U.S. patent number RE34,109 [Application Number 07/724,597] was granted by the patent office on 1992-10-20 for propeller blade.
This patent grant is currently assigned to IMC Magnetics Corp.. Invention is credited to Robert J. Gornstein, Dennis J. Patrick.
United States Patent |
RE34,109 |
Gornstein , et al. |
October 20, 1992 |
Propeller blade
Abstract
A propeller having blades, designed to achieve maximum thrust at
lower static velocities in both forward and reverse pitch modes is
disclosed. Each blade has a generous leading edge radius leading to
an upper surface having positive camber and a lower surface having
positive camber adjacent the leading edge and negative camber
adjacent the trailing edge. Each blade also has a high
thickness-to-chord ratio and lower span-to-chord ratio or aspect
ratio. This thicker blade profile and aft loading leads to a more
uniform pressure distribution and laminar airflow across the blade
surface resulting in a higher coefficient of lift, increased blade
life and reliability, and lower noise levels.
Inventors: |
Gornstein; Robert J. (Dockton,
WA), Patrick; Dennis J. (Seattle, WA) |
Assignee: |
IMC Magnetics Corp. (Jericho,
NY)
|
Family
ID: |
27111008 |
Appl.
No.: |
07/724,597 |
Filed: |
July 1, 1991 |
Related U.S. Patent Documents
|
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
Reissue of: |
875292 |
Jun 17, 1986 |
04844698 |
Jul 4, 1989 |
|
|
Current U.S.
Class: |
416/223R;
416/242 |
Current CPC
Class: |
B64C
11/18 (20130101) |
Current International
Class: |
B64C
11/00 (20060101); B64C 11/18 (20060101); B64C
003/26 () |
Field of
Search: |
;416/223R,223A,232,242,243,DIG.2,237 ;244/35R,35A ;415/119 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
|
|
|
|
|
|
|
0103478 |
|
Mar 1984 |
|
EP |
|
329903 |
|
May 1930 |
|
GB |
|
1079606 |
|
Aug 1967 |
|
GB |
|
Primary Examiner: Denion; Thomas E.
Attorney, Agent or Firm: Christensen, O'Connor, Johnson
& Kindness
Claims
The embodiments of the invention in which an exclusive property or
privilege is claimed are defined as follows:
1. A propeller comprising a plurality of blades, each of said
blades having a root, a tip, a leading edge, a trailing edge, means
defining an upper surface that is contoured and constructed
adjacent said tip to achieve a substantially uniform negative
pressure profile chordwise across said upper surface, means
defining an upper surface that is contoured and constructed
adjacent said root to achieve a negative pressure profile that has
a broad, rounding peak preceding the leading edge and that
decreases toward the trailing edge point on said upper surface
means, and said means defining an upper surface and means defining
a lower surface further being constructed and oriented to provide a
twist to said propeller that defines a substantially uniform pitch
spanwise from said root to said tip, said means defining an upper
surface and said means defining a lower surface being further
constructed and contoured to achieve a positive lift coefficient
.[.greater than.]. .Iadd.of .Iaddend.about 1.7 .Iadd.maximum
.Iaddend.at said root and .[.greater than.]. .Iadd.of
.Iaddend.about 1.0 .Iadd.maximum .Iaddend.at said tip, said lift
coefficient varying nonlinearly spanwise when each of said blades
is in a forward pitch mode, and to achieve a negative lift
coefficient greater than about -1.0 from said root to said tip,
said negative lift coefficent varying nonlinearly spanwise when
each of said blades is in a reverse pitch mode.
2. The propeller of claim 1, wherein said means defining an upper
surface is further contoured and constructed to have an airfoil
cross section with a positive upper camber throughout the chord
length of each of said blades.
3. The propeller of claim 2, wherein said means defining a lower
surface further is constructed and contoured to have both positive
lower camber adjacent the leading edge of said blade and negative
lower camber adjacent the trailing edge of said blade.
4. The propeller of claim 3, wherein said means defining a lower
surface being further constructed and contoured to have a positive
lower camber occurring within the leading 65 percent of the chord
length measured from said leading edge, and a negative lower camber
occurring within the trailing 35 percent of the chord length of
each of said blades.
5. The propeller of claim 4, further comprising a hub, each of said
blades being rotatably mounted on said hub for rotation about a
radial axis.
6. The propeller of claim 1, wherein each of said blades is
constructed and contoured to achieve a planform profile having an
aspect ratio between about 2.0 and about 2.2.
7. The propeller of claim 6, wherein said blades are further
constructed and contoured to achieve a planform profile having a
thickness-to-chord ratio varying nonlinearly spanwise between about
.[.21.]. .Iadd.23 .Iaddend.percent at said root and about .[.10.].
.Iadd.9 .Iaddend.percent at said tip.
8. A propeller comprised of a plurality of blades, each of said
blades having a root, a tip, a leading edge, a trailing edge, means
defining an upper surface, and means defining a lower surface being
constructed and oriented to provide a twist to said propeller,
which defines a substantially uniform pitch spanwise from said root
to said tip, and said means defining an upper surface that is
contoured and constructed adjacent said tip to achieve a negative
pressure profile chordwise across said upper surface substantially
as shown in FIG. 6, wherein area .[.A.]. .Iadd.C .Iaddend.is a
representation of the negative pressure distribution over said
means defining an upper surface adjacent said tip, and said means
defining an upper surface being further constructed and contoured
adjacent said root to achieve a negative pressure profile chordwise
across said upper surface substantially as shown in FIG. 5, wherein
area .[.C.]. .Iadd.A .Iaddend.is a representation of the negative
pressure distribution over said means defining an upper surface at
said root, said means defining an upper surface and said means
defining a lower surface being further constructed and contoured to
achieve a positive lift coefficient .[.greater than.]. .Iadd.of
.Iaddend.about 1.7 .Iadd.maximum .Iaddend.at said root and
.[.greater than.]. .Iadd.of .Iaddend.about 1.0 .Iadd.maximum
.Iaddend.at said tip, said negative lift coefficient varying
nonlinearly spanwise when each of said blades is in a forward pitch
mode, and to achieve a negative lift coefficient greater than about
-1.0 from said root to said tip, and said negative lift coefficient
varying nonlinearly spanwise when each of said blades is in a
reverse pitch mode.
9. The propeller of claim 8, wherein said means defining an upper
surface is further constructed and contoured to have an airfoil
cross section achieving a positive upper camber throughout the
chord length of each of said blades.
10. The propeller of claim 9, wherein said means defining a lower
surface area is further constructed and contoured to have an
airfoil cross section achieving both positive lower camber adjacent
said leading edge of said blade and negative lower camber adjacent
said trailing edge of said blade.
11. The propeller of claim 10, wherein said positive lower camber
occurs within the leading 65 percent of the chord length as
measured from said leading edge, and said negative lower camber
occurring within the trailing 35 percent of the chord length of
each of said blades.
12. The propeller of claim 11, further comprising a hub, each of
said blades being rotatably mounted on said hub for rotation about
a radial axis.
13. The propeller of claim 8, wherein each of said blades is
further constructed and contoured to achieve a planform profile
having an aspect ratio between about 2.0 and about 2.2.
14. The propeller of claim 13, wherein each of said blades is
further constructed and contoured to achieve a planform profile
having a thickness-to-chord ratio varying nonlinearly spanwise
between about .[.21.]. .Iadd.23 .Iaddend. percent at said root and
about .[.10.]. .Iadd.9 .Iaddend.percent at said tip.
Description
BACKGROUND OF THE INVENTION
This invention relates to propeller blades designed to
aerodynamically generate thrust and more particularly to propeller
blades used on hovercraft, airboats, dirigibles, or other vehicles
requiring maximum thrust at static or low vehicle velocities.
Existing fixed-wing propeller blades and helicopter rotor blades
must meet the operational requirements of take-off, climb, cruise,
and, in the case of helicopters, autorotation. As a result, current
blades are designed to efficiently move aircraft at the relatively
high airspeeds necessary to maintain flight. These prior art
blades, however, have failed to effectively and efficiently develop
the thrust needed to meet the design parameters for vehicles such
as hovercraft.
One major drawback of conventional propellers and rotors is the
lack of thrust developed at static velocities. This is due to
current blade configurations that operate in an
over-80-percent-stalled condition at zero velocity. In addition,
current blades cannot generate the reverse thrust that hovercraft
require to reduce speed, reverse direction, or maneuver. Another
significant problem is poor environmental durability. The sand and
dust particles, as well as water droplets, which are present around
hovercraft operating low to the surface, quickly erode the leading
edges and tips of low-profile, high-aspect ratio blades. Finally,
prior art blades generate high noise levels due to blade tip
speed.
SUMMARY OF THE INVENTION
The present invention provides a propeller blade design that
generates a greater maximum thrust at static velocities in both
forward and reverse pitch modes than has previously been achieved.
Further, the present invention achieves the foregoing and
simultaneously lowers the noise level by reducing the tip speed of
the propeller blades and decreasing the local sonic velocity over
the blade surface. Additionally, the present invention, while still
achieving the foregoing, increases the propeller blade life and
reliability by reducing erosion and pitting due to the impact of
water, sand, and insects. This is accomplished by decreasing the
propeller blade tip speed, increasing the thickness of the leading
edge of the propeller blade, and achieving a more uniform pressure
distribution across the upper surface of the propeller blade.
The foregoing advantages are achieved in accordance with this
invention by providing a propeller blade profile having a higher
lift coefficient that avoids the blade-stalling condition that can
occur at maximum thrust and low or static velocities. This is
accomplished by increasing the radius of the leading edge and
allowing a greater thickness-to-cord ratio along the span of the
propeller blade. In addition, the planform profile of the propeller
blade has a lower aspect ratio achieved by shortening the span and
increasing the chord of the blade. Further, the contour of the
lower propeller blade surface has a positive lower camber adjacent
the leading edge but then reverses to a negative lower camber
adjacent the trailing edge to provide aft loading. The aft loading
and higher blade profile causes a redistribution of the negative
pressure across the upper surface by eliminating a sharp or steep
negative pressure gradient peak that exists over the leading edge
of prior art blades and shifting the negative pressure to apply
more evenly across the upper surface of the propeller blade.
In accordance with this invention, a unique propeller blade profile
is provided wherein a generous leading edge radius leads to a
thicker airfoil cross section having positive upper camber
throughout the chord length of the upper surface and both positive
and negative lower camber throughout the chord length of the lower
surface. In addition, the aspect ratio of the propeller is lower
than other propellers to reduce the tip speed while, at the same
time, the increased chord length provides greater lift. In
addition, a substantially uniform angle of initial pitch is
achieved spanwise by nonlinearly varying the twist of the propeller
blade from the root to the tip. This substantially uniform pitch
avoids propeller blade stalling and more efficiently develops
maximum static thrust. Further, the negative lower camber, combined
with the greater leading edge radius and increased thickness of the
airfoil, increases the aft loading on the propeller blade and
lowers and broadens the negative pressure spike or peak on the
upper leading edge of the blade. This results in a lower noise
level and a decreased amount of erosion on the leading edge of the
blade and blade tip from water droplets, dust, insects, and other
particles. The increased leading edge radius further allows for the
attachment of a protective device along the leading edge of the
propeller blade to increase its life.
In the preferred embodiment of this invention the leading edge
radius of the propeller blade is relatively large when contrasted
with prior blades. The thickness of the blade cross section in
terms of chord length varies nonlinearly from about 23 percent at
the root to about 9 percent at the tip of the blade. A twist angle
is also provided to achieve a uniform pitch spanwise from about 28
degrees at the root to about 6 degrees at the tip. The maximum
attainable lift coefficient at static velocities varies nonlinearly
spanwise across the blade from a maximum of about 1.7 at the root
to about 1.0 at the tip when the blade is rotated to a forward
pitch mode, and a maximum attainable negative lift coefficient of
about -1.0 from the root to the tip when the propeller blade is
rotated about the hub to a reverse pitch mode. The aspect ratio
varies between 2.0 and 2.2. A substantially uniform negative
pressure distribution exists at the tip of the propeller blade and
a slightly more rounded and broader negative pressure curve exists
at the root.
A practical advantage realized through implementation of this
propeller blade design is increased aerodynamic performance.
Current propeller systems used on hovercraft limit the vehicle to
50 percent of its design goal speed and 75 percent of its design
payload. By utilizing the new propeller blade, which increases
aerodynamic performance through increased lift and efficiency, cost
rebuilding of hovercraft propulsion systems can be avoided with
easier and more economical retrofitting. This unique blade not only
gives greater forward speed, but the increased reverse thrust that
is available renders hovercraft more maneuverable, easier to stop,
and gives greater reverse speed capabilities.
An additional important feature is the increased life and
reliability. The larger leading edge radius better prevents
chipping and cracking that can occur from erosion and pitting along
the leading edge and tip. Durability can be further increased
because the leading edge contour provides an excellent shape for
attaching a protective device if that proves needed.
Additionally, the blade is safer and quieter. The reduced diameter
and tip speed, as well as redistribution of negative pressure
gradients, keeps noise levels lower, thus not jeopardizing hearing
and being less of a disturbance to crew and passengers. The smaller
diameter propeller is also safer because of increased strength and
reliability.
Finally, the present invention provides a cost savings as a result
of more efficient use of available horsepower and resulting
reduction in fuel consumption. The new blade design is easily
adaptable to retrofitting on current hubs and propeller systems. No
changes would need to be made in drive system or mounting
structures. The increased power, maneuverability, and speed would
allow hovercraft to operate with greater payloads and
capabilities.
BRIEF DESCRIPTIONS OF THE DRAWINGS
Other objects and advantages of the present invention will become
apparent to one skilled in the art after a reading of the following
description taken together with the accompanying drawings, in
which:
FIG. 1 is a plan view of a propeller showing the upper surface of
the attached blades;
FIG. 2 is a cross-sectional view of a propeller blade;
FIGS. 3a through 3d are four cross-sectional contours of the
propeller blade taken along section lines 3a--3a through 3d--3d of
FIG. 1;
FIG. 4 shows the cross-sectional contours of FIGS. 3a-3d
superimposed on one another;
FIG. 5 is a cross-sectional illustration of the pressure
distributions at the root station of the propeller blade;
FIG. 6 is a cross-sectional illustration of the pressure
distributions at the tip station of the propeller blade;
FIG. 7 is a graph illustrating the lift characteristics of a prior
art propeller blade; and
FIG. 8 is a graph illustrating the lift characteristics of the
present invention.
DETAILED DESCRIPTION
Referring to FIG. 1, a propeller 10 has blades 12 mounted to hub
14. Although a three-bladed configuration is shown, it is to be
understood that other configurations such as two-bladed propellers
or four-bladed propellers may be used without departing from the
spirit of the present invention. In addition, blades 12 may either
be rotatably or fixedly mounted to hub 14, although blades 12 are
optimally designed to operate in a forward- and reverse-pitch mode,
thus requiring rotation.
The unique planform profile of blade 12 is typically measured as an
aspect ratio or the ratio of the span of blade 12, represented by
line a, divided by the chord of blade 12, represented by line b.
The present invention achieves greater thrust at lower static
velocities by having a chord length varying from 18 inches to 26
inches and, more preferably, a constant chord length of about 21
inches spanwise, and by having a radius measured from hub 14
varying between 40 inches to 60 inches, and more preferably a
constant span along leading edge 20 and trailing edge 22 of 45
inches. This corresponds to an aspect ratio in the range of 1.5 to
3.3 and preferably in the range of 2.0 to 2.20.
Referring now to FIG. 2, the currently preferred construction of
the propeller blade 12 is illustrated. Basically, the blade 12 has
a moncoque shell with a leading edge bull nose 20 that is
configured in accordance with the aerodynamic requirements of the
blade. Upper and lower skins 24a and 24b are fixed in a
conventional manner to the bull nose 20 and extend rearwardly to a
trailing edge extrusion 25 configured again in accordance with the
aerodynamic requirements of the blade. The trailing edge extrusion
25 is joined to the upper and lower skin 24a and 24b again by
conventional fasteners. The outer tip of the wing carries a
conventional fairing. The blade 12 has a full rib 27a at its root
and partial ribs 27b, 27c and 27d at spanwise locations outwardly
from the root rib 27a. Ribs 27b and 27c are located at spanwise
stations 16 inches and 22 inches from the center of rotation. The
last rib 27d is positioned approximately midspan of the blade 12. A
cylindrical spar or torque tube 29 extends up the center of the
blade through apertures in root rib 27a and partial ribs 27b, 27c
and 27d. The torque tube terminates inwardly from the root in a
conventional blade shank 29 that couples in a conventional manner
to the blade hub.
Referring now to FIGS. 3a through 3d, four sections of blade 12 at
various stations along its span are illustrated FIG. 3a is taken at
the root station; FIG. 3b is taken at an inboard station; FIG. 3c
is taken at the outboard station; and FIG. 3d is taken at the tip
station. In the embodiment shown, the root, inboard, outboard, and
top profiles are taken at 10, 32, 38, and 45 inches, respectively,
from the center of rotation.
Referring first to FIG. 3a, blade 12 has a leading edge 20 and
trailing edge 22. Chord line 30 represents the distance from the
outside of leading edge 20 to the outside of trailing edge 22.
Upper surface 32 defines the upper camber of blade 12 and lower
surface 34 defines the lower chamber of blade 12. The radius of
leading edge 20 and the magnitude of the upper and lower camber, as
measured relative to chord 30, will determine the maximum lift
potential of the blade, denominated as the coefficient of lift. The
upper camber is a distance typically measured from upper surface 32
perpendicularly to chord 30, at its greatest point; and the lower
camber is typically measured from lower surface 34 perpendicularly
to chord 30, at its greatest point. Positive camber occurs when a
surface maintains a convex curve away from chord 30, and negative
camber occurs when a surface maintains a concave curve toward chord
30 as illustrated by curve 36 along lower surface 34.
The present invention achieves a higher coefficient of lift than
prior art blades by having an upper surface that defines a positive
camber throughout the chord length of blade 12, and having both a
positive and negative lower camber along the chord length of blade
12. Adjacent the root, positive lower camber should preferably
occur within the first 65 percent of the chord length measured
adjacent leading edge 20; and negative lower camber should
preferably occur in the remaining 35 percent of the chord length
adjacent trailing edge 22 as shown by curve 36. This negative
camber adjacent the root progresses to a positive camber between
the root and inboard sections.
Another important ratio in evaluating lift and efficiency is the
thickness of the cross section profile between upper surface 32 and
lower surface 34, typically measured perpendicularly through chord
30 at the point of greatest distance. In the present embodiment,
the thickness-to-chord ratio of blade 12 would vary nonlinearly
spanwise from about 23 percent at the root station to about 9
percent at the tip station. The cross-sectional profile of the
blade is accurately depicted at stations 10, 32, 38 and 45,
respectively. The curvature of the upper and lower surfaces at each
of those stations is accurately defined by the following formula:
##EQU1## wherein x is the location on the chord measured from the
leading edge and y is the distance of the upper or lower surface
from the chord line. The parameters a,b, c,d,e, and f are set forth
in Table I below:
TABLE ______________________________________ Station 10 Station 32
Station 38 Station 45 ______________________________________ UPPER
SURFACE a .23766 .15921 .16177 .11887 b .15650 .12547 -.013454
.0068211 c -1.4017 -.78829 -.31984 -.29976 d 2.3766 1.3817 .54537
.52469 e -2.0787 -1.4758 -.66596 -.5987 f .70347 .59913 .29375
.24937 LOWER SURFACE a -.3381 -.1769 -.13873 -.12229 b .5976 .30416
.22654 .20148 c -.62181 -.55967 -.44032 -.41682 d -.33374 .95425
.78215 .77721 e 1.6804 -.92532 -.7364 -.72913 f -.99487 .40228
.30649 .28928 ______________________________________
Referring now to FIG. 4, the four section profiles are overlaid to
illustrate the geometric twist using the plane of rotation of the
blade 40 as a reference. Root chord 30 is the chord at the root
station; inboard chord 37 is the chord at the inboard station;
outboard chord 38 is the chord at the outboard station; and tip
chord 39 is the chord at the tip station of blade 12. The twist
angle, local sonic velocity and Reynold's number at each station
along the span of the blade is set forth in the following Table
II:
TABLE II ______________________________________ Reynold's No.
Station Location Twist MI (.times. 10.sup.6)
______________________________________ Root 10 +27.5.degree. 0.05
0.7 Inboard 32 +9.2.degree. 0.32 4.0 Outboard 38 +7.5.degree. 0.47
6.6 Tip 45 +6.5.degree. 0.64 8.0
______________________________________
The location represents the radial distance in inches of a given
station from the center of the hub 14. Twist is the initial fixed
angle of incidence with respect to the plane of rotation at each of
the stations. Typically, the angle is positive when leading edge 20
is higher than trailing edge 22. In Table I the twist angle is
measured between chord 30 and reference plane 40. M1 is the free
stream local sonic velocity measured at a given station. The
Reynold's number represents the airflow characteristics at each
blade station.
Referring to FIGS. 5 and 6, the positive and negative pressure
distributions achieved by the present invention are shown
diagrammatically. FIG. 5 illustrates the blade profile at the root
station as having a negative pressure distribution over upper
surface 32 substantially as shown by Area A. Adjacent the leading
edge, the negative pressure peak of Area A is significantly lower
and broader than has been achieved by prior art blades. The lower
and broader peak leads to a lower local sonic velocity resulting in
a quieter blade. Moreover, shockwave producing peaks or spikes are
eliminated, also leading to a quieter blade. The magnitude of the
negative pressure decreases near the trailing edge, but still
provides substantial loading adjacent the trailing edge, which
increases blade efficiency. Area B represents the positive pressure
distribution across the lower surface 34. Negative camber curve 36
creates greater positive pressure under trailing edge 22. This aft
loading helps to further reduce the negative pressure peak over
leading edge 20 and shift some negative pressure over upper surface
32 adjacent trailing edge 22.
The positive and negative pressure distribution over the tip
station of blade 12 is shown in FIG. 6. Area C represents the
negative pressure distribution over upper surface 32. The pressure
distribution is substantially uniform across the entire chord
length of upper surface 32, which keeps drag lower than current
blades and minimizes boundary layer separation, especially over
trailing edge 22. Prior art blades have failed to achieve this
laminar airflow under static or low velocities, resulting in a
stalled condition and increased drag and decreased performance.
FIGS. 7 and 8 provide a comparison of the lift characteristics for
prior art airfoils, such as the Clark Y and NACA 16 series
airfoils, currently used in hovercraft propeller systems, and the
lift characteristics of a propeller constructed in accordance with
the present invention. Referring first to FIG. 7, the lift
coefficient is plotted versus spanwise blade station shown in
increments with the value 1 being at the blade tip and value 0.3
being adjacent the root. The theoretical coefficient of lift for
the prior art blade is calculated in accordance with the blade
element method and provides the spanwise distribution of the
theoretical maximum coefficient of lift. The theoretical
coefficient of lift is calculated at static conditions, that is a
zero velocity, with the application of 1,200 horsepower at 1,900
RPM. The calculations are conducted in a forward and reverse thrust
mode with a blade setting of 18.degree. and -24.degree.,
respectively, at the 0.8 station.
The line labeled C.sub.L MAX is the maximum coefficient of lift
that can actually be developed under these operating conditions.
The region between the C.sub.L MAX line and the theoretical
coefficient of lift line represents the region over which the prior
art blade is stalled. As can readily be observed, in the forward
thrust mode, the prior art blades are typically stalled over
greater than 60 to 70% of the blade span. Similarly, in reverse
thrust mode, the prior art blade because of its cross section is
rather inefficient, generating a maximum actual coefficient of lift
of about 0.7. Again, as can be observed, the prior art blade is
stalled over about 50% of its span in the reverse operating
mode.
If more horsepower of RPM or both were applied in the reverse mode,
the maximum coefficient of lift would change little. The
theoretical coefficient of lift, however, moves downwardly thus
increasing the stalled region and decreasing the efficiency of the
blade. In the reverse mode, the region between the C.sub.L MAX and
the theoretical coefficient of lift represents a margin whereby
increased horsepower does not result in increased stall. However,
since the theoretical coefficient of lift line will actually move
downwardly, the percentage of the blade which is stalled will
increase with increased application of power.
Referring to FIG. 8, the lift characteristics of a blade
constructed in accordance with the present invention at the same
operating conditions as the prior art blade are illustrated. The
graphs are generated for forward and reverse modes at blade
settings of 23.degree. and -30.degree., respectively, at station
0.8. In the forward thrust mode, the blade constructed in
accordance with the present invention is designed so that at the
operating conditions tested, the theoretical coefficient of lift as
well as the actual coefficient of lift achievable under those
conditions is approximately the same. Thus, no portion of the blade
in the forward thrust mode is operating in a stalled region. If,
however, additional horsepower or RPM or both were applied to the
blade constructed in accordance with the present invention, the
theoretical as well as the actual coefficient of lift will
increase, thus moving C.sub.L ACTUAL up toward the line labeled
C.sub.L MAX. The latter line represents the maximum coefficient of
lift achievable prior to stalling the wing, with a blade designed
in accordance with the invention. Thus, it can be seen that the
coefficient of lift of the blade constructed in accordance with the
present invention is approximately the same of that of the prior
art blades. However, since none of the blade constructed in
accordance with the present invention is operating in a stalled
condition, the blade is operating in a much more efficient mode
resulting in a higher thrust output for a given power input.
In the reverse thrust mode, the blade is designed so that the
maximum achievable coefficient of lift is about 1.0 across the
entire span of the blade. Again, at these operating conditions, the
actual coefficient of lift and the theoretical coefficient of lift
are again about the same. Under these conditions, only about 25-30%
of the blade is operating in the stalled region with the remainder
of the blade operating under lift conditions that are significantly
greater than those achievable by prior art blades. Thus, one of the
most important characteristics of the blade constructed in
accordance with the present invention is fully observable, that is,
it has very superior reverse thrust characteristics.
Having described the invention in its preferred embodiment, it is
to be realized that changes and modifications therein may be made
without departing from the essential concepts representing the
advancements in this art. For example, shrouding the propeller will
further increase thrust by preventing spillage of air spanwise over
the tip. Shrouding also affords greater protection from physical
injury to crew or passengers. It is therefore intended that the
scope of the claims that follow be limited by their definitional
terms and equivalents thereof.
* * * * *