Note: Descriptions are shown in the official language in which they were submitted.
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LOW VOLUMETRIC COMPRESSION RATIO INTEGRATED
TURBO-COMPOUND ROTARY ENGINE
FIELD OF THE INVENTION
The present invention relates to gas turbine and rotary engines and,
in particular, to turbo-compounded rotary engines or turbo-compounded internal
combustion engine.
BACKGROUND OF THE INVENTION
Topping of the gas turbine engine cycle is well-known in the art.
US 4,815,282, US5,471,834 and US5,692,372, for example, show the prior
attempts at integrating gas turbine with cycle-topping devices, such as piston-
type
internal combustion engines and eccentric rotary engines such as the so-called
Wau~el engine. Such cycle topping devices promise much-improved fuel
efficiency for the integrated engine. All of the integrated engines disclosed
in the
above mentioned patents require an intercooler to cool the air before it
enters the
compressor section of the engine. Such intercooler are lmow to be bulky,
heavy,
etc. and, thus, not ideal for airborne applications.
For gas turbines destined for airborne applications, integration
must not only successfully address nnprovements in cycle efficiency9 but also
provide a compact and lightweight package, and preferably one which does not
significantly alter the envelope required versus that of a regular (i.e. non-
compounded) gas turbine engine. Prior art attempts have not been as successful
in these areas, and hence there exists a need for improved compact devices
which
offer not only improved efficiency, but also better power density,
reliability,
operability and so on.
Various types of cycle topping devices are lmown, including both
non-rotating and rotating types. The present application is particularly
concerned
with eccentric rotary machines of all types useful in providing cycle-topping
benefits to a gas turbine engine. Exaanples are shown 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 well-understood by the spilled
reader.
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SUMMARY OF THE INVENTION
It is an aim of the present invention to provide a compound cycle
engine better suited for airborne applications than the prior art.
One general aspect of the present invention covers an integrated
cycle topping device and gas turbine engine (the "integrated engine") designed
for
low voliunetric compression ratio (<3.5) which allows pre-mixed fuel upstream
of the cycle topping device without the need of an inter-cooler. It provides
for
improved thermal efficiency and improved specific power.
In accordance with a further general aspect of the present
invention, there is provided a compound cycle engine comprising a compressor
and a turbine section, and at least one rotary engine providing an energy
input to
said turbine section, wherein said at least one rotary engine is mechanically
linked
to said turbine section to provide a common power output.
In accordance with am~ther general aspect of the present invention,
there is provided a compound cycle engine comprising a compressor section, a
rotary engine section and a turbine section in serial flow communication with
one
another, and a primary output shaft providing the primary power output of the
engine, wherein the rotary engine section and the turbine section are both
drivingly connected to the primary output shaft.
In accordance with another general aspect of the present invention
there is provided a method of providing a non-intercooled cycle for a compound
cycle engine, the engine having a rotary engine and a gas turbine connected in
series, the method comprising the steps of a) compressing air in a compressor
section of the gas turbine, b) further compressing the air in the rotary
engine,
wherein the volumetric compression ratio in the rotary engine is below 3.5,c)
mixing fuel with the compressed air to obtain an air/fuel mixture, d)
combusting
the air/fuel mixture, e) extracting energy from the combusted air/fuel mixture
through expansion in the rotary engine, and f) further extracting energy from
the
combusted air/fuel mixture using a turbine section of the gas turbine.
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In accordance with another general aspect of the present invention,
there is provided a compound cycle engine comprising a compressor and a
turbine section, and at least one cycle topping device providing an energy
input to
said turbine section, said compressor section compressing the air according to
a
pressure ratio PRgt, said at least one cycle topping device further
compressing the
air according to a volumetric compression ratio R,,~, and wherein PRgt x
R~~l'3 <
30.
. hl accordance with a sill further general aspect of the present
invention, there is provided a method of providing a non-intercooled cycle for
a
compound cycle engine, the engine including a cycle topping device and a gas
turbine connected in series, the method comprising the steps of:
a) compressing air in a compressor section of the gas turbine using a pressure
ratio PRgt, b) further compressing the air in the cycle topping device using a
volumetric compression ratio R"~, c) mixing fuel with the compressed air to
obtain an air/fuel mixture, d) combusting the air/fuel mixture, e) extracting
energy from the con busted air/fuel mixture through expansion in the topping
device, and f) further extracting energy from the combusted air/fuel mixture
using
a turbine section of the gas turbine, wherein the relationship between PRgt
and RVc
is maintained such that PRgt x R"Ci.s < 30.
In accordance with a 5ti11 further general aspect of the present
invention, there is provided a method of providing a cycle for a compound
cycle
engine, the engine including a rotary engine and a gas turbine connected in
series,
the method comprising the steps of: a) determining an auto-ignition limit of a
fuel/air mixture; b) determining a pressure ratio associated with the auto-
ignition
limit; c) determining respective pressure ratios for a compressor section of
the gas
turbine and for the rotary engine; d) and selecting a combination of pressure
ratios
for the compressor section and the rotary engines, which provides an overall
pressure ratio inferior to the pressure ratio determined in step b).
It is understood that the term "cycle topping device", as used
throughout this application and the attached claims, applies to any device
adapted
to provide an input to the turbine cycle, and not just rotary cycle topping
devices
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such as a Wankel engine, sliding or pimled vane rotary machine (such as those
disclosed in US5,524,587 or US5,522,356, respectively). Also, the term
"compound cycle engine" as used throughout this application and the attached
claims is intended to refer to an engine wherein at least two different types
of
engine (e.g. rotary engine and gas turbine, etc.) are integrated together to
provide
a common output. Further, the term "rotary engine", as is used in the art and
as is
used herein, is used to refer to an engine in which gas compression and
expansion
occur in a rotary direction, rather than in a reciprocating manner such as in
a
piston-style internal combustion engine.
BRIEF DESCRIPTION OF THE DRAWINGS
Reference is now made to the accompanying Figures depicting
aspects of the present invention, in which:
Fig. 1-3 are schematic diagrams of single shaft embodiments of an
integrated engine comprising a gas turbine engine turbo-compounded by a rotary
cycle topping device;
Fig. 4 is a Temperature-Entropy diagram of a turbo-compounded
rotary engine cycle;
Fig. 5 is a Thermal Efficiency-Overall Pressure Ratio diagram
illustrating the sensitivity of an intercooler thermal efficiency vs. the
rotary
engine volumetric ratio and the gas turbine pressure ratio;
Fig. 6 is a Combustion Inlet Temperature vs. Combustion Inlet
Pressw-e diagram illustrating the sensitivity to auto-ignition vs. rotary
engine
volumetric ratio and gas turbine pressure ratio;
Fig. 7 is a schematic diagram of a free turbine embodiment of an
integrated engine.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
Integrated engine embodiments are shown in Figs. 1-3 for single
shaft concepts where one (1) or two (2) closed volume combustion rotary
engines
can be coupled to a power turbine via a gearbox. Fig. 1 shows an integrated
engine or compound cycle engine wherein the rotary engines are mounted at 90
degrees to the main engine axis. Fig. 2 shows another possible configuration
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wherein the rotary engines are mounted parallel to the main engine axis. Fig.
3
shows a rotary engine mounted in-line with the main engine axis.
Referring now more particularly to Fig. 1, there is disclosed a
single shaft engine 10 which includes an AGB/RGB 12 (accessory
gearbox/reduction gearbox), a compressor 14, two rotary machines or engines 16
and a power turbine 18 connected on a single shaft 20. The turbine shown is a
radial turbine, though other configurations are possible. The rotary engines
16 axe
connected to the shaft 20 by separate tower shafts 22 and 24. The compressor
14
is preferably a centrifugal compressor, though need not necessarily be so, and
is
fed by an intake 26. The compressor 14 communicates with the rotary engines 16
via an inlet scroll 28, and the rotary engines 16 in turn communicates with
the
power turbine 18 via an outlet scroll 30, to thereby provide a continuous gas
path
between compressor intake 26 and turbine exhaust 27, as will be understood by
the skilled reader. The compressor 14 acts as a turbocharger to the rotary
engines
16. A fuel pre-mixer 32 is integrated to the inlet scroll 28 of each rotary
engine.
As shown in Fig. 1, the shaft 20 1S c~llJolntly drl~ell by the power
turbine 18 and the rotary engines 16. The rotary engine output shafts 22 and
24
can be mechaalically linked to the shaft 20 by means of bevel gearing 34.
Each rotary engine 16 includes a housing 23 which is liquid-
cooled in a suitable manner, and having an associated cooling inlet 25 and
outlet
27. The cooling liquid, fOr lllStallce o11, is circulated through the rotary
engine
housing 23. As the liquid travels through or over the housing 23, it picks up
excess heat. The liquid is then pumped to a liquid cooler (not shown) where
the
liquid is cooled before being re-circulated back into the rotary engines 16.
As can be readily appreciated from Fig. 1, in use ambient air
entering the gas turbine intalce 26 is compressed by the compressor 14, then
it is
routed to the pre-mixers) 32 where fuel is premixed with the air. The fuel/air
mixture then enters the rotary engines 16, gets further compressed with volume
reduction. The compressed mixture is then igiuted in the rotary engines,
according to known techniques, before being expanded, the energy of such
expansion further driving the rotary engine. The rotary engine exhaust gases
are
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then ducted to the power turbine 18 for powering the turbine to produce
further
work before exhausting to the atmosphere via the turbine exhaust 27.
The power developed by the rotary engines 16 and the power
turbine 18 is used to drive a common load via the AGB/RGB 12. As will be
appreciated by the skilled reader, and is shown in with respect to the
embodiment
of Fig. 7, the load can take the form of a propeller, a helicopter rotor, load
compressor or an electric generator depending whether the engine is a
turboprop,
a turboshaft or an APU (Auxiliary Power Unit).
Figs. 2 and 3 respectively show other embodiments of a single
shaft engine wherein like components are identified by like reference
numerals. A
duplicate description of these components is herein omitted for brevity, as
the
slcilled reader does not require such to understand the concepts disclosed.
The embodiment shown in Fig. 2 essentially differs from the
embodiment shown in Fig. 1 in that the rotary engines 16 are mounted parallel
to
the main engine axis. The output shafts 22 and 24 of the rotary engines 16 are
mechanically linked to the power turbine shaft 20 through the AGB/RGB 12.
As can be clearly seen in Fig. 3, the single shaft engine 10 can also
be configured so that a single rotary engine 16 is mounted in-line with the
power
turbine shaft 20. According to this reverse-flow configuration, the turbine
shaft
20 is drivingly corrected to the AGB/RGB 12 through the rotary engine output
shaft 20. Gearing (not shown) is provided to mechanically connect the power
turbine shaft 20 to the rotary engine output shaft 22.
As can be seen from Figs. 1-3, the rotary engines) can be mounted
such that their shaft axes are either parallel or perpendicular to the gas
turbine
shaft axis.
Fig. 7 shows a flee turbine embodiment where the rotary engine 16
(which can be either one or two rotary, or more, rotary engines, but referred
to
here in the singular for convenience) is coupled to the power turbine 18 only.
The compressor 14 is mounted on a separate shaft 15 and is independently
driven
by a compressor turbine 17 coaxially mounted on the shaft 15. The compressor
14
and the compressor turbine 17 act as a turbocharger to the rotary engine 16.
The
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outputs of the rotary engine 16 and power turbine 18 are linked mechanically
through the AGB/RGB 12 to drive a common load (for instance a helicopter
rotor., a propeller or a generator). The AGB/RGB provides the required speed
reduction (if any, as desired) to permit coupling of the high speed power
turbine
18 to the slower rotary engine 16. The power turbine 18 and the rotary engine
16
both cooperate to provide the shaft horsepower required to drive the load
coupled
to the AGB/RGB 12. This flee turbine configuration is advantageous in that it
provides the ability to have a high speed turbomaclune section (more compact
and efficient) since it is not directly mechanically coupled to the slower
rotary
engine. A smaller stauter 39 can also be used on the free turbine
configuration as
the starter 39 can be provided on the output RGB (see Fig. 7) rather than
having
to drive the entire compound machine.
A cooling fan 34 is preferably drivingly comlected to the rotary
engine output shaft 22 to push cooling air through via appropriate ducting 36
to
provide cooling air to the air cooled rotor 31 of the rotary engine. The
cooling air
is then expelled from the rotor to cool the cavity 35 between the compressor
14
and the hot scroll 30. The machine housing 23 is cooled with suitable cooling
liquid circulated through a suitable liquid conduit or housing jacket 37,
extending
between the cooling inlet and outlet 25 and 27, to thereby also extract excess
heat
from the housing of rotary engine 16.
As is apparent from Figs. 1-3 and 7, the disclosed embodiments do
not include an intercooler between the gas. turbine compressor and the rotary
engines. The prior art required an intercooler (see for example, U.S. Fatents
Nos.
4,815,282 and 5,4.71,834) to cool the air before it enters the rotary machine
in
order to prevent pre-ignition of the fuel/air mixture, as the skilled reader
will
recognize that as a fuel/air mixture is increasingly compressed, in becomes
susceptible to igniting. The embodiments of Figs. 1-3 and 7 were not possible
in
the prior art, but are now possible through use of the cycle improvements
according to another aspect of the present invention, as will now be
described.
~ Figs. 4 and 5 illustrate the high efficiency and specific power of
the non-intercooled cycle. The results shown in Fig. 4 are for a constant
volume
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combustion (CVC) rotary engine having a volumetric expansion pressure ratio
(Rve) twice its volumetric compression ratio (R,,~), with no intercooler and a
temperature T4 at the exit of the rotary engines 16 set at 3100 °F, the
rotary engine
being used with a gas turbine engine having a compressor pressure ratio (PR-
GT)
of 6. The temperature-entropy relations were obtained for five different
values of
volumetric compression ratio (RV~ 1.2, R,,~ 1.5, R~~ 2.0, R~~ =3.0, and R~~
=5).
Fig. 4 also shows the value of the ratio r~th/SHP/W1 (r~th:. thermal
efficiency;
SHP: shaft horse power; W1: airflow at the compressor intake) at the peals
temperature of each curve.
The results in Fig. 5 are also for a constant volume combustion
rotary engine with a peak temperature T4 of 3100 °F, the rotary engine
having a
volumetric expansion pressure ratio (Rve) twice its volumetric compression
ratio
(RVC), and wherein the compressor pressure ratio (PR-GT) and the volumetric
compression ratio (R~C) are varied for constant leakages. The term "Net Shaft"
in
the axis "Thermal efficiency hIet Shaft" is intended to mean directly on the
output
shaft of the engine. Fig. 5 shows three (3) curves for different values of
compressor pressure ratio (PR-GT= 8; PR-GT= 6; and PR-GT= 4) when no
intercooler is used and three (3) additional curves for the same three
different
values of compressor pressure ratio (PR-GT= 8; PR-GT= 6; and PR-GT= 4) but
this time when an intercooler is used. ~n each curve, five different values of
the
volumetric compression ratio of the rotary engine (R"C 1.2; R"C 1.5; RVC 2;
R~C 3; and RVC 5) are provided.
More particularly, the inventor has found that, and Fig. 5 clearly
demonstrates that, when no intercooler is used, the thermal efficiency is
optimal
when the overall pressure ratio of the engine is about 40. When the overall
pressure ratio increases over 50, the thermal efficiency drops. From Fig. 5,
it can
thus be readily seen that under specific conditions (i.e. when the overall
pressure
ratio is below 50), the intercooler provides very little advantage to thermal
efficiency which is more offset by its weight, size and cost. It can also be
seen
that after a certain point, the thermal efficiency starts to decrease as the
volumetric compression ratio (R~~) of the rotary engines 16 increases.
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Considering the much-additional weight and size that an intercooler entails,
according to the present invention preferably, R,,~ is kept below 3.5 to
provide
optimal thermal efficiency without the need of an intercooler. Fig. 5 also
clearly
shows that the thermal efficiency of an integrated engine with no intercooler
and
having an R~~ of 3 with a compressor pressure ratio (PRgt) of 6 is almost as
good
as the thermal efficiency of an integrated engine with an intercooler.
However, if
the compressor is designed with a PRgt of 8, the R"~ must be reduced to 1.2 to
provide a thermal efficiency equivalent to an integrated engine with an
intercooler.
Fig. 6 shows four curves for two different values of the
compressor pressure ratio (PR-gt= 6 and PR-gt = 4), the first pair of curves,
which extends into the auto ignition zone, on the graph being for an engine
with
no intercooler and the two remaining curves at the bottom of the graph being
for
an engine with an intercooler. ~n each curve, five different values of the
volumetric compression ratio of the rotary engine (R~,C 1.2; RVC 1.5; R~,C 2;
RVC 3; and R"C 5) are provided.
As can be clearly seen in Fig. 6, in accordance with the present
invention, a lnnit line (shown with a thick stippled line in the Figure)
between an
"Auto-Ignition done" and a normal zone can be determined, based on the
properties of the fuel and fuel/air mixture used. As demonstrated by Fig. 6, a
careful selection of overall pressure ratio, and a careful allocation of
pressure
ratios between the gas turbine and the rotary engines, can be used to achieve
an
"auto-igniti~n-free" cycle. If no intercooler is being used, the volumetric
compression ratio (R"C) in the rotary engines has to be kept below
approximately
3 for a compressor pressure ratio (PRgt) of 6 and below approximately 3.5 for
a
PRgt of 4. in order to be out of the auto-ignition zone. The analysis of Fig.
6,
clearly show that by reducing the compression ratio, the air heats up less and
is
then further away from auto-ignition temperature, thereby obviating the need
for
an intercooler.
Im view of the foregoing, it appeaxs that a clear advantage of
limiting the volumetric compression ratio in the rotary engine below 3.5 is
that
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while the high thermal efficiency is maintained, the reduced pressure and
temperature prior to combustion allows to pre-mix the fuel with air prior to
the
rotary engines 16 to be done without auto-ignition and no need of an
intercooler
which is too bully for many aerospace applications, and particularly so for
commercial and commuter aircraft. As will be appreciated by the skilled
reader,
these cycle limitations are also applicable, and provide similar advantages,
to a
fuel injected configuration with spark ignition.
The low overall pressure ratio, i.e. preferably less than 50, with
low rotary engine compression volumetric ratio, i.e. preferably less than 3.5,
and
gas turbine pressure rati~, i.e. preferably less than 6, gives a compact
optimum
thermal efficiency cycle, easier to design with lower loads, less stress and
with
reduced leakage in seals and gaps. This cycle is particularly attractive to
rotary
machines designed with controlled rotating gaps as opposed to high speed seals
which are subject to wear.
It is noted that the rotary engine compression is described herein as
a "volumetric compression ratio" because it is readily measurable in such
closed
volume combustion engines by reason of its closed volume combustion design,
whereas the gas turbine compression described as a "pressure ratio" because of
the gas turbine's continuous flow design, in which pressures are more easily
measured instead of volume ratios.
The criteria to have a non-intercooled cycle with high thermal
efficiency (40-45°/~) in a compact engine package with improved power
to weight
ratio can be defined as follows:
PI~~t x lZ,,Ci.s < 30
where PR~t is the pressure ratio of the compressors) or gas turbine
engine compression stages) feeding the rotary engine, and
R~~ is the volumetric compression ratio of the rotary engine.
Typical values for optimum cycle efficiency are: PRgt= 3 - 6 and Rvc= 2 - 3.5,
and
full range of interest to meet above criteria 1.2 < PRgt < 9 and 1.2 < Rvc <
12
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As long as the above conditions are met, it will be possible to
operate without an intercooler to cool the air before it enters the rotary
engines
16. This advantageously provides for a very compact integrated engine package.
Furthermore, limiting the overall pressure ratio below 50 also contributes to
reduce the weight in that otherwise the wall thicl~ness of the rotary engines
would
have to be thicker and heavier.
The above-described combination of compression ratio in the
rotary engines and the gas turbine engine ensures that the temperature of the
pre-
mixed air/fuel mixture just prior to the combustion is below 1100 °F.
It is noted
that the above "pressure rules" applies to diesel or kerosene/jet engines type
of
fuel.
The above description is meant to be exemplary only, and one
slcilled in the art will recognise that changes may be made to the embodiments
described without departing from the scope of the invention disclosed. For
example, it is understood that the rotary engine could be replaced by several
rotary engines in parallel or series, or by other types of turbine cycle
topping
devices. For instance, a reciprocating engine could be used as well as a wave
engine coupled to a combustor. Rotary engines are however preferred for
compactness and speed compatibility (rotary engines have higher rotational
speed
potential vs. reciprocating engines). Another example is that instead of using
pre-
mix air/fuel upstream of the topping device, other configurations with fuel
injection directly into the topping device after air compression, to be
ignited with
sparlc ignition, may also be employed. The terms "accessory gearbox" and
"reduction gearbox" are used herein as those are familiar terms of gas turbine
art,
however the slcilled reader will appreciate that the gearbox provided may be
any
suitable transmission system, and may or may not include speed reduction,
depending on the application. Though one compression and one turbine stage is
shown, any suitable number of stages may be provided as desired. Still other
modifications which fall within the scope of the present invention will be
apparent to those skilled in the art, in light of a review of this disclosure,
and such
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modifications are intended to fall within the equivalents accorded to the
appended
claims.
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