Note: Descriptions are shown in the official language in which they were submitted.
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TITLE OF THE INVENTION
Rotary ramjet engine
FIELD OF THE INVENTION
[0001] The present invention relates to the general field of engines
for producing mechanical power and is particularly concerned with a rotary
ramjet engine.
BACKGROUND OF THE INVENTION
[0002] The conversion of chemically stored energy into useful work
has been the goal of engine designers since the creation of internal
combustion
engines utilizing the Otto cycle. As is well known, the most widely available
devices nowadays for converting fossil fuel energy into rotational power are
the
Otto and Diesel cycle reciprocating internal combustion engines and the
Brayton cycle gas turbines.
[0003] The reciprocating or piston-type internal combustion engines
typically demonstrate relatively low fuel efficiency. Indeed, typically less
than
20% of the chemical energy stored in the fuel is transformed into useful
mechanical energy. This relatively low efficiency is imputable to many
factors.
[0004] For example, complex mechanical systems are required to
transform the reciprocating motion of the piston into the rotary motion of the
drive shaft. Also, the mechanical properties of the materials used for
building
conventional combustion chambers significantly limits the allowable
temperature and pressure in the combustion chamber and, thus, limits the
thermal efficiency of the engine.
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[0005] Furthermore, with most conventional reciprocating internal
combustion engines, the combustion of the fuel occurs at ordinary rates.
These ordinary rates of combustion result into prolonged heating of the
combustion chamber which produces more degradation at the chamber wall
per unit volume of fuel burned than when the fuel is burned at a faster rate.
This, in turn, limits the specific power or power-to-weight ratio of the
engine.
[0006] Rotary engines compared to reciprocating engines
significantly reduce the mechanical complexity by eliminating the need to
transform the reciprocating piston motion to rotary motion of the drive shaft.
The prior art is replete with various types of rotary engines in which the
rotor
has a plurality of circumferentially spaced combustion chambers formed with
ducts to exhaust combustion products in order to provide reaction forces.
However, most conventional rotary engines do riot provide substantially
improved efficiency or power-to-weight ratio over the conventional
reciprocating
engines probably at least in part because the combustion of fuel still occurs
at
ordinary rates.
[0007] In view of the scarcity and high costs of engine fuels, engine
designers and engineers have been grappling mostly with the fundamental
problems of exhaust emissions and pollutants and increased fuel economy, yet
striving to improve performance in these areas without sacrificing already
compromised engine performance and efficiency. Over time, numerous
proposals have been set forth to reduce pollution and increase engine and fuel
performance. Each has some distinct disadvantage because of its interaction
with other engine parameters inherent in the Otto or Diesel cycle engines.
[0008] Although engine designing efforts have been directed mostly
towards these fundamental problems and, in particular, towards improving the
efficiency of the fuel-to-work conversion, other engine parameters may in some
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situations be considered at least as important. For example, in some specific
settings, the so-called specific power or power-to-weight ratio is sometimes
considered a crucial design and operational criteria.
[0009] For example, it is well known that with various types of
vehicles, weight may become a critical factor. Indeed, extra weight in
vehicles
such as automobiles and airplanes in particular require substantial propulsion
power and also reduce maneuverability. When the specific power of the engine
is not optimized, the engine itself may constitute an important fraction of
the
total vehicle weight. Accordingly, many types of vehicles could greatly
benefit
from a simple propulsion system having a relatively high power-to-weight ratio
in order to minimize the overall vehicle mass.
[0010] Furthermore, there also exist many situations wherein
engines are only required to provide power for limited periods of time. Some
of
these situations would also greatly benefit from an overall reduction of the
engine weight and, hence, optimization of the specific power since the engine
may be considered as a "dead mass" for most of the operational cycle.
[0011] In the field of aerospace, typical examples of situations
wherein optimization of the specific power could prove to be particularly
interesting include upholding of the gyroscopic stabilization of satellites,
generating power for space weapons and tools, providing power for future
human space stations, motorizing some of the mechanisms during spatial
machines take-off, providing propulsion for personal flight vehicles and so
forth.
[0012] In the automotive industry, hybrid vehicles in particular could
potentially greatly benefit from an engine able to demonstrate a relatively
high
specific power. Indeed, in such hybrid vehicles, piston engines are sometimes
used as generators and peak power leveler engines to counterbalance the
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lower power and relatively low autonomy of conventional batteries. Such
vehicles would hence greatly benefit from auxiliary engines having a
relatively
high specific power that could be used as a peak leveler. Also, electrically
powered vehicles such as an hydrogen fuel cell electric vehicle could benefit
from an engine adapted to act both as a constant or near constant energy
storage flywheel and as a peak power supplier when needed such as during
acceleration.
[0013] In the field of biomechanics various devices such as limb
prosthesis or the like could also greatly benefit from a relatively light
actuator
able to provide relatively short bursts of power. In general, any application
wherein there exists a need for a providing peak power for relatively small
periods of time and without sacrificing weight criterias could benefit from an
engine optimized for specific power.
[0014] As is well known, of the currently available energy converters,
the so-called gas turbine appears as one of the best design in terms of
specific
power. Conventional gas turbine engines have generally included a stationary
combustion chamber or burner where injected kerosene or other fuel and air
from a compressor is mixed and burned. The burnt gas may pass through a
duct which directs it against the blade of a rotating turbine blade wheel that
delivers through its shaft. High efficiency and high power output from such
engines depend on the use of gas jets of high energy being directed at the
turbine blade wheel. However, if the jet energy increases to a substantially
high level, large thermal and mechanical stresses are imposed on the blade
which may cause mechanical failure.
[0015] Indeed, relatively large mechanical stresses are produced in
available engines by reason of the high rotational speeds for typical turbine
blade wheel diameters. When these mechanical stresses are combined with
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high temperatures imposed by the gas jet, the conditions become close to the
limit of strength of the best available turbine blade materials.
[0016] Higher gas jet energy can be readily obtained, but they
cannot be efficiently used since if the blade wheel is allowed to rotate
faster,
this causes excessive mechanical stress, while if the gas jet is moving too
fast
relative to the blade, for example more than about Mach 1, then this causes
excessive heating of the blade.
[0017] Typically, with conventional gas turbine engines, long cycle
life is obtained by using large quantities of extra air to produce gas
released at
moderate temperature and velocity and by using moderate rotational speed of
the turbine blade wheel. The result is relatively low efficiency even for
large
units and limited power output for a given size of engine.
[0018] The mechanical resistance of the materials used for building
the gas turbines is hence often considered the limiting factor that hinders or
limits the specific power performances of conventional gas turbines. Also,
conventional gas turbines are usually not considered as a commercially
interesting high specific power propulsion system because of their inherent
high
costs and complexity.
[0019] The need for providing engines capable of demonstrating a
relatively high specific power while has been recognized in the prior art. One
of
the proposed solutions involves the use of shock waves instead of mechanical
compressors for compressing the combustible fluid. Since shock waves allow
for relatively high rotational speeds, relatively low pressures are required
to
obtain potentially interesting power values as opposed to conventional turbo-
engines requiring relatively large forces because of their relatively low
rotational
speed.
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[0020] Various types of thrusters using shock wave combustion to
provide thrust are known. For example, ramjets are widely known and have
been used primarily in aerospace applications since the 1940. In an aerospace
propulsion ramjet, air in ingested into an engine inlet at supersonic speeds
caused by the forward motion of an airplane or missile. The air is rammed into
a smaller opening between a center-body and the engine side wall generating a
series of shock waves. These shock waves compress and decelerate the air to
subsonic speeds while, at the same time, dramatically raising working flow
pressure and temperature.
[0021] The ramjet effect may also be achieved in a stationary
platform by passing an accelerated flow of air over raised sections machined
on the rim of a rotor disc. Combined with the high rotation rate of the rotor,
this
produces a supersonic flow relative to the rotor rim. Interaction between the
raised sections of the rim which are rotating at supersonic speeds and the
stationary engine case creates a series of shock waves that compress the air
stream in a manner similar to ramjet inlets on a supersonic missile or
aircraft.
[0022] Over the years, the combustion chamber configuration, the
configuration of ramjet engine inlets, the fuel injection requirements and
ignition
requirements have been the subject of much studies and technical
development. Ramjets have also been experimentally employed to assist in
the rotation of helicopter blades about a central shaft.
[0023] The prior art has further shown examples of rotary motors
using shock wave combustion produced by the so-called ramjet effect. For
example, U.S. Patent 5,372,005 et 5,709,076 both naming S.P. Lawlor as
inventor and issued respectively in 1994 and 1998 disclose both a method and
an apparatus for generating power using rotating ramjets which compress inlet
air and expand exhaust gases against stationary peripheral walls.
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[0024] The engine disclosed in the hereinabove mentioned patents
is commonly referred to as a ramgen. The ramgen typically includes a pair of
thrusters mounted in a diametrically opposed relationship relative to each
other
on the engine rotor
[0025] The tangential thrust produced by the thrusters provides the
rotational output power. Although interesting, the ramgen nevertheless suffers
from inherent drawbacks. One of these drawbacks relate to the fact that the
inlet air of a given thruster may be potentially contaminated by the exhaust
air
of the other thruster. A large fan positioned upstream relative to the ramgen
inlet must hence be used in order to discharge exhaust gases.
[0026] Also, ramgen-type engines are mainly designed for the
production of electrical and mechanical power at medium size electrical or
mechanical power plants. Such medium size mechanical or electrical power
plants, typically in the range of 10 to 100 megawatts are typically required
in
industrial applications including stationary electric power generating units,
rail
locomotives, marine power systems and the like.
[0027] Power plants in this general size range are also typically
suited for use in industrial co-generation facilities increasingly employed to
service industrial thermo-power needs while simultaneously generating
electrical power. Obviously, ramgen-type engines are mainly concerned with
the efficient conversion of fuel input to electrical output as opposed to
being
concerned with specific power. Accordingly, design choices such as the use of
only two thrusters inherently limit the ability of ramgen-type engines to
provide
high specific power.
[0028] The prior art has also shown some examples of other types of
rotary supersonic combustion chambers. For example, U.S. Patents 3,971,209
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and 4,199,296 both naming R.S. Chair as inventor and issued respectively in
1976 and 1990 disclose respectively various embodiments of gas generators
and gas turbine engines. The structure disclosed in the above-mentioned
patents is commonly referred to as a rambine.
[0029] The rambine, although somewhat interesting, may be
considered as "self-propelled compressor-combustion chamber combination"
as opposed to an autonomous engine. Furthermore, the angle of attack at the
inlet diffuser of the rambine typically does not coincide with the radial
direction
of the engine as is the case with conventional compressors. This may
potentially limit the overall performance of the rambine.
[0030] Accordingly, there exists a need for an improved engine
allowing for the conversion of chemically stored energy into output power with
relatively high specific power characteristics.
OBJECTS OF THE INVENTION
[0031] It is a general object of the present invention to provide a
mechanical engine having a relatively high power-to-weight and high power-to-
volume ratio when converting chemically stored energy into mechanical output
power.
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SUMMARY OF THE INVENTION
[0032] In accordance with the present invention, there is provided a
rotary ramjet engine for producing an output power about an output shaft
extending substantially along a shaft axis.
[0033] The engine comprises an annular shape thruster base
defining a radially inwardly located thruster base inner surface and an
opposed
radially outwardly located thruster base outer surface. The thruster base
includes at least one thruster extending substantially radially and inwardly
from
the thruster base inner surface. The thruster base is disposed for rotary
motion
along a substantially circular thruster path positioned in a rotary plane
substantially perpendicular to the shaft axis. The thruster base is capable of
generating a thruster torque about the shaft axis, the thruster base
generating a
thruster centrifugal force acting thereon when rotating along the thruster
path.
[0034] The engine also comprises a mechanical thruster-to-shaft
coupling means operatively coupled to both the thruster base and the output
shaft for transmitting the thruster torque to the output shaft.
[0035] The engine further comprises a centrifugal force
compensating and annular-shaped reinforcement wall operatively coupled to
the thruster base for reacting to the centrifugal force and compensating for
the
latter so as to maintain the thruster base in the thruster path when the
thruster
base is rotated in the rotary plane. The annular-shaped reinforcement wall is
a
piece distinct from the thruster base and from the mechanical thruster-to-
shaft
coupling means. The annular-shaped reinforcement wall defines a radially
inwardly located reinforcement wall inner surface. The annular-shaped
reinforcement wall is so configured that at least a portion of the inner
surface of
the reinforcement wall is in abutting contact with the thruster base outer
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surface. The reinforcement wall includes a one-dimensional carbon-based
composite material
[0036] The mechanical thruster-to-shaft coupling means and the
centrifugal force compensating and annular-shaped reinforcement wall are
allowed to perform their respective force transmitting and compensating
function substantially independently from each other so as to substantially
reduce the need for the mechanical thruster-to-shaft coupling means to react
to
and compensate for the centrifugal force.
BRIEF DESCRIPTION OF THE DRAWINGS
[0037] The foregoing aspects and other features of the present
invention are explained in the following non-restrictive description of an
illustrative embodiment thereof, taken by way of example only in connection
with the accompanying drawings, wherein:
[0038] Figure 1, in an exploded view, illustrates some of the
components of a rotary ramjet engine in accordance with an embodiment of the
present invention;
[0039] Figure 2, in an axially transversal cross-sectional view,
illustrates some of the components of a rotary ramjet engine in accordance
with
an embodiment of the present invention;
[0040] Figure 3, in a radially transversal cross-sectional view,
illustrates some of the components of the rotary ramjet engine shown in Figs.
1
and 2;
[0041] Figure 4, in a partial axially transversal cross-sectional view,
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illustrates some of the dimensional characteristics of part of a rotor
component
incorporated in a rotary ramjet engine in accordance with an embodiment of the
present invention;
[0042] Figure 5, in a partially exploded sectional view, illustrates
some of the features of a rotor, thruster blades, a thruster base and a
reinforcement wall, all part of a rotary ramjet engine in accordance with an
embodiment of the present invention;
[0043] Figure 6, in a perspective view, illustrates a flow guiding
blade used for forming a thruster incorporated in a rotary ramjet engine in
accordance with an embodiment of the present invention;
[0044] Figure 7, in a perspective view, illustrates a set of flow
guiding blades and a thruster base part of a rotary ramjet engine in
accordance
with an embodiment of the present invention;
[0045] Figure 8, in an elevational view, illustrates the deployed
configuration of a flow duct formed by a ramjet thruster, part of a rotary
ramjet
engine in accordance with an embodiment of the present invention;
[0046] Figure 9, in a partial elevational view, illustrates a set of three
flow guiding blades positioned in an adjacent relationship relative to each
other
to define a corresponding set of flow ducts, part of corresponding ramjet
thrusters incorporated in a rotary ramjet engine in accordance with an
embodiment of the present invention;
[0047] Figure 10, in an elevational view, illustrates the deployed
configuration of a pair of ramjet ducts formed partly by a corresponding pair
of
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flow guiding blades, part of a rotary ramjet engine in accordance with an
embodiment of the present invention;
[0048] Figure 11, in a cross-sectional view taken along arrows 11-11
of Fig. 10, illustrates the configuration of cooling and bleeding channels,
part of
ramjet thrusters incorporated in a rotary ramjet engine in accordance with an
embodiment of the present invention;
[0049] Figure 12, in a top end view of Fig. 10, illustrates some of the
features of ramjet thrusters, part of a rotary ramjet engine in accordance
with
an embodiment of the present invention;
[0050] Figure 13, in a left side view of Fig. 10, illustrates some of the
features of ramjet thrusters, part of a rotary ramjet engine in accordance
with
an embodiment of the present invention;
[0051] Figure 14, in a right side view of Fig. 10, illustrates some of
the features of ramjet thrusters, part of a rotary ramjet engine in accordance
with an embodiment of the present invention;
[0052] Figure 15, in a partial schematic view, illustrates the
configuration and positioning of shock waves generated in a thruster, part of
a
rotary ramjet engine in accordance with an embodiment of the present
invention at a speed of 1.95 Mach;
[0053] Figure 16, in a detailed view taken along arrows 16-16 of Fig.
15, illustrates in greater detail, the configuration and positioning shown in
Fig.
15;
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[0054] Figure 17, in a partial schematic view, illustrates the
configuration and positioning of shock waves within the thruster, part of a
rotary
ramjet engine in accordance with an embodiment of the present invention at a
speed of 2.5 Mach;
[0055] Figure 18, in a detailed view taken along arrows 18-18 of Fig.
17, illustrates in greater detail the positioning and configuration of the
shock
waves shown in Fig. 17;
[0056] Figure 19, in a schematic axially transverse cross-sectional
view, illustrates the relationship between some of the components of a rotary
ramjet engine in accordance with an embodiment of the present invention;
[0057] Figure 20, in a schematic partially axially transverse cross-
sectional view, illustrates an engine assembly also part of the present
invention, including four rotary ramjet engines in accordance with an
embodiment of the present invention;
[0058] Figure 21, in a partial perspective view with sections taken-
out, illustrates a rotor component having cooling channels in accordance with
an alternative embodiment of the invention;
[0059] Figure 22, in an exploded view, illustrates a rotary ramjet
engine in accordance with yet another embodiment of the invention and,
[0060] Figure 23, in an axial cross-sectional view, illustrates the
relationship between some of the components of the rotary ramjet engine
shown in Figure 22.
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DETAILED DESCRIPTION
[0061] In an embodiment, thruster-to-shaft coupling means includes
a mechanical coupling component configured and sized for extending
substantially radially between the output shaft and the thruster; a coupling
component-to-thruster attachment means for attaching the mechanical coupling
component to the thruster; a coupling component-to-shaft attachment means
for attaching the coupling component to the output shaft.
[0062] Typically, the thruster-to-shaft coupling means allows the
thruster and the mechanical coupling component to expand and retract
substantially radially substantially independently from each other.
[0063] Conveniently, the mechanical coupling component includes a
generally disc-shaped rotor defining a pair of opposed rotor side surfaces and
a
radially outermost rotor peripheral surface, the rotor defining a rotor
rotational
axis, the rotor rotational axis being in a substantially collinear
relationship
relative to the shaft axis, the rotor being rotatable about the rotor
rotational axis.
Typically, the rotor side surfaces are configured to reduce aerodynamical drag
thereon when the rotor is rotated about the rotor rotational axis.
[0064] Conveniently, the engine further comprises a vacuum
creating means fluidly coupled to the engine for creating at least a partial
vacuum substantially adjacent at least a portion of at least one of the rotor
side
surfaces.
[0065] Typically, the centrifugal force compensating means includes
a reinforcement wall, the reinforcement wall having a substantially annular-
shaped configuration, the reinforcement wall being configured and sized so as
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to be positioned radially outwardly relative to the thruster and in abutting
contact with the latter about a reinforcement wall-to-thruster interface.
[0066] Typically, each of the flow guiding blades defines a radially
outwardly located blade contacting edge in contact with the thruster base and
a
radially opposed blade spaced edge positioned in a radially inwardly spaced
relationship relative to the thruster base; the blade spaced edge being spaced
from the casing inner wall by a blade-to-inner wall running clearance.
[0067] Conveniently, the ramjet thruster includes an inlet aperture
for receiving the combustible fluid, the inlet aperture leading flow-wise into
a
convergent inlet diffuser having a flow-wise decreasing effective diffuser
cross-
sectional area, the inlet diffuser leading flow-wise into a combustion
chamber,
the combustion chamber leading flow-wise into a divergent exhaust nozzle
having a flow-wise increasing effective diffuser cross-sectional area and the
exhaust nozzle leading flow-wise to an exhaust aperture; the inlet diffuser
being
provided with a bleeding means for bleeding a bleedable portion of the
combustible fluid from the inlet diffuser.
[0068] Typically, the inlet diffuser is an internal perforated diffuser,
the bleeding means including bleeding apertures extending through the thruster
base in the region of the inlet diffuser, the bleeding apertures being
configured,
positioned and sized for allowing an outflow volume of bleedable fluid
therethrough, the outflow volume of bleedable fluid being inversely
commensurable with the speed of flow of the combustible fluid in the inlet
diffuser.
[0069] Conveniently, each of the ramjet thrusters defines a
corresponding ramjet channel extending therethrough, the ramjet channel
being generally angled relative to the tangential direction towards the radial
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direction; the ramjet channel having substantially the configuration of the
mirror
image of a substantially deployed "S" shape wherein the "S" shape is
substantially deployed substantially along the tangential direction.
[0070] Typically, each of the ramjet thrusters allows exhaust
therefrom of combustion products resulting from the combustion of the
combustible fluid; each of the ramjet thruster including an inlet aperture for
receiving the combustible fluid, the inlet aperture leading flow-wise into a
convergent inlet diffuser having a flow-wise decreasing effective diffuser
cross-
sectional area, the inlet diffuser leading flow-wise into a combustion
chamber,
the combustion chamber leading flow-wise into a divergent exhaust nozzle
having a flow-wise increasing effective nozzle cross-sectional area and the
exhaust nozzle leading flow-wise to an exhaust aperture.
[0071] Conveniently, the inlet diffuser defines a substantially
rectilinear extrados and a substantially concave intrados both merging towards
each other for forming a diffuser throat; the exhaust nozzle defining a
substantially rectilinear intrados and an opposed substantially concave
extrados, both emanating from a nozzle throat, the combustion chamber having
a substantially biconcave configuration between the diffuser and nozzle
throat;
the extrados of the inlet diffuser being angled from the tangential direction
towards the radial direction by an inlet extrados angle; the intrados of the
exhaust nozzle being angled from the tangential direction towards the radial
direction by an exhaust intrados angle; whereby the combustible fluid
penetrates into each of the ramjet thrusters with an average flow direction
substantially parallel to the extrados of the inlet diffuser and the
combustion
products are ejected from each of the thrusters with an average flow direction
substantially parallel to the intrados of the exhaust nozzle, the
substantially
concave intrados and extrados respectively of the inlet diffuser and exhaust
nozzle being respectively adapted to act as compression and expansion ramps
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for respectively decelerating and accelerating the fluids flowing adjacent
thereto.
[0072] The proposed rotary ramjet engine allows for the conversion
of chemically stored energy into rotating mechanical power with a relatively
high specific power or power-weight or volume ratio. Also, the proposed
engine not only allows for the generation of mechanical power but also for
mechanical power storage independently or in combination with the mechanical
power generation.
[0073] Also, in at least one embodiment, the thrusters are designed
so as to be relatively compact lengthwise and positionable relative to one
another so as to optimize the number of thrusters for a given thruster
supporting diameter. This leads to an optimization of the power obtainable for
a given engine diameter, again, in turn, leading to optimization of the power-
to-
weight or volume ratio.
[0074] In at least some embodiments, the engine is also
characterized by its adaptability to various types of fuels including
conventional
hydrocarbon based fuels and hydrogen. Some of the combustible material
usable by the proposed engine allows the latter to maintain relatively high
specific power output while reducing the generation of undesirable products of
combustion such as nitrogen oxide or the like or incomplete products of
combustion such as carbon monoxide or the like. The proposed engine is
hence usable in various settings, particularly in locales having strict
environmental regulations,
[0075] Still further, the proposed engine is designed so as to reduce
the risks of breakdown in order to reduce the overall operational costs and to
allow for its usage on equipment where reliability of operation is important.
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Generally, the proposed engine allows for the generation of power to be done
in a simple and direct manner. In at least one embodiment, the proposed
engine is designed so as to have a minimal number of moving parts and so as
to avoid complex sub-systems. The rotating components of the proposed
device are designed so as to be able to withstand the stresses and strains of
rotating at relatively high tip speeds.
[0076] In at least one embodiment, the proposed engine is designed
so that the centrifugal force imparted on its torque creating components by
the
rotation of the latter is compensated or counter-balanced by a toroidal or
ring-
shaped reinforcement component specifically designed to efficiently withstand
relatively high stresses particularly when the latter are oriented in a
predetermined direction. By using a radially outwardly located reinforcement
wall, the usually larger rotating component used for transmitting the torque
to
the output shaft may be greatly reduced in size and weight.
[0077] The reduction in weight of the conventionally more massive
rotational components, mainly the rotor, leads to a substantial decrease of
the
overall engine mass and, hence, to a substantially improved specific power.
The reduction in weight of the larger components such as the rotor, in turn,
lead
to a reduction in the size and weight of other components required for
supporting the rotor such as the frame, the bearings and the like.
[0078] In at least one embodiment, the rotor of the proposed engine
is specifically configured to reduce parasitic aerodynamical drag thereon.
Also,
in at least one embodiment, the proposed engine is further provided with
vacuum creating means for creating at least a partial vacuum at strategic
locations in order to further reduce aerodynamical drag on its rotating
components.
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[0079] In at least one embodiment, the proposed engine is provided
with cooling means for cooling some of its potentially heat sensitive
components, allowing the latter to be used in relatively high temperature
environments.
[0080] Still furthermore, in at least one embodiment, the proposed
engine is provided with means for insuring a relatively smooth operation
thereof
even during operational start-up conditions. In at least one embodiment, a
perforated internal diffuser is used to allow for bleeding in inverse
proportionality to the rotational speed during initial rotational
acceleration.
[0081] Furthermore in accordance with at least one embodiment, the
thruster configuration is such that the cowling-blade interface is positioned
on
the interior surface of the blade as opposed to the conventional exterior wall
positioning used with conventional turbo machines. The pressure on the inner
wall being lower than on the external wall, efficiency losses may be reduced
compared to conventional gas turbines.
[0082] The design of some of its components is such that rupture
thereof during use may be less catastrophic than the rupture of conventional
equivalent components.
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[0083] Referring to Fig. 1, there is shown, in an exploded view, some
of the components of a rotary ramjet engine 10 in accordance with an
embodiment of the present invention. The engine 10 is intended to be used for
producing an output power about an output shaft 12. The output shaft 12
extends substantially along a shaft axis 14 and is mounted so as to be
rotatable
about the latter.
[0084] The engine 10 includes at least one, and preferably a plurality
of thrusters 16. The thrusters 16 shown throughout the Figures are of the
ramjet type. It should, however, be understood that any suitable type of
thruster could be used without departing from at least one inventive concept
of
the present invention.
[0085] As illustrated in Fig. 1, the thrusters 16 are typically disposed
substantially symmetrically relative to the shaft axis 14 for rotary motion
along a
common substantially circular thruster path positioned in a common rotary
plane substantially perpendicular to the shaft axis 14. Each individual
thruster
16 is capable of generating a corresponding thruster torque about the shaft
axis
14. Also, as is well known in the art, the rotation of the thrusters 16 along
the
thruster path generates thereon a corresponding thruster centrifugal force.
[0086] The engine 10 also includes a thruster-to-shaft coupling
means operatively coupled to both the thrusters 16 and the output shaft 12 for
transmitting the respective thruster torques to the output shaft 12. The
engine
10 further includes an independent centrifugal force compensating means
operatively coupled to the thrusters 16 for reacting to the centrifugal force
generated thereby and compensating for the centrifugal force so as to maintain
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the thrusters 16 in the thruster path when the thrusters 16 are rotated in the
rotary plane.
[0087] In at least one embodiment of the invention, the thruster-to-
shaft coupling means and the centrifugal force compensating means are
allowed to perform their respective torque transmitting and centrifugal force
compensating functions substantially independently from each other so as to
substantially reduce the need for the thruster-to-shaft coupling means to
react
and to compensate for the centrifugal force.
[0088] The thruster-to-shaft coupling means may take any suitable
form including non-mechanical coupling such as magnetic, electric or electro-
magnetic coupling (not shown) or the like without departing from the scope of
the present invention. When a mechanical coupling component is used,
various types of configurations such as spokes (not shown) or the like may be
used without departing from the scope of the present invention. Typically, the
thruster-to-shaft coupling means includes a generally disc-shaped rotor 18.
[0089] Regardless of the type of mechanical coupling component
used, the mechanical coupling component typically defines a radially innermost
located coupling component inner edge and a radially outermost located
coupling component outer edge. Typically, a coupling component-to-thruster
attachment means is used for attaching the mechanical coupling component to
the thruster 16. Similarly, a coupling component-to-shaft attachment means is
typically used for attaching the coupling component to the output shaft 12.
Again, it should be understood that various types of coupling component-to-
thruster and shaft attachment means, including structural, adhesive, magnetic,
electric, electro-magnetic and other means could be used without departing
from the scope of the present invention.
CA 02426906 2003-04-16
22
[0090] When a mechanical coupling component is used, the latter is
typically configured and sized for extending substantially radially between
the
output shaft 12 and at least one of the thrusters 16. In such instances, a
coupling component-to-thruster attachment means is provided for attaching the
mechanical coupling component to the at least one thruster 16 and a coupling
component-to-shaft attachment means is also provided for attaching the
coupling component to the output shaft 12.
[0091] When a generally disc-shaped rotor 18 is used, the rotor 18
defines a pair of opposed rotor side surfaces 20 (only one of which is shown
in
Fig. 1) and a radially outermost rotor peripheral surface 22. The rotor 18
further defines a rotor rotational axis extending in a substantially co-linear
relationship relative to the shaft axis 14.
[0092] The rotor side surfaces 20 are typically configured to reduce
aerodynamical drag thereon when the rotor 18 is rotated about the rotor
rotational axis. In one possible embodiment of the invention shown. more
specifically in Figs. 2 and 4, the cross-sectional configuration of the rotor
18 is
dividable into a pair of rotor cross-section half portions 25 substantially
symmetrically configured and positioned relative to the rotor rotational axis.
Each one of the rotor cross-section half portions 25 defines a half portion
proximal region 24 and an integrally extending half portion distal region 26
located respectively radially, proximally and distally relative to the rotor
rotational axis.
[0093] Typically, although by no means exclusively, the half portion
proximal region 24 has a substantially frusto-triangular configuration
tapering
radially outwardly while the half portion distal region 26 has a substantially
rectangular configuration. The illustrated rotor configuration not only
reduces
parasitic drag on the rotor 18 during use but also facilitates dimensionning
of
CA 02426906 2003-04-16
23
the rotor 18 at different dimensional scales without the need for complex
calculations.
[0094] Figure 4 indicates, in millimeters, typical dimensional
relationships between the axial thickness and radial position of a typical
rotor
half portion 25. It should, however be understood that these dimensional
values are only given by way of example and that other dimensional values and
rotor cross-sectional configurations could be used without departing from the
scope of the present invention.
[0095] In order to further reduce aerodynamical drag on the rotor
side surfaces 20, the engine 10 may be provided with a vacuum creating
means (not shown) fluidly coupled to the engine 10 for creating at least a
partial
vacuum substantially adjacent at least a portion of at least one of the rotor
side
surfaces 20. Preferably, when a vacuum creating means is used, the vacuum
creating means creates a vacuum substantially adjacent the substantially full
area of both rotor side surfaces 20.
[0096] The thruster-to-shaft coupling means preferably allows the
thrusters 16 and the mechanical coupling component to expand and retract
radially substantially independently from each other. The thruster-to-shaft
coupling means is hence typically designed so as to reduce or eliminate the
stresses created by discrepancies between the centrifugal radial expansions of
the at least one thruster 16 and the mechanical coupling component such as
the rotor 18. Typically, the component-to-thruster attachment means allows the
mechanical coupling component and the at least one thruster 16 to rotate
solidarity with each other while allowing a relative radial movement between
the
coupling component outer edge and the at least one thruster 16.
CA 02426906 2003-04-16
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[0097] Typically, although by no means exclusively, the coupling
component-to-thruster attachment means includes an inter-engaging tongue
and groove combination including at least one tongue 28 and at least one
groove 30. The tongue and groove combination extends between the coupling
component outer edge and the at least one thruster 16. The tongue 28 and the
groove 30 are configured and sized for maintaining the tongue 28 in
operational
contact with the groove 30 while allowing relative movement between the
tongue 28 and the groove 30 in a substantially radial direction and while
preventing relative movement between the tongue 28 and the groove 30 in
other directions.
[0098] In situations wherein a disc-shaped rotor 18 is used, at least
one and preferably a plurality of tongues 28 extend integrally and
substantially
radially from the rotor peripheral surface 22. Correspondingly, at least one
and
preferably a plurality of grooves 30 are defined by the thrusters 16.
[0099] As illustrated more specifically in Figs. 5 through 7, each
tongue 28 typically has a substantially parallelepiped-shaped tongue
configuration while each corresponding groove 30 typically has a substantially
complementary parallelepiped-shaped groove configuration. In at least one
embodiment of the invention, the tongue and groove combination is provided
with biasing means such as a spring component and dampening means such
as a shock-absorbing component (both not shown) for ensuring the dynamic
stability of the rotor 18.
[0100] In situations such as shown throughout the figures wherein
more than one thruster 16 is used, the engine 10 further includes a thruster-
to-
thruster coupling means extending between adjacent thrusters 16 for
mechanically coupling together the thrusters 16 and for maintaining the latter
in
a predetermined spaced relationship relative to each other. Typically, the
CA 02426906 2003-04-16
thruster-to-thruster coupling means includes a substantially annular-shaped
thruster base 32.
[0101] As shown more specifically in Fig. 7, the thruster base 32
defines a radially inwardly located thruster base inner surface 34 and an
opposed radially outwardly located thruster base outer surface 36. The
thrusters 16 typically extend substantially radially and inwardly from the
thruster
base inner surface 34.
[0102] As shown more specifically in Figs, 1 through 3 and 5, the
centrifugal force compensating means typically includes a reinforcement wall
38 having a substantially annular-shaped configuration. The reinforcement wall
38 is configured and sized so as to be positioned radially outwardly relative
to
the thruster base 32 and in abutting contact with the latter about a
reinforcement wall-to-thruster interface 41.
[0103] As shown more specifically in Figures 5 and 10 through 14,
the reinforcement wall 38 typically defines a radially inwardly located
reinforcement wall inner surface 40 and an opposed radially outwardly located
reinforcement wall outer surface 42. In situations wherein a plurality of
thrusters 16 are coupled together by an annular thruster base 32, the
reinforcement wall 38 is configured and sized so that at least a portion and
preferably most or all of the reinforcement wall inner surface 40 contacts at
least a portion and preferably most or all of the thruster base outer surface
36
or protrusions extending therefrom about a reinforcement wall-to-thruster base
interface 41 indicated in Figure 5.
[0104] When a mechanical coupling component such as a set of
spokes (not shown) or a rotor 18 is used as thruster-to-shaft coupling means,
the latter is typically made out of a carbon/carbon composite material coated
at
CA 02426906 2003-04-16
26
strategic locations with a substantially oxidation resistant coating for
providing
suitable mechanical characteristics in a high temperature and oxidizer rich
operational environment. The oxidation resistant coating typically includes
silicium carbide and tetra-ethyl-ortho-silicate. It should, however, be
understood that the rotor 18 may be made out of any other suitable material
and coated with any suitable coating without departing from the scope of the
present invention.
[0105] The reinforcement wall 38 is typically made out of a one
dimensional carbon-based composite material having its carbon fibers oriented
in a predetermined direction. In one embodiment of the invention, the carbon-
based composite material may include a matrix made out of epoxy. In another
embodiment of the invention, the carbon-based composite material may include
a matrix made out of polyimide.
[0106] Alternatively, the reinforcement wail 38 may be made out of
monofilament carbon fibers or other high strength fibers windings without any
linking matrix. The carbon fibers could be wound or coiled around any suitable
structure such as external protrusions protruding from the external surface of
the thrusters 16.
[0107] As shown in Figs. 1 and 2, the engine 10 typically further
includes a casing 44. The casing 44 has a substantially cylindrical casing
inner
wall 46. Typically, the casing inner wall 46 includes two inner wall half
sections
extending substantially, axially and outwardly from a position located
respectively on each side of the rotor 18 adjacent the rotor outer peripheral
surface 22. As shown in Fig. 2, the thruster base 32 and the casing inner wall
46 are in a substantially concentric relationship relative to each other so as
to
define a casing inner wall-to-thruster base radial spacing 48 therebetween.
CA 02426906 2003-04-16
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[0108] As illustrated more specifically in Figs.1, 5 and 7, at least one
and preferably a plurality of pairs of flow guiding blades 50 extend from the
thruster base 32 towards the casing inner wall 46. The flow guiding blades 50
are configured, sized and spaced relative to each other so as to define,
together with portions of the thruster base inner surface 34 and the casing
inner
wall 46 extending therebetween, corresponding ramjet thrusters 16.
[0109] As shown more specifically in Figures 5 and 10 through 14,
each of the flow guiding blades 50 defines a radially outwardly located blade
contacting edge 62 in contact with the thruster base inner surface 34 and a
radially opposed blade spaced edge 64 positioned in a radially inwardly spaced
relationship relative to the thruster base inner surface 34.
[0110] As shown more specifically in Fig. 2, the blade spaced edge
64 is spaced from a corresponding casing inner wall 46 by a blade-to-inner
wall
running clearance 66. Similarly, each half section of the casing inner wall 46
is
spaced from a corresponding adjacent rotor side surface 20 by an inner wall-to-
rotor running clearance 68.
[0111] Referring now more specifically to figure 8, there is shown a
two-dimensional top view of a deployed ramjet channel formed by a ramjet 16.
The reference numeral 70 is used to designate a schematic representation of
an average or equivalent flow path of a fluid traveling through the ramjet
channel. Each ramjet channel is generally angled relative to the tangential
direction 72 towards the radial direction 74. Also, typically, each ramjet
channel typically forms substantially the mirror image of a substantially
deployed or stretched "S"-shaped configuration wherein the "S" shape is
substantially deployed or stretched substantially along the tangential
direction.
CA 02426906 2010-09-09
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[0112] Each ramjet channel typically extends from an inlet aperture
76 for receiving a combustible fluid to an exhaust aperture 78 for exhausting
combustion products resulting from the combustion of the combustible fluid.
Typically, the combustible fluid is a mixture of air and a combustible gas
such
as hydrogen or the like. Typically, the mixture is formed prior to being drawn
through the inlet aperture 76.
[0113] The inlet aperture 76 leads in a flow-wise direction into a
convergent inlet diffuser 80 having a flow-wise decreasing effective diffuser
cross-sectional area. The inlet diffuser 80 leads flow-wise into a combustion
chamber 82. The combustion chamber 82 leads flow-wise into a divergent
exhaust nozzle 84 having a flow-wise increasing effective nozzle cross-
sectional area. The exhaust nozzle 84, in turn, leads flow-wise to the exhaust
aperture 78.
[0114] Typically, each inlet diffuser 80 defines a substantially
rectilinear extrados 86 (a negative pressure or suction surface) and an
opposed
substantially concave intrados 88 (a positive pressure surface). The diffuser
extrados 86 and intrados 88 both merge towards each other for forming a
diffuser throat 90.
[0115] Conversely, the exhaust nozzle 84 typically defines a
substantially rectilinear intrados 92 and an opposed substantially concave
exhaust nozzle extrados 94. The exhaust nozzle intrados and extrados 92, 94
both emanate from a nozzle throat 96.
[0116] The combustion chamber 82 typically has a generally
lenticular configuration defined by a pair of substantially concave and
opposed
combustion surfaces 98 extending between the diffuser and nozzle throat 90,
96.
1930872.1
CA 02426906 2003-04-16
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[0117] The extrados 86 of the inlet diffuser 90 is typically angled
from the tangential direction 32 towards the radial direction 74 by an inlet
extrados angle 100. Similarly, the intrados 92 of the exhaust nozzle 84 is
typically angled from the tangential direction 72 towards the radial direction
74
by an exhaust intrados angle 102.
[0118] Typically, the inlet extrados angle 100 has a value of from
about 8 to 20 degrees. Preferably, the inlet extrados angle has a value of
approximately 12 degrees. Typically, the exhaust intrados angle 102 has a
value of from about 8 to about 30 degrees. Preferably, the exhaust intrados
angle 102 has a value of about 17.33 degrees.
[0119] Typically, the combustible fluid penetrates into each of the
ramjet thrusters 16 with an average or equivalent flow path 70 substantially
parallel to the extrados 86 of the inlet diffuser 80. Similarly, the
combustion
products are ejected from each of the thrusters with an average or equivalent
flow path 70 substantially parallel to the intrados 92 of the exhaust nozzle
84.
As is well-known in the art, the substantially concave intrados and extrados
88,
94 respectively of the inlet diffuser 80 and exhaust nozzle 84 are
respectively
adapted to act as compression and expansion ramps for respectively
decelerating and accelerating the combustion fluid and combustion products
flowing adjacent thereto.
[0120] Figs. 15, 16 and 17, 18 schematically illustrate the shock and
expansion pattern for two potentially limit inlet Mach numbers respectively of
1.95 and 2.5. The lower limit typically corresponds to the onset of the
rotation
startup and the upper limit to the typical maximum rotational velocity. It
should
be understood that the hereinabove mentioned upper and lower limit values are
only given by way of example for a specific fluid mixture and that other
values
are compatible with the present invention without departing from its scope.
CA 02426906 2003-04-16
[0121] Referring back to Fig. 2, there is shown that the radial
spacing 48 typically varies with axial position. The increase in the radial
spacing 48 about the outlet nozzle 84 allows for the obtention of a suitable
ratio
between the effective area of the inlet diffuser throat 90 and the effective
area
of the exhaust surface 78. Typically, the ratio of the effective area of the
diffuser throat to the effective area of the exhaust surface 78 is in the
range of
one half to allow for suitable flow characteristics. Also, as is well-known in
the
art, the combustion chamber 82 is typically provided with at least one and
preferably a set of flame-holding protrusions 166 protruding from the thruster
base inner surface 36.
[0122] In order to facilitate the flow in the inlet diffuser 80 during
startup, the inlet diffuser 80 is typically provided with a bleeding means for
bleeding a bleedable portion of the combustible fluid from the inlet diffuser
80.
Typically, the bleedable portion includes gases left over by an external
normal
shock in the un-started mode. This, in turn, will facilitate the flow through
the
inlet diffuser 80 during startup at relatively low Mach numbers.
[0123] As illustrated in Figures 5 and 10 through 14, the inlet diffuser
80 is typically an internal perforated diffuser having bleeding apertures 106
extending through the thruster base 32. The bleeding apertures 106 are
configured, positioned and sized for allowing an outflow volume of bleedable
fluid 110 therethrough. The bleeding apertures 106 are typically configured,
positioned and sized so that the outflow volume of bleedable fluid 110 is
inversely commensurable with the speed of flow of the combustible fluid in the
inlet diffuser 80.
[0124] Typically, the bleeding apertures 106 are chamfered. It
should be understood that although the bleeding apertures 106 are shown as
having a generally slot-like configuration, the bleeding apertures 106 could
CA 02426906 2003-04-16
31
have other configurations such as a generally cylindrical configuration
without
departing from the scope of the present invention.
[0125] When the inlet diffuser 80 is idled, having an upstream
normal shock, the flow through the inlet diffuser 80 is typically subsonic
with a
high pressure. In such a state, the bleeding apertures 106 bleed a relatively
large quantity of fluid. Conversely, in steady state, when the normal shock is
located just upstream relative to the diffuser throat 90, the bleeding
apertures
106 bleed a relatively small portion of bleedable fluid.
[0126] Indeed, as is well known, the fluid flowing at supersonic
speed adjacent the bleeding apertures must rotate in the latter in order to
bleed
therethrough. When the tangential speed at the surface of the bleeding
apertures 106 is relatively high, the bleeding discharge is relatively
inefficient,
hence only allowing for the bleeding of a relatively low mass of bleeding
fluid.
[0127] As illustrated more specifically in Fig, 5, bleeding flow driving
ribs 108 typically extend substantially radially from the thruster base outer
surface 36 substantially adjacent the bleeding apertures 106. The bleeding
flow guiding ribs 108 are configured, sized and position for guiding the
bleeding
flow 110 towards a thruster base intake peripheral edge 112.
[0128] As shown more specifically in Figs. 2, the casing 44 typically
also includes a substantially annular flow partitioning wall 111. The
partitioning
wall 111 is intended to separate the intake flow 113 of combustible fluid
drawn
towards the inlet apertures 76 from the outflow 110 of bleedable fluid flowing
out of the bleeding flow guiding ribs 108. The flow partitioning wall 111
typically
extends substantially axially from a position located substantially adjacent
the
thruster base intake peripheral edge 112 to a position located substantially
away from the inlet apertures 76.
CA 02426906 2003-04-16
32
[0129] When the reinforcement wall 38 is made of a composite
material, for example, the engine 10 is preferably further provided with a
temperature insulating means for thermically insulating the reinforcement wall
38 against the relatively higher temperatures generated adjacent the latter
during use. The temperature insulating means make take any suitable form
including layers of suitable insulating material (not shown).
[0130] In the embodiment shown throughout the figures, the
temperature insulating means includes a cooling means for cooling the
reinforcement wall 38 so as to maintain the structural characteristics of the
composite material matrix despite the relatively higher temperatures generated
adjacent the latter. Typically, the cooling means includes a thruster base
cooling channel extending substantially radially through the thruster base 32
for
allowing the pumping of a cooling fluid towards the thruster base outer
surface
36.
[0131] Various types of pumping means may be used for pumping
the cooling fluid towards the thruster base outer surface 36. In the
embodiment
shown throughout the figures, the thruster base cooling channel is configured
and sized for allowing centrifugal pumping therethrough of the cooling fluid
at a
sufficient cooling fluid rate through the rotation of the thruster base 32.
[0132] As shown more specifically in Figures 5 and 10 through 14,
the cooling means typically further includes cooling baffles 116 extending
substantially radially from the thruster base outer surface 36. The cooling
baffles 116 are configured so as to define outer surface cooling channels
therebetween on the thruster base outer surface 36.
[0133] The outer surface cooling channels are configured so as to
be in fluid communication with the thruster base cooling channel. The outer
CA 02426906 2003-04-16
33
surface cooling channels typically include a substantially circumferential
distributing channel 118 in communication with the thruster base cooling
channel and a plurality of auxiliary channels 120 extending at an angle
relative
to the distributing channel 118 in fluid communication therewith.
[0134] Since the cooling baffles 116 typically abuttingly contact the
reinforcement component inner surface 40, the pumping of a cooling fluid 105
at a suitable cooling fluid flow rate within the distributing and auxiliary
channels
118, 120 allows for thermal insulation of the reinforcement component 38 from
the substantially high temperatures generated in the ramjet thrusters 16
during
use. The cooling fluid 105 may take any suitable form. Typically, the cooling
fluid 105 includes ambient air pumped centrifugally during use from a location
positioned radially inwardly relative to the ramjet thrusters 16.
[0135] As shown more specifically in Figures 2, the cooling fluid 105
is typically discharged from the auxiliary channels 120 with an outflow
velocity
having at least partially an axially oriented component. Also, the auxiliary
channels 120 typically discharge the cooling fluid 105 substantially adjacent
the
casing outlet peripheral wall 56. The flow of cooling fluid 105 adjacent the
casing outlet peripheral wall 56 may hence be used to cool the latter and
prevent overheating thereof by the exhaust flow 104 of the relatively hot
products of combustion.
[0136] By way of example, in the embodiments shown in Figures 1
through 5, the engine further includes at least one, and preferably a
plurality of
rotor cooling channels 122 extending at least partially and typically fully
therethrough. The rotor cooling channels 122 define a rotor cooling channel
outlet end 124 for discharging the cooling fluid substantially adjacent the
reinforcement component 38.
CA 02426906 2003-04-16
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[0137] Typically, a rotor cooling channel 122 extends substantially
radially through each tongue component 28 and has a corresponding rotor
cooling channel outlet end 124 positioned adjacent the radially distal end of
the
corresponding tongue component 28. Each rotor cooling channel 122 also
defines a rotor cooling channel inlet end 126 in fluid communication with the
ambient air of the external environment typically adjacent the rotor
rotational
axis.
[0138] As illustrated in Fig. 1, the output shaft 12 may be provided
with a shaft cooling channel 128 extending substantially longitudinally and at
least partially therethrough. The shaft cooling channel 128 is in fluid
communication with at least one and preferably a set of radially disposed
shaft
fluid discharge apertures 130.
[0139] Also, typically, the rotor 18 has a substantially centrally
located shaft receiving aperture 132 extending therethrough. The shaft
receiving aperture 132 is configured and sized for substantially fittingly
receiving the output shaft 12. The rotor cooling channel inlet ends 126 lead
into the shaft receiving aperture 132 and are positionable in fluid
communication with the shaft fluid apertures 130.
[0140] Again, it should be understood that the rotor and shaft cooling
channels 128, 132 are only illustrated by way of example since it is expected
that such rotor and shaft cooling channels 128, 132 could create relatively
high
stress concentrations. Alternatively, as illustrated in Figure 21, the rotor
cooling
channel 122' extends only partially radially through the rotor 18 and a rotor
cooling channel inlet end 126' is positioned on either one or both of the
rotor
side surfaces substantially adjacent the rotor peripheral surface 22. When a
pair of rotor cooling channel inlet ends 126' are positioned on both of the
rotor
CA 02426906 2003-04-16
side surfaces 20 they may be symmetrically or asymmetrically disposed relative
to each other.
[0141] As illustrated in Figure 2, the output shaft 12 is typically
rotatably mounted to the casing 44 using bearings 160 supported by a pair of
supporting flanges 164 extending from a corresponding pair of casing radial
walls 162. The bearing-to-casing interface is also typically used for reacting
to
the axial thrust generated by the thrusters 16. It should be understood that
other type of shaft supporting means for rotatably supporting the output shaft
12 could be used without departing from the scope of the present invention.
[0142] In operation, at start-up, the rotation of the rotating
components is initiated through the use of an external driving means (not
shown). As shown more specifically in Fig. 2, the inlet fluid schematically
illustrated by arrows 113 is supplied to a circumferential inlet fluid supply
plenum 52 formed by a substantially annular casing outer wall 54.
[0143] The inlet fluid 113 is drawn from the casing inlet fluid supply
plenum 52 towards the ramjet inlet aperture 76. It should be noted that the
inlet
extrados 86 is angled by the inlet surface angle 100 relative to both the
tangential and axial axes 72, 74. The inlet fluid 113 penetrates through inlet
aperture 76 in a substantially parallel relationship relative to the inlet
extrados
86 and, hence, with both axial and tangential inlet fluid velocities. The
rotational speed of the thruster 16 relative to the adjacent static surfaces
is
such that the tangential inlet velocity is typically supersonic.
[0144] The supersonic ramjet inlet uses the kinetic energy inherent
in the inlet fluid mass due to the relative velocity between the ramjet inlet
and
the supplied combustible fluid to compress the latter typically via oblique
shock
waves illustrated, by way of example, at two different speeds of Mach 1.95 and
CA 02426906 2003-04-16
36
2.5 respectively in Figs. 15, 16 and 17, 18. As the combustible fuel passes
through the shock waves, it is subjected to a thermal shock sufficient to
initiate
combustion thereof.
[0145] Depending on the type of combustible fuel used, ignition
could also be assisted or performed with an ignition means (not shown) such
as electrical ignition means or the like mounted on a suitable location such
as
on the engine housing 44 or on the rotor 18. Typically, the inlet diffuser is
configured and sized so that when the inlet mixture reaches the combustion
chamber 82, the flow has decelerated to a subsonic flow speed.
[0146] Figs. 15 and 16 illustrate an example of the configuration of
the oblique shock waves within the inlet diffuser 80 at a flow speed of Mach =
1.95. The reference numeral 134 and 136 are used to designate lines of limit
flow being expanded prior to the equivalent shock wave. The reference
numeral 138 is used to indicate the line of typical flow. The reference
numerals
140, 142 and 144 are used to indicate respectively first, second and third
shock
waves having corresponding angles of 33.4, 35.5 and 48.7 degrees. The
reference numeral 146 is used to indicate an equivalent shock wave used for
design purposes extending at an angle of 51.3 degrees.
[0147] Similarly, Figs, 17 and 18 illustrate an example of the
configuration and distribution of shock waves at a flow speed Mach number of
2.5 in the inlet diffuser 80. Similar reference numerals identified with a
prime
are used to denote similar flow lines and shock waves. The angles of the
first,
second, third and equivalent shock waves are respectively 25.8, 27.2, 37.0 and
40.4 degrees.
[0148] In steady state mode, the position of the normal shock is
stabilized at the throat 96 or slightly upstream and has a value of
substantially
CA 02426906 2003-04-16
37
Mach = 1 or slightly higher. The products of combustion after discharge from
the combustion chamber 82 are expanded in the exhaust nozzle 84 and
exhausted through the exhaust aperture 78 still containing tangential kinetic
energy hence resulting in the rotational thrust of the rotor 18. The
tangential
thrust of the thrusters 16 is transmitted to the output shaft 12 by the rotor
18
while the centrifugal force created by the rotation of the thrusters 16 and
rotor
18 is compensated by the reinforcement wall 38. Typically, the speed at the
exhaust aperture is in the range of Mach = 2.15.
[0149] The products of combustion after discharge through the
exhaust aperture 78 are guided axially away from the engine 10 by an annular
casing outlet peripheral wall 56 as indicated by arrow 104 in Figure 2. The
axial thrust of the thrusters 16 is transmitted to the bearings 160. A casing
annular bulge extending circumferentially and having a substantially "U"-
shaped cross-sectional configuration is provided for protectively housing the
reinforcement wall 38.
[0150] The casing annular bulge typically includes a pair of bulge
side walls 58 extending substantially radially outwardly from the casing inlet
and outlet outer walls 54, 56 and a bulge peripheral wall 60 extending
substantially radially between the distal ends of the bulge side walls 58.
Although the configuration of the casing 44 shown throughout the figures is
substantially compact so as to favour a high power-to-volume ratio, it should
be
understood that other casing configurations could be used without departing
from the scope of the present invention.
[0151] Also, it should be understood that the hereinabove mentioned
values with regards to shock waves and speeds are only given by way of
example and that any suitable number of shock waves having any suitable
value including an isentropic state free of shock waves as well as other
suitable
CA 02426906 2003-04-16
38
flow and moving parts speed could be used or result from the present invention
without departing from its scope.
[0152] As shown schematically in Figure 19, the output shaft 12 may
optionally be coupled to a driven shaft 172 through a gear mechanism 168 or
any other suitable means.
[0153] As mentioned previously, the fuel and combustion air may be
premixed prior to feed to the ramjet inlet aperture 76. For example, fuel
injectors (not shown) may add the fuel to an inlet fluid which may be a fuel
free
oxidant containing steam or which may contain some high value fuels such as
hydrogen or some low value fuels such as biomass-produced fuel gas, sub
quality natural gas, methane or the like.
[0154] Referring now more specifically to Fig. 20, there is shown a
quad rotor engine 148 in accordance with an embodiment of the present
invention. The quad rotor engine 148 is substantially similar to the
hereinabove
disclosed rotary ramjet engine 10 and, hence, similar reference numerals will
be used to denote similar components. One of the main differences between
the quad rotor engine 148 and the rotary ramjet engine 10 resides in that the
quad rotor engine 148 uses four distinct rotor components 18 and associated
peripheral ramjets 16 allowing for a proportional specific power increase in
the
order of 25% compared to the rotary ramjet engine 10.
[0155] Also, the coupling of two rotors 18 on the same driving shaft
12 eliminates the necessity to support the rotor 18 with bearings or similar
components on each side of the rotors 18. Accordingly, the cowling or casing
44 can be made relatively lighter since it does not need to provide structural
support, merely providing a protective wall for separation from the external
environment.
CA 02426906 2003-04-16
39
[0156] Yet another advantage of coupling the pair of rotors 18
through a common driving shaft 12 is the cancellation of the parasital axial
loads by symmetry. Still furthermore, a reduction gear box 150, typically of
the
planetary type, may also be made substantially lighter since it is located at
the
centre of the common output shaft 12. Hence, it allows the use of a compact
structural housing 44' integrating all of the rotating components of the
reduction
gear boxes 150 and rotor shafts 12. Such a housing 44' may provide a weight
reduction in the order of 80% compared to the housing 44 of the one-rotor
rotary ramjet engine 10.
[0157] Still furthermore, the vectors of the angular momentum of the
rotors 18 could cancel each other out by having output shafts 12 coupled to
opposite rotors 18 rotate in opposite direction relative to each other.
Accordingly, such an engine would be less subjected to gyroscopic effects
linked to the angular momentum change during an angular manoeuvre.
[0158] Referring now more specifically to Figures 22 and 23, there is
shown a rotary ramjet engine 10' in accordance with yet another alternative
embodiment of the invention. The engine 10' is substantially similar to the
engine 10 and, hence, similar reference numerals will be used to denote
similar
components. Two of the main differences between the engines 10 and 10'
resides in that the guiding blades of the engine 10' extend substantially
radially
from the rotor outer peripheral edge 22 in a direction leading radially
outwardly
instead of extending from the thruster base 32 in a direction leading radially
inwardly.
[0159] Consequently, pressure values at the guiding blade interface
with its supporting structure may be higher. Also, since the flow guiding
blades
50 must be supported along their entire length by the rotor outer peripheral
surface 22 and since the rotor 18 must compensate for centrifugal forces
CA 02426906 2003-04-16
generated on the rotor 18 and the flow guiding blades 50 during rotation
thereof, the size and mass of the rotor 18 part of the engine 10' are
substantially greater then that of the engine 10.
[0160] Although the specific power of the engine 10' is expected to
be substantially less impressive then that of the engine 10, the engine 10'
nevertheless is usefull, for example, in the making of more easily available
prototype models for studying the fluid dynamics of the thrusters 16. It is
indeed expected that the fluid dynamics of the thrusters 16 will remain
substantially similar for both the engines 10 and 10'.
[0161] Another main difference between the engines 10 and 10'
resides in that the engine 10' is provided with a variable geometry inlet
diffuser
80' having a sleeve valve 152 including a movable diffuser ring 154 instead of
the perforated inlet diffuser 80 of the engine 10 having bleeding apertures
106.
As shown more specifically in Figure 23, the sleeve valve 152 also includes a
circumferential valve seat 156 extending between the inlet diffuser 80 and a
circumferential bleeding channel 158.
[0162] As indicated by arrows 160, the diffuser ring 154 is movable
axially using a suitable diffuser ring moving means (not shown) for
selectively
restraining or blocking the flow of bleeding fluid from the inlet diffuser
across
the valve seat 156 and into the bleeding channel 158. Apart from functional
differences resulting inherently from the above-mentionned structural
differences, the operation of the engine 10' is substantially similar to that
of the
engine 10 and, hence, will not be further disclosed.
[0163] The present invention also relates to a method for optimizing
the performance of a rotary ramjet engine as hereinabove disclosed for a
particular application and size. The method involves estimating a maximum
CA 02426906 2003-04-16
41
operational rotational speed taking into consideration the limits set by the
strength of the materials as a function of both the geometric characteristics
of
the components and the mechanical properties of the materials used for
manufacturing such components.
[0164] The method also involves determining an inlet Mach number
for a predetermined efficiency range (the Mach number being defined as the
ratio of flow speed at the inlet, measured in the reference frame of the
rotor,
over the speed of sound of the inlet gas).
[0165] Once the maximum operational rotational speed and the
Mach number have been determined taking into consideration respectively the
strength of materials and a target efficiency, the desired speed of sound of
the
inlet fluid is determined. The method further involves controlling the speed
of
sound of the inlet fluid by controlling the nature and relative proportion of
the
fuel and oxidizer.
[0166] Indeed, the engine 10 may be operated at various Mach
numbers depending on the type of fuel used and on other variables such as the
use of a premixed fuel-air mixture as opposed to a fuel mixed with a fluid
such
as air within the engine. For example, when pre-mixed hydrogen is used
together with a perforated inlet diffuser 80 the operational inlet limit Mach
number may vary between 1.86 and 2.6. With a variable diffuser the
operational range may vary between Mach I and 2.6.
[0167] With a pre-mixed stoichiometric mixture containing gaseous
octane upper limits may reach values in the neighborhood of Mach 4 or 5 since
the speed of sound of the mixture is substantially reduced. The use of octane
as fuel would also allow for a substantial reduction in the rotational speed
of the
engine. This reduction in rotational speed could, in turn, allow for the use
of
CA 02426906 2003-04-16
42
more conventional materials such as metals or ceramics for building prototypes
or actual models.
[0168] Although the present invention has been described
hereinabove by way of preferred embodiments thereof, it can be modified,
without departing from the spirit and nature of the subject invention as
defined
in the appended claims.
