Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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COMPACT COMPOUND ENGINE PACKAGE
BACKGROUND OF THE INVENTION
Field of the Invention
The present invention relates to compound cycle engines and, more
particularly, to a compact compound engine pacl~age suitable for aircraft
applications.
Description of the Prior Art
There have been attempts to developed compound cycle engines
having internal combustion engines and turbine engines, coupled together to
provide
a common output. For example, see U.S. Patent No. 4,815,282. However, to date,
proposed compound cycle engine designs have been bullcy and therefore failed
to
detail a complete solution to the integration of a cycle topping device, such
as a rotary
combustion engine, with a gas turbine in a compact pacl~aging suitable for
aero
applications, such as aviation.
Moreover, prior art compound cycle engine designs have been wear in
providing solutions to the cooling of internal combustion engines which are
practical
and realistically viable in an aircraft environment.
SUMMARY OF THE INVENTION
It is therefore an aim of the present invention to integrate a cycle
topping device into a compact compound engine pacl~age.
It is also an aim of the present invention to provide a compact self
cooling system for a rotary combustion device.
In one aspect, this disclosure covers the integration of a rotary topping
device into a turbo-compounded pacl~age using a bevel drive to facilitate
compact
pacl~aging particularly suitable for aircraft use. A few embodiments of the
pacl~aging
are disclosed for different aircraft applications, including turboprop,
turboshaft and
auxiliary power unit (APU).
Therefore, in accordance with a general aspect of the present
invention, there is provided a compound cycle engine comprising a compressor
and a
turbine section, said turbine section having a turbine shaft, and at least one
cycle
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topping device providing an input to said turbine section and cooperating
therewith to
provide shaft horsepower, said at least one cycle topping device having an
output
shaft extending at an angle to said turbine shaft, and wherein a bevel gearbox
mechanically linlcs the turbine shaft and the cycle topping device output
shaft
together.
In accordance with a further general aspect of the present invention,
there is provided an aircraft engine comprising a compressor section and a
turbine
section turbocompounded with at least one cycle topping device oriented at an
angle
with respect thereto, the turbine section having a turbine shaft mechanically
linked to
an output shaft of the cycle topping device through bevel gearing to provide a
common output.
In accordance with a still further general aspect of the present
invention, there is provided a compound cycle engine comprising a compressor
and a
turbine section, topping means for providing an energy input to said turbine
section to
permit operation thereof, said topping Illeans being oriented at an angle to
said
turbine section, and bevel gearing for mechanically linlcing said topping
means and
said turbine section in order to provide a common output to drive a load.
In accordance with a still further general aspect of the present
invention, there is provided a compound cycle engine comprising a compressor
and a
turbine section, at least one cycle topping device providing an input to said
turbine
section and cooperating therewith to provide shaft horsepower, said at least
one cycle
topping device being fed with pressurized air from said compressor section,
and
wherein a valve is provided for selectively bypassing said cycle topping
device.
In accordance with a still further general aspect of the present
invention, there is provided a compound cycle engine comprising a compressor
and a
turbine section, at least one cycle topping device providing an input to said
turbine
section and cooperating therewith to provide shaft horsepower, and gas turbine
engine accessories, wherein said at least one cycle topping device drives said
gas
turbine engine accessories.
BRIEF DESCRIPTION OF THE DRAWINGS
Reference will now be made to the accompanying drawings, in which:
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Fig. 1 is a schematic diagram showing the integration of the gas
turbine engine and rotary machine using bevel drive and direct drive blower.;
Fig. 2 is a functional schematic diagram of the compound cycle engine
illustrating the air flow path through the gas turbine and the rotary topping
device;
S Fig. 3 is a conceptual isometric view of a turbo compound engine
package;
Fig. 4 is a conceptual isometric view of the engine of Figure 3,
showing the outer air cooling ducts in place;
Figs. 5a and Sb shows the device of Figure 4 in both turboshaft and
turboprop installations; and
Fig. 6 is a schematic cross-sectional side view of a rotary topping
device with integrated toroidal cooler and cooling fan.
DESCRIPTION ~F THE PREFERhED EMEODIMENTS
Fig. 1 is a schematic representation of a compound cycle engine 10 of
a type preferably provided for use in a variety of aero applications, such as
turboshaft,
turboprop or APU (auxiliary p~wer unit) applications. l2efernng to Figure l,
it can be
seen that the compound cycle engine 10 generally comprises at least one rotary
cycle
turbine topping device (TTD) 12 (preferably 1 or 2, as indicated in Fig. 1)
and a gas
turbine engine 14, which acts as a turbocharger. Turbocharger 14 comprises a
compressor 19, a first stage turbine 20 and a second stage or power turbine
16. A
hollow shaft 22 connects first stage turbine 20 to compressor 19. The power
turbine
16 is preferably a free turbine and includes a power turbine shaft 28
concentrically
disposed within the hollow shaft 22 for independent rotation with respect
thereto. The
shaft 28 connects power turbine 16 to rotary machine 12 via a bevel gearset
34.
Preferably the bevel gear set 34 has a reduction gear ratio of 3:1. It is
understood that
the gear ratio could however be any desired gear ratio. Compressor 19
communicates
with an air intake 35 and a compressor scroll 24, the compressor scroll 24
leading to
aal inlet 37 of the rotary cycle topping device 12. The compressor scroll 24
preferably
consists of a split scroll (2 * 180 deg half scrolls), as indicated in Fig. 2.
An air outlet
39 of rotary machine 12 communicates with an exhaust duct 41 to a turbine
volute 30
leading to the turbines 20 and 16. A wastegate 32 selectively connects
compressor
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scroll 24 and turbine volute 30 in fluid flow communication. The. wastegate 32
preferably includes a selectively openable blow-off valve. As best shown in
Fig. 2,
the waste gate or blow-off valve 32 selectively allows the compressor
discharge to
bypass the rotary cycle topping device 12 and to "blow off' directly into the
turbine
volute 30 in order to prevent surge at low rotary cycle topping device speed.
As shown in Fig. 1, the rotary cycle topping device 12 has an output
shaft 26. To facilitate neat and compact packaging suitable for aircraft use,
it is
herein proposed to set the output shaft 26 of the rotary cycle topping device
12 at an
angle B of about 90 degrees, and preferably less, to the power turbine shaft
28. The
angle ~ is preferably 45 degrees or greater. The bevel gearset 34 is used to
mechanically linlc the rotary topping device output shaft 26 and the power
turbine
output shaft 28 together. The use of the bevel gearset 34 advantageously
provides for
very short ducting from the compressor 19 to the rotary cycle topping device
12 and
the rotary cycle topping device 12 to the compressor turbine 20 and the power
turbine
16 while at the same time providing a compact transmission. The compressor
exit
and turbine entry ducting 24 and 30 is hot, heavy and expensive and, thus, is
preferably as short as possible. The length of the ducting should also be
minimal in
order to minimize heat and pressure losses which negatively affect the overall
engine
efficiency.
~y so orienting the rotary cycle topping device 12 with respect to the
power turbine shaft 28 and by using a bevel gearset, the envelope and frontal
area of
the engine 10 can also be minimized. This can be readily appreciated from Fig.
3
which shows a pair of rotary cycle topping devices 12 installed at the front
of the
engine 10 on opposed sides of a gearbox 17, the rotary cycle topping devices
12
being oriented at approximately 90 degrees to the main engine axis. The
resulting
paclcage very much resembles commercially available turboprops and turboshaft
gas
turbine applications and can conceivably be installed in existing aircraft
nacelles or
engine bays.
Another advantage provided by the above compact packaging
configuration is that cooling air can be drawn along the main cycle air from a
single
inlet 35 (Fig. 3) at the front of the engine 10.
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While the rotary cycle topping devices) could be placed in parallel
with the compressor turbine rotor with a relatively short ducting to the
compressor
and turbine, a potentially heavy idler geax train would be needed. Also the
resultant
frontal area would be high and not so suitable for aero engine installation.
Inline placement of the rotary topping devices) tends to lead to long
ducts from either the turbine or compressor to the rotary topping devices) as
well as
potentially requiring long installation. The above-described used of a bevel
gearset to
mechanically lined the power turbine shaft 28 to the output shaft 26 of the
rotary
cycle topping device 12 is, thus, advantageous as compared to the other
contemplated
alternatives.
More specifically, as shown in Fig. 1, the bevel gearset 34 generally
comprises a first bevel gear 36 rigidly mowzted to the rotary cycle topping
device
output shaft 26 for meslung engagement with a second bevel gear 38 provided at
the
front end of the power turbine shaft 28. A third gear 40 is rigidly mounted at
the
distal end of the rotary cycle topping device output shaft 26 for meshing
engagement
with a fourth bevel gear 42 provided on a shaft 18, which is connected to a
load, such
as a propeller (Fig. 5b), a generator, a tachometer, a helicopter rotor (Fig.
5a), a
starter (Fig. 3), an oil pump (Fig. 3), a fuel pmnp (Fig. 3), a cooling fan,
and a load
compressor. Accordingly, the shaft 18 is directly drivingly connected to the
rotary
cycle topping device output shaft 26 and indirectly drivingly connected to the
power
turbine shaft 28 through the rotary topping device output shaft 26. The
outputs of the
rotary cycle topping device 12 and power turbine 16 are thus linl~ed
mechanically to
drive the shaft 18. They both cooperate to provide the shaft horsepower
required to
drive the load coupled to the shaft 18. At engine start-up, the rotary cycle
topping
device 12 does most of the worlc, whereas under normal operating conditions,
the
power turbine 16 contributes significantly to the total power output on the
shaft 18.
It is understood that the gearset 34 does not need to be a double gearset
and that any gearset that permits coupling of two non-parallel shafts could be
used as
well. All shafts 18, 22, 26 and 28 have suitable bearings 51.
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As shown in Fig. l, rotary engine shaft 26 is also connected to a fan or
blower 44 having a fan air inlet 46 and a fan air outlet 47 communicating via
ducting
49 to an oil cooler 50.
The rotary cycle topping device 12 may be of any suitable design, such
as those disclosed in US5,471,834, US5,522,356, US5,524,587 and US5,692,372,
to
name a few, though there are certainly others available as well, as will be
understood
by the slcilled reader. The contents of all of these documents are hereby
incorporated
into this disclosure by reference. It is noted that the cycle topping device
does not
necessarily have to be an internal combustion engine, the only requirement
being that
it produces the input (i.e. hot stream of gas) needed for the turbines to
operate. For
instance, a wave rotor engine coupled to a combustor could potentially be used
for
topping or providing an energy input to the gas turbine cycle.
As shown in Fig. 2, the rotary cycle topping device 12 preferably
generates rotary movement through a sliding vane rotor 29 to drive the output
shaft
26. A fuel-air mixer 27 is provided in the ducting between the compressor
scroll 24
and the rotary cycle topping device inlet 39 to inject fuel in the compressed
air before
it flows into the rotary cycle topping device 12. A low speed enrichment
throttling
valve 31 is provided in the ducting just upstream of the fuel-air mixer 27 to
adjust the
quantity of air entering into the fuel and air mixture. It can be readily
appreciated
from Fig. 2, that the gas generator (i.e. the compressor 19 and the compressor
turbine
20) does not drive any accessories, its main function being to turbocharge the
rotary
cycle topping device 12. It is the rotary cycle topping device 12 and the
power turbine
16 that provides the require shaft horsepower to chive the accessories. via a
gearbox
17 (Fig. 3). The shaft 26 act as a power tale-off shaft for driving the
accessories. The
term "accessories " is herein intended to generally refer to gas turbine
components
that need to be driven but which does not provides any propulsive forces. For
instance, the accessories could talce the form of a fuel pump, an oil pump, an
air
pump, a starter, a tachometer, a generator and a load compressor.
Referring now to Figures Sa and Sb, shown is the compound cycle
engine 10 in turboshaft and turboprop installations, respectively. In Fig. 5a,
the
engine 10 is used to drive a helicopter rotor 53. In Fig. 5b, the engine
drives a
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propeller 55. In the turboshaft application, the air intake is located at the
top of the
engine 10, whereas in the turboprop application, the air intake 35 is located
on the
front side of the engine 10.
In use, incoming air flowing through the air intake 35 is compressed
by the compressor 19 and directed to the inlet 39 of the cycle topping device
12 via
compressor scroll 24. Fuel is introduced into the compressed air flow
immediately
prior to its entry into the rotary cycle topping device 12 by lcnown means as
schematically depicted at 27 in Fig. 2. The low speed enrichment throttling
valve 31
(Fig.2) adjusts the quantity of air entering into the fuel and air mixture.
The fuel/air
mixture is then further compressed by the rotary motion of the rotor 29 before
being
ignited. The resultant combustion gases are then expanded to drive the rotor
29 and,
thus, the shaft 26, before being exhausted. The combustion gases are directed
into the
compressor turbine 20 and the power turbine 16 via the turbine volute 30. The
compressor turbine 20 and the power turbine 16 extract energy from the
expanding
combustion gases, converting the energy into shaft horsepower to respectively
drive
the compressor 19 and the shaft 18 as well as other accessories. In use, shaft
22
typically rotates at about 60000-70,000 rpm while shaft 28 rotates at about
50,000
rpm. The bevel gearset 34 will provide a reduction of about 3:1. While output
bevel
gearset 34 will provide an output shaft speed as required, such as 6,000 rpm
for a
turboshaft or 2,000 rpm for a turboprop. The rotary machine will have a
rotational
speed of about 15,000 rpm. The above speeds are given for exemplary purposes
only
and are thus not intended to be exclusive.
In operation, the blow-off valves 32 are typically opened where there
is a mismatch between the flow capacitors of the rotary cycle topping device
12 and
the turbocharger 14 such as might occur at part speed operating conditions.
The above-described combined cycle engine offers high thermal
efficiency because of high cycle pressure ratio and temperature provided by
the
closed volume combustion of the rotary cycle topping device 12. The combined
cycle
also provides for the reduction of the size and the weight of the
turbomachinery as
compared to a conventional single cycle gas turbine engine at the same
horsepower
shaft or thrust because of the increased power per unit mass airflow.
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Furthermore, the integration of a rotary combustion device (i.e. the
preferred embodiment of the cycle topping device 12) into a gas turbine engine
is
significantly advantageous in term of fuel efficiency particularly when
operating at
reduced power.
The rotary combustion cycle topping devices) 12 generates) heat
during operation which must be dissipated in order to prevent overheating
thereof.
Cooling requirements of such rotary internal combustion engines can be lugher
than
gas turbine engines and therefore achieving very compact arrangements for
cooling
are important to making a practical device for aviation and automotive
applications.
As shown in Figs. 1 and 6, a compact self cooling system for the rotary
cycle topping device 12 can be achieved by integrating the axial blower or
cooling
fan 44 directly on the output shaft 26 of the rotary cycle topping device 12.
This is
made advantageous by the relatively high output rpm of the rotary cycle
topping
device 12 (about 16000 rpm) which makes a high flow compact fan practical.
There
is n~ need f~r intermediate gears, chains or pulleys for driving the fan 449
as the fan
44 is directly mounted on the rotary topping device output shaft 26, thereby
providing
for a very compact cooling arrangement.
As shown in Fig. 4, the direct drive cooling fan 44 draws ambient air
through air cooling ducts 46 via cooling air intakes 48. The cooling air
intakes 4~ are
1~cated at the front of the engine 10 at a higher elevation or outboard
position than
the engine air intalLe 35 and laterally with respect thereto. The compressor
intake 35
corresponds to the lowest flow section. This advantageously provides for more
direct
flows of cooling air, which are much more difficult to design for. The ducting
46 to
the fan 44 can advantageously be very short from the air intalces 4.~ due to
the 90
degrees rotary topping device placement. The cooling air is exhausted
laterally
through an exhaust port 59 of each side of the engine 10. An optional duct aft
61 can
be connected to the exhaust port 59 to discharge the cooling air axially along
the
sides of the engine 10.
The air cooling ducts 46 channel the air through a heat exchanger or an
oil cooler 50 (Figs. 1, 2 , 3 and6) to piclc up the excess heat absorbed by
oil or other
coolant as it is circulated by a pump 51 (Fig.6) through cooling passages 52
defined
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in the rotary topping device casing, as shown in Fig. 2. As shown in Fig. 6,
the pump
51 is mounted in a closed loop circuit 55 with the device 12 and the oil
cooler 50 to
ensure a continuous re-circulation flow of oil. As the oil travels through the
oil cooler
50, it gives off heat to the forced air passing through the .oil cooler 50.
Then, the so
cooled oil is re-circulated through the cooling passages 52 in the rotary
topping
device casing to extract excess heat therefrom. The oil cooling is about 5% of
the fuel
input.
As shown in Fig.6, the oil cooler 50 is preferably provided in one
embodiment in the form of a toroidal oil cooler surrounding the fan inlet on a
downstream side of the rotary topping device 12. The toroidal oil cooler 50 is
integrated to the rotary topping device 12 and extends rearwardly therefrom.
The
toroidal cooler 50 is concentrically mounted about the output shaft 26 and
located
radially outwardly of the device 12. The toroidal oil cooler 50 provides a
toroidal
cooling path for the oil circulated by the pump 51, which is conveniently
driven from
the reduction gearbox 17. The fan 44 draws air radially inwardly through the
toroidal
oil cooler 50. The toroidal cooler 50 defines air exliaust passages 53
defining a bend
from radial to axial. The hot air leaving the toroidal cooler 50 is rejected
axially
rearvaardly of the fan 44. As illustrated in Fig. 6, the fan 44 is used mainly
for the
purpose of providing forced air through the oil cooler 50 in order to improve
cooling
efficiency. fIowever, suction air from the fan inlet or delivery air may also
be used to
cool the core of the rotary topping device 12 by causing air to flow through
axially
extending passages 54 defined through the rotary topping device 12.
The embodiments of the invention described above are intended to be
exemplary. Those slcilled in the art will therefore appreciate that the
forgoing
description is illustrative only, and that various alternatives and
modifications can be
devised without departing from the spirit of the present invention. For
example, the
compressor and turbine configuration shown is only one of many possibilities.
The
ducting arrangement between successive components need not be exactly as
shown,
nor does the relative arrangement of components. Though the description refers
generally refers to one rotary machine, it will be, understood that one or
more could
be used in parallel or series. Accordingly, the present is intended to embrace
all such
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alternatives, modifications and variances which fall within the scope of the
appended
claims.
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