefficients cross section is identified by indicia comprising three
respectively, similar to those of FIGS. 13 and 14 for a numerals setting forth the design lift coefficient multi
prior art corresponding NACA Series 16 airfoil; plied by 10 (first numeral) and the thickness coefficient
FIGS. 17 and 18 are graphical representations of lift multiplied by 100 (the last two numerals). Thus, for
and drag coefficients and lift-to-drag ratios respectively example, the uppermost airfoil section, is characterized
for one of the airfoil sections of the blade of the present by a design lift coefficient of 0.4 and a thickness ratio of
invention and a corresponding NACA Series 16 airfoil; 0.04, the second airfoil section having a design lift coef
FIGS. 19 and 20 are performance maps of efficiency ficient of 0.6 and a thickness ratio of 0.06, the third
and power coefficient at high Mach numbers plotted 15 airfoil section having a design lift coefficient of 0.7 and
against advance ratio for a propeller having the airfoil a thickness ratio of 0.08, the fourth airfoil section having
sections of the present invention and a propeller having a design lift coefficient of 0.7 and a thickness ratio of
NACA Series 16 airfoil sections respectively; and 0.12, the fifth airfoil section having a design lift coeffici
FIG. 21 is a plot of efficiency versus advance ratio at ent of 0.6 and a thickness ratio of 0.20 and the sixth
lower Mach numbers for the propellers whose perfor 20 airfoil having a design lift coefficient of 0.4 and a thick
mance maps are illustrated in FIGS. 19 and 20. ness ratio of 0.30. Still referring to FIG. 3, illustrative
BEST MODE OF CARRYING OUT THE locations on a single propeller blade of the airfoil sec
INVENTION tions are shown. It is seen that the 404 airfoil section is
In general, propeller blade section thrust is character 25
taken substantially at the tip of the blade, the 430 section
ized by the expression: proximal to the blade root and the 620 section is taken at
a location approximately 0.175 of the length of the blade
Toc ClbV2 longitudinal axis from the root portion thereof. The
remaining sections are taken at approximately 0.425 of
Wherein: 30 the length of the axis from the root thereof, 0.625 of the
T is thrust, length of the axis and 0.825 the length of the axis. It will
CL is lift coefficient, of course be understood that while the chords of the
b is chord length of the section, airfoil sections are illustrated as being of a common
V is section relative velocity. length, design considerations regarding blade taper will
An examination of this expression indicates that as dictate the relative sizes of the airfoil sections and the
chord b is reduced for minimum weight, and section 35 present invention shall not be limited to any specific size
relative velocity V is reduced for low noise, section lift relationship between the airfoil sections.
coefficient CL must be increased to maintain a given Those cross sections of the blade between the airfoil
thrust. Accordingly, it is apparent that the lift coeffici sections shown in FIG. 3 are defined by a transition
ent must be maximized to achieve a given section thrust 40 surface connecting corresponding portions of any two
output when chord and section relative velocity are adjacent airfoil shapes. The airfoil cross sections will, of
reduced for minimization of weight and noise. At the course, be angularly displaced from one another in a
same time, it will be appreciated that for cruise modes of manner well known in the art to impart sufficient twist
operation at low operating lift coefficients and high to the blade to establish varying blade angles of attack
section Mach numbers, the airfoil sections must be char 45 dictated by aerodynamic performance requirements.
acterized by high lift-to-drag ratios. The following tables list precise dimensionless coor
Heretofore, it has been extremely difficult to achieve dinates of a number of airfoil sections of the blade of the
high aerodynamic performance at both takeoff and present invention wherein the x/c values are dimension
cruise conditions with an airfoil blade having cross-sec less locations on the blade chord line, y/c upper are the
tional shapes from an existing airfoil family. Referring dimensionless heights from the chord line of points on
to FIG. 1, the shaded regions of the curve are indicative 50
of the performance output capabilities of a typical low the suction surface and y/c lower are the dimensionless
speed airfoil and a typical high speed airfoil at lift coeffi heights from the chord line of points on the pressure
surface.
cients representative of propeller takeoff (T.O.) and
climb conditions. It is seen that a classic "low speed' TABLE I
airfoil exhibits a much greater lift coefficient and sub 55
HS-404
stantially less drag for takeoff conditions than a classic (y/c) (y/c) (y/c) (y/c)
"high speed' airfoil and would therefore, be more desir xac upper lower x/c upper lower
able than the high speed airfoil. However, referring to 0.00000 000009 0.00004 0.44000 0.04914 0.01.16
FIG. 2, wherein the shaded region is indicative of the 0.00050 0.007 -0.00113 0.46000 0.04893 0.01.129
performance output capabilities of the same two airfoils 60 0.0000 000250
0.00200 000372
-0.00155
-0.00210.
0.48000
0.50000
0.04860
0.04818
0.01137
0.01139
at cruise modes of operation, it is seen that the high 0.00300 000470 -0.00253 0.54000 0.04705 0.01132
speed airfoil is much more desirable than the low speed 0.00500 0.00628 -0.00323 0.56000 004635 0.01122
airfoil, since it exhibits substantially higher lift to drag 0.00750 000788 -0.00395 0.58000 0.04555 0.01.08
ratios at lift coefficients corresponding to normal cruise 0.01000 000923 -0.00455 0.60000 0.04461 0.0109
conditions. In FIGS. 1 and 2, the curves indicated by 65 0.02000 0.01339 -0.00620 0.64000 0.04223 0.01043
0.03000 0.01655 -0.00709 0.66000 0.04.076 0.01010
the dotted line exemplify the performance of the airfoil 0.04.000 0.0922 -0.00751 0.68000 0.0391 0.0097.
HS1 of the present invention. As is readily noted from 0.05000 002158 -0.00761 0.70000 0.03728 0.00924
these curves, this airfoil exhibits the takeoff and climb 0.06000 0.02375 -0.00756 0.74000 0.0339 0.00809
 4,519,746
5 6
TABLE I-continued TABLE III-continued
HS1-404 HS-708
(y/c) (y/c) (y/c) (y/c)
x/c upper lower x/c upper lower xAc upper lower x/c upper lower
0.07000 0.02S77 -0.0024 0,76000 0.03097 0.00741 0.08000 0.0503 -0.01588 0.78OOO 0.05062 0.00650
0.08000 0.02766 -0.00686 0,78000 0.02867 0.00668 0.09000 0.05344 -0.01530 0.80000 004649 0.00557
0.09000 0.02943 -0.00639 O.80000 0.02631 0.00591 0.10000 0.05634 -0.01456 0.82000 0.0422 0.00459
0,10000 0.0307 -0.00584 0.82000 0.02389 0.0052 0.12000 0.06151 -0.0127 0.84.000 003798 0.00356
0.2000 0.03403 -0.0045 0.84.000 0.0243 0.00429 0.4000 0.06595 -0,01050 086000 0.03360 0.00248
0.14000 0.03659 -0.00295 0.86OOO 0.0892 0.00343 10 0,6000 0.06977 -0.00808 0.88000 0,02916 0.0033
0.16000 0.03883 -0.00123 0.88000 0.01637 0.00252 0.8000 0.0308 -0.00SS9 0.90000 0,024.62 0.00006
0.18000 O,04077 0.00055 0.90000 0.01376 0.0056 0.20000 0.07687 --00038 0.91000 0,022.31 -0.00058
0.2OOOO 0.04244 0.00230 0.91000 0.01243 0.0006 0.22000 0.07824 -0.00092 0.92OOO 0.01996 -0.00128
0.22000 0.04386 0.00393 0.92000 0.01.108 0.00054 0.24000 008021. -O.OO113 0.93OOO 0.0757 --0,001.96
0.24000 0.04504 0.00536 0.93000 0.00970 0.00000 O.26000' 0.0882 -0.00293 094000 0.052 -0.00268
0.26000 0.04601 0.00658 0.94000 OOO829 --0.00054
0.28000 0.04680 0.00758 0.95000 0.00684 -0.00110 15 0.28000
0.30000
0.0833
0,0842
-0.00450
0.0.0833
0.95000
0.96000
00262
001005
--0.00341
0.0046
0.30000 0.0445 0.00840 0.95000 000536 0.0067 0.34000 0.08583 0.0079 0.97000 000741 -0.00493
0.34,000 0.04844 O.OO961 0.97000 OOO384 -0.00225 0.36000 0.08640 0.00868 0.98000 000471 -000572
0.36000 0.04880 0.0007 0.98000 000229 --000285 0.38000 0.08680 0.00932 0.99000 000196 -0.00652
0.38000 O,04906 0.0044 0.99000 00000 -0.00346 0.40000 0.08O 0.00984 100000 .0,00086 -0.0034
O.4OOOO 0.04920 0.004 OOOOO -0.00092 -OOO408
20
TABLE IV
TABLE II
HS-712
HS1-606 (y/c) (y/c) (y/c) (yac)
x/c upper lower x^c upper lower
xAc upper lower xAc upper lower 25
0.00000 00000 0.00000 0.44000 0.10556 -0.01.105
0.00000 000015 0.00015 0.44000 007100 0.0387 0.00050 0.00430 -0.00377 0.46000 0.0600 --0.01060
O.00050 000259 -0.00620.46000 0.07O69 0.04.06 0.0000 0.00600 -0.00519 0.48000 0.0526 - 0.0020
OOOOO 0.00379 -0.00220
0.48000 O,0702 0.04.19 0.00200 0.00903 -0.00709 0.50000 0.10423 -0.00985
0.00200 0.00562 -0.00296
0.50000 OO6959 0.0425 0.00300 002 -0.00342 0.54000 0.1054 -0.00919
O.O.O.300 000708 -0.00354
0.54000 OO6796 0.0420 30 0.00500 0.0490 -0.0044 0.56000 0,09965 0.00888
OOO500 000938 -0.00455
0.56000 OO6695 0,,O1409 0.00750 0.01861 -0.0230 0.58000 0,09795 0.00856
0.00750 001164 -0.00563
0.58000 0.06578 0.01394 0.0000 0,0280 -0.0376 0.60000 0.09587 -0.00824
OOOOO 0.01352 -0.00656
0.60000 0.06442 0.01373 0.02000 0.0386 -0.01767 0.64000 0,09125 -0.00756
0.02000 0.01926 -0.00925
0.64000 OO6098 0.034 0.03000 0.03967 -0.02006 0.66000 0.08872 -000721
0.03000 0.02371 -0.01079
0.66000 0.05886 0.0273 0,04000 0.04688 -0.02109 0.68000 0.08605 -0.00687
0.04.000 0.02756 -0.01600.68000 0.05647 0.01224 0,05000 0.05192 -0.02283 0.70000 0.08322 -0.00658
0.05000 0.0305 -0.090 O.70000 0.05383 0.065 35
0.06000 0.06698 -0.02353 0.74000 0.07789 -0.00619
0.06000 0.03429 -0.018 0.4000 0.04794 0.01.018 0.07OOO 0.00651 -0.0247 0.76OOO 0.07677 -0.0061
0.07000 003731 -0.016 0.76000 0.0444 0,00930 0.08000 0.06562 -0.02450 0.78000 0.07028 --000809
0.08000 0.04.012 -0.018 0.78000 0.0442 0.00835 0.09000 OO6937 -0.02466 0.80000 0.06668 -0.0061
0.09000 0.04273 -0.01060 O.80000 0.03802 0.0034 0.10000 0.07280 -0.02469 0.82000 0.06262 -0.00614
0.0000 0.04515 .000988 0.82000 0.03453 0.00530 0.12000 0.07886 -0.02440 0.84.000 0.05834 -0.00621
0.2000 0.04946 .000807 0.84.000 0.03099 0.00522 40 0.14000 0.0890 -0.02374 0.86000 0.08368 --0.00633
0.4000 0.05318 -0.00588 0.86000 0.02738 0.00408 O,6000 0.08839 -0.02279 0.88000 0.04856 -0.00653
0,16000 0.05638 -0.00347 0.88000 00237 0.00288 O.8000 OO921 -0.0259 0.90000 0.04299 -0.00683
0.18000 0.0595 -0.00099 0.90000 0.01996 0.0061 0.20000 009625 -0.02030 09000 0.03997 -0.00701
0.2OOOO 0.06153 0.0043 0.91000 0.01804 0.0009.4 0.22000 0,09786 -0.090. 0.92000 0.03686 --000721
0.22000 0.06354 0.00367 0.92000 0.016.6 0.00025 0.24000 001000 -0.0786 0.93000 0.03364 --000742
0.24000 0.06522 0.00565 0,93000 00141 - 0.00046 45 0.26OOO 0.10173 -0.01688 0.94000 0.03032 -0.00753
0.26000 0.06659 0.00734 0.94000 001209 -0.00119 0.28000 0.10313 -0.01607 0.95000 0.02889 -0.00785
0.28000 0.06772 0.00875 0.95000 0,000 --000193 0.30000 0.0428 -0.01536 0.96000 0.02335 -0.00807
0.30000 0.06865 0.00992 0.96000 0.00788 --000289 0.34000 0.0542 -0.01489 0.97000 0.01968 0.00829
0.34000 0.07005 0.01.167 0.97000 000570 -0.00346 0.36000 0.0632 -0.01344 0.98000 0.01970 -0.00882
0.36000 0.07.056 0,0231 0,98000 0.00346 --00042.5 0.38000 0.0587 -0.01279 0.99000 0.01971 -0.00576
0.38000 0.07092 0.01284 199000 0.00118 --000508 SO 0.40000 0.10700 -0.01216 1.00000 000720 -OOOOO
0.40000 0.07112 0.0326 100000 -0.00115 -0.00589

TABLE V
TABLE III HS1-620
HS1-708 55 (y/c) (y/c) (y/c) (y/c)
xAc upper lower x/c upper lower
XAc upper lower x/c upper lower 0.00000 000001 0.0000 0.44000 0.13904 -0.05705
0.00000 0.00013 0.00013 0.44000 0.0868 0.0060 0.00050 000694 -0.00650 0.46000 0,384 -0.05626
0.00050 000330 -0.00244 0.46000 0.08640 0.01087 0.00100 0.00993 -0.00907 0.48000 0.13699 -0.05545
0.00100 000479 -0.00338 0.48000 0.08579 0.01.06 0.00200 0.0424 -0.01258 0.50000 0.13556 -0.05459
0.00200 000704. .000485 0.50000 0.08502 0.020 60 0.003OO 0.01760 -0.01520 0.54000 0.1389 -0.05269
0.00300 000883 -0.0056 0.54000 008301 0.029 0.005OO 0.02300 -0.01922 0.56OOO 0.2963 -0.0516
O.00500 0.01173 -000715 0.56000 00876 0.01.125 0.00750 002845 -0.02308 0.58000 0.271 -0.05045
0.00750 001461 -0.00865 0.58000 008033 0.0118 0.0000 0.03306 -0.0262 0,60000 0.12435 0.0499
OOOOO 0.01704 -0.00990 0.60000 0.07865 0.01.105 0.02000 0.04.737 -0.0352 0.64000 0.1823 -0.04645
0.02000 0.02448 -0.01322 0.64000 0.07443 0.01065 0.03000 0.05820 -0.0439 0.66000 0.1489 -0.04499
0.03000 0.03017 -0.01506 0.66000 0.0784 0.0034 65 O,04000 0.06713 -0.04609 0.68000 0.1138 --0.04348
0.04.000 0.0350 -0.01603 0.68000 0.06892 0.00996 0.05000 0.07477 -0.04983 0.70000 0.0770 -0.0497
0.05000 0.03932 -0.01645 0.70000 OO6570 0.00947 0.06000 0,0814 -0.05289 0.74000 0.0998 -0.03900
0,06000 0.04326 -0.01650 0.74000 0.05853 0.00818 0.0000 0.08742 -0.0554 0.76000 0.09560 -003755
0.07000 0.04692 -0.01629 0.76000 0.05465 0.00738 0.08000 0.09274 -0.05751 0.78000 0.0918 -0.03610
 4,519,746
7 8
TABLE V-continued maximum camber at approximately 0.5 x/c. The 0.08
HS1-620 thickness ratio airfoil has a maximum thickness at ap
(y/c) (y/c) (y/c) (y/c) proximately 0.33 x/c and a maximum camber at approx
x/c upper lower x/c upper lower imately 0.50 x/c. The 0.12 thickness ratio airfoil has a
0.09000 0.09755 -0,05925 O.80000 0.08651 -0.03462 maximum thickness at approximately 0.32 X/c and a
0.10000 0,10190 -0.06070 0.82000 0.08153 -0.03307 maximum camber at approximately 0.38 x/c. The 0.20
0.12000 0.10943 -0.06280 0.84000 0.07616 -0.03144 thickness ratio airfoil has a maximum thickness at ap
0.4000
0.16000
0.11566
0.12079
-0.06404
-0.06459
0.86000
0.88000
0.07031
0.06391
- 0.02972
-0.02790
proximately 0.315 x/c and a maximum camber at 0.30
0.18000 0.12501 -0.06463 0.90000 0.05693 -0.02600 10
X/c. The 0.30 thickness ratio airfoil has a maximum
0.20000 0.2844 - 0.06430 0,9000 0.05322 -0.02502 thickness at approximately 0.310 x/c and a maximum
0.22000 0.3121 - 0.06376 O. 92000 0.04939 -0.02401 camber at approximately 0.29 x/c. Each of the airfoils is
0,24000
0.26000
0.13344
0,13524
-0.06317 0.93000
-0.06261 0.94000
0.04545
0.04142
-0,02297
-0.0289
further characterized by a trailing edge thickness equal
0.28000 0.13670 -0.062 11 0.95000 003727 -0.02078 to approximately 10% of the maximum section thick
0.30000 0.13789 -0.06163 0.96000 0.03295 -0.0963 eSS.
0.34000 0.3955 -0.06061 0.97000 0.02840 -001843 15 From the foregoing and referring to FIG. 5 which
0.36000 0.14000 -0.06000 0.98000 0.02353 -0.01716 illustrates the 708 airfoil, it is seen that the airfoils of the
0.38000 0.14017 -0.05932 0.99.000 0.01831 -0.01580 present invention are characterized along substantially
0.40000 0.4007 -0.05859 .00000 0.0272 -0.0427
the entire length of the chord thereof by cross-sectional
airfoil shapes, each comprising a rounded generally
TABLE VI 20 parabolic leading edge portion 10 fairing into a pressure
HS-430
surface 15 having a leading convex portion 20 which
(y/c) (y/c) (y/c) (y/c) fairs into a trailing portion 25. The leading portion also
x/c upper lower x/c upper lower fairs into a convex suction surface 30, the pressure and
0.00000 000000 0.00001 0.44000 0.17626 0.1742 25 suction surfaces merging into a slightly blunt trailing
0.00050 0.01024 -0.0099 0.46000 0.17497 -0.11622 edge 35. As shown in FIG. 3, for thickness ratios less
0.0000 0.01456 -0.01393 0.48000 0.17342 -0.11489 than approximately 0.15 the trailing portion 25 of the
0.00200
0.00300
0.0207
0.02547
-0.01951
-0.02372
0.50000
0.54000
0.1756
0.6680
-0.11340
-0.10987
pressure surface is concave in shape thus defining the
0.00500 0.03303 -0.03028 0.56000 0.16388 - 0.0780 leading portion of the pressure surface as being of a
0.00750 0.04057 -0.03668 0.58000 0.16062 -0.10553 . protuberant nature. At such thickness ratios, the convex
0.01000 0.04692 -0.041.96 0.60000 0.15707 -0.10307 30 protuberant portion fairs into the concave trailing por
0.02000
0.03000
0.06627
008068
-0.05751
-0.06858
0.64000
0.66000
0.4920
0.4493
-0,09765
-0.09472
tion at a distance from the airfoil leading edge of ap
0.04000 0.0924 -0,07728 0.68000 0.14045 -0.0968 proximately 10-15% of the airfoil chord length. At
0.05000 0.01.235 -0,08441 0.70000 0.13577 -0.08857 thickness ratios greater than 0.15 trailing portion 25 is
0.06000 0.11097 -0.09041 0.74000 0.12588 -0.08220 COVeX.
0.07000
0.08000
0,11857
0.12534
-0.09554
-0.09995
0.76000
0.78000
0.2064
0.1518
-0.07895
-0.07562
35 It is also noted that the relatively rounded leading
0.09000 0.13138 -0.10376 O.80000 0.10942 -0.07216 edge of the airfoil of the present invention minimizes the
0.10000 0.13680 -0,10707 0.82000 0.10329 --006850 risk of damage due to normal handling and impact with
0.12000 0.14595 -0.11231 0.84.000 0.09670 -0.06459 foreign objects.
0.14000 0.15325
0.16000 0.15904
- 0.11605 0.86000
-0.1859 0.88000
0.08955
0,08178
-0.06040
-0.05590 40
FIG. 6 is illustrative of the general shape of an
0.18800 0.16517 -0.12064 0.90000 0.07334 -0.0509 NACA Series 16 airfoil, a shape used widely in propel
0.20000 0.16723 -0.12114 0.91000 0.06385 -0.04855 lerblades of present day turbo-prop powered commuter
0.22000 0.701 -0.12164 0.92000 0.06420 -0.04594 aircraft. It is seen that the shapes of the airfoils of the
0.24000
0.26000
0.17243
0.17433
-0.12190
-0.12202
O.93000
0.94000
0.05938
0.05440
-O,04323
-0.04.043
present invention are readily distinguishable from the
0.28000 0.17587 -0.12204 0.95000 0.04927 -0.03754 45 shape of the Series 16 airfoil. It is first noted that the
0.30000 0.17708 -0.12197 0.96000 0.04396 --O.03455 Series 16 airfoil includes a concave pressure surface
0.34000 0.1785 -0.1245 0.97000 0.03846 -0.03145 along the entire blade chord while the airfoil shapes of
0.36000
0.38000
0,17872
0.17857
-0.12096
-0.2030
0.98000
0.99000
0.03267
0.02649
-0.0288
-0.02467
the blade of the present invention include the convex
0.40000 0.17808 - 0.1948 1.00000 0.0969 -0.02080 portion extending at least along the leading 10-15% of
the blade pressure surface. It is also seen that the NACA
5O Series 16 airfoil includes a relatively sharp leading edge
FIG. 4 is a graphical representation of the camber portion while the airfoil shapes of the present invention,
and thickness lines of various airfoil sections of the especially those shapes of a thickness ratio greater than
present invention, x/c being indicative of dimensionless 0.06 include rounded leading edge portions and rela
locations on the chord line, y/c being indicative of the tively blunt trailing edges for higher critical Mach num
dimensionless height of the camber line from the airfoil 55 bers at leading portions thereof and enhanced pressure
chord and t/c being the total dimensionless thickness of recovery at the trailing portions thereof.
the airfoil at the corresponding chord location. h/b The enhanced performance of the airfoil of the pres
indicates the thickness ratios of the various airfoil sec ent invention (HS1) is compared with the performance
tions. of the conventional NACA Series 16 airfoil in FIGS. 7
Each airfoil section has a unique location of maxi 60 and 8. As shown in FIG. 7, on takeoff, at high angles of
mum thickness and camber and when these airfoils are attack, the sharp nose of the Series 16 airfoil produces a
incorporated in a single propellerblade, smooth, contin shock at the leading edge which causes extensive sepa
uous upper and lower surfaces result. ration of the boundary layer along the suction surface of
The airfoils are characterized as follows. The 0.04 the blade, the airfoil of the present invention having a
thickness ratio airfoil has a maximum thickness at ap 65 much rounder leading edge does not establish such high
proximately 0.35 x/c and a maximum camber at approx local Mach numbers and therefore maintains a gentler
imately 0.50 x/c. The 0.06 thickness ratio airfoil has a pressure gradient which permits the boundary layer to
maximum thickness at approximately 0.34 x/c and a remain attached throughout most normal angles of at
 4,519,746
9 10
tack. It is noted from FIGS. 7 and 8 that a climb condi cruise and a 40-60% higher lift/drag ratio in climb than
tions, both airfoils perform reasonably well, each oper the Series 16 airfoil.
ating at its design lift coefficient. However, in cruise Model propellers, one comprising four blades with
modes of operation, the airfoils are operating at low lift airfoils of the present invention and the other incorpo
coefficients and high Mach numbers. In this case, the 5 rating four NACA Series 16 blades were tested in the
highly cambered NACA Series 16 airfoil operates "nose United Technologies Subsonic Wind Tunnel Test Facil
down' to the relative air velocity, the sharp nose of this ity in East Hartford, Conn. Both models were 3.25 ft. in
airfoil producing a leading edge shock on the pressure diameter. Other than airfoil shape, and a slight differ
side of the airfoil causing boundary layer thickening or ence in camber level, both models were geometrically
separation over the forward portion of the airfoil, 10 identical, being of solid aluminum and incorporating the
thereby adversely affecting effficiency (lift/drag) ratios. same planform, thickness ratio and twist distribution,
On the other hand, the more rounded leading edge with a 91 activity factor. The integrated design lift
portion and the protuberant leading pressure surface coefficients of the Series 16 blades were adjusted
portion of the airfoil of the present invention results in slightly to compensate for the higher effective camber
lower Mach numbers under cruise conditions with no 15 levels of the airfoils of the present invention. The model
strong shock waves and therefore no associated for propellers were tested in both 8 and 18-ft. throats of the
ward boundary layer thickening or separation. As the above-noted wind tunnel. Testing in both throats per
data which follow indicates, the efficiency of a propel mitted data to be gathered for conditions ranging from
ler defined by the airfoil blades of the present invention 0.03 Mach numbers up to and including Mach numbers
represents at least a 2-4% increase in takeoff efficiency of 0.6 and blade angles of from -20° to +81 with
and a 1-2% increase in cruise efficiency. propeller speeds in the range of normal operation.
Referring to FIGS. 9 through 12, each of the Figs. FIGS. 19 and 20 are representative of portions of the
illustrates the variation in pressure coefficient, Calong data obtained from this wind tunnel test and clearly
the airfoil chord, x/c for both a representative airfoil show that at cruise Mach numbers of 0.4 the perfor
shape of the present invention and a close correspond 25 mance of the airfoil blade of the present invention is
ing NACA Series 16 blade. It should be noted that as substantially better than that of the Series 16 blade as
shown in FIG. 9, the Series 16 airfoil develops a very evidenced by the exhibited breadth of high efficiency
large leading edge Mach spike when operated in a take region on the maps.
off condition at high positive angles of attack (FIG. 9) Referring to FIG. 21, a comparison of the two pro
and high negative angles of attack required in a cruise 30 pellers at low Mach numbers (up to 0.10) is shown. This
(FIG. 12) mode of operation due to the relatively sharp plot represents a blend of the efficiency maps derived
leading edge of this airfoil. Experience has shown that a from the wind tunnel test data at Mach numbers of 0.03
surface Mach number in excess of 1.3-1.4 most often to 0.10. A study of this plot shows that as power coeffi
results in a strong shock wave resulting in boundary cient is increased the airfoil blades of the present inven
layer separation and inefficient performance. Thus, in 35 tion become progressively more efficient than the prior
FIGS. 9 and 12, the 2.88 and 2.2 Mach numbers experi art Series 16 propeller blades. For example, it is noted
enced by the Series 16 airfoil most likely result in sepa that at a power coefficient of 0.10 the airfoil blades of
ration and the poor performance resulting therefrom. the present invention exhibit an improvement of 1% in
On the other hand, it is noted that the peak surface efficiency over the Series 16 airfoil while at a power
Mach numbers for the airfoils of the present invention 40 coefficient of 0.26 and a Mach number in the 0.06 to
are much lower, only the surface Mach number at take 0.10 range, the blade of the present invention exhibits a
off exceeding the desired 1.3 to 1.4 Mach number range. 6% improvement in efficiency.
However, the gradual pressure recovery over the aft On the basis of these and various other test data, the
portion of the airfoil suction surface indicated by the propeller of the present invention was shown to be
upper righthand portions of the curves in FIGS. 9-12 45 more efficient than the Series 16 propeller over a wide
indicates that flow separation is probably minimized range of Mach numbers, advance ratios and power
despite the surface Mach number of 1.76 during takeoff coefficients, representative of commuter aircraft pro
conditions for the HS1 airfoil. peller operating conditions.
FIGS. 13 and 14 are graphical representations of Having thus described the invention, what is claimed
wind tunnel test data illustrating the relationship be 50 1S
tween lift and drag coefficients for various Mach num 1. An airfoil blade having along substantially the
bers and angles of attack for the 606 airfoil of the blade entire length thereof, a cross-sectional airfoil shape
of the present invention. As indicated in FIG. 13, since comprising a rounded, generally parabolic leading edge
there are no abrupt losses in lift or step increases in drag portion fairing into a pressure surface; said pressure
near maximum lift there is no indication that any shock 55 surface having a leading convex portion fairing into a
wave induced separation occurs on the 606 airfoil of the trailing portion; said leading edge portion also fairing
present invention despite the relatively high local mach into a convex suction surface; said pressure and suction
number at takeoff conditions. FIGS. 15 and 16 are surfaces merging into a blunt trailing edge; one of said
graphical representations of wind tunnel test data illus airfoil shapes having a thickness ratio of 0.04, a maxi
trating similar relationships of lift and drag coefficients 60 mum thickness at approximately 0.35 x/c and a maxi
to angles of attack for the Series 16 airfoil the pressure mum camber at approximately 0.5 x/c; another of said
coefficients of which are plotted in FIGS. 9-12. airfoil shapes having a thickness ratio of 0.06, a maxi
FIGS. 17 and 18 represent a comparison of data re mum thickness at approximately 0.34 x/c and a maxi
duced from FIGS. 13–16. FIG. 17 clearly shows that at mum camber at approximately 0.50 x/c; another of said
a takeoff condition, the airfoil of the present invention 65 airfoil shapes having a thickness ratio of 0.08, a maxi
exhibits a 20% increase in maximum lift coefficient mum thickness at approximately 0.33 x/c and a maxi
while FIG. 18 indicates that the airfoil of the present mum camber at approximately 0.50 x/c; another of said
invention exhibits 60-70% higher lift-to-drag ratio in airfoil shapes having a thickness ratio of 0.12, a maxi
 4,519,746
12
mum thickness at approximately 0.32 x/c and a maxi portion of a thickness equal to approximately 10% of
mum camber at approximately 0.38 x/c; another of said the maximum shape thickness.
2. The airfoil blade of claim 1 wherein for thickness
airfoil shapes having a thickness ratio of 0.20, a maxi ratios less than approximately 0.15, said trailing portion
mum thickness at approximately 0.315x/c and a maxi 5 of said pressure surface is concave in shape and said
mum camber at approximately 0.30 x/c and another of convex leading suction surface portion is protuberant
said airfoil shapes having a thickness ratio of 0.30, a from said pressure surface.
maximum thickness at approximately 0.310 x/c and a 3. The airfoil blade of claim 1 wherein said leading
maximum camber at approximately 0.29 x/c wherein convex portion fairs into said trailing portion at a dis
10 tance from the airfoil leading edge of approximately
x/c is the dimensionless chord length of said airfoil 10-15% of the airfoil chord.
shapes; all of said airfoil shapes having a trailing edge ck k k k -k

15

20

25

35

40

45

50

55

60

65