U.S. patent application number 17/497606 was filed with the patent office on 2022-04-14 for aircraft.
This patent application is currently assigned to ROLLS-ROYCE plc. The applicant listed for this patent is ROLLS-ROYCE plc. Invention is credited to Thomas S. BINNINGTON, Daniel BLACKER, David A. JONES, Natalie C. WONG.
Application Number | 20220112866 17/497606 |
Document ID | / |
Family ID | 1000006063306 |
Filed Date | 2022-04-14 |
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United States Patent
Application |
20220112866 |
Kind Code |
A1 |
WONG; Natalie C. ; et
al. |
April 14, 2022 |
AIRCRAFT
Abstract
An aircraft comprises a machine body which encloses a turbofan
gas turbine engine. The turbofan gas turbine engine includes a heat
exchanger module, fan assembly, compressor module, turbine module,
and exhaust module. The heat exchanger module communicates with the
fan assembly by an inlet duct. The heat exchanger module includes
first heat transfer elements that transfer heat energy from a first
fluid within the transfer elements to an airflow passing over a
surface of the transfer elements before entry of the airflow into a
fan assembly inlet. The first fluid contained within transfer
elements has a temperature, and the airflow passing over the
transfer element surface has a temperature. The turbofan gas
turbine engine further includes at least one second heat transfer
element, with the or each second heat transfer element transfers
heat energy from the first fluid to a second fluid.
Inventors: |
WONG; Natalie C.; (Bristol,
GB) ; BINNINGTON; Thomas S.; (Bristol, GB) ;
JONES; David A.; (Bristol, GB) ; BLACKER; Daniel;
(Bristol, GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
|
GB |
|
|
Assignee: |
ROLLS-ROYCE plc
London
GB
|
Family ID: |
1000006063306 |
Appl. No.: |
17/497606 |
Filed: |
October 8, 2021 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02K 3/06 20130101; B64D
27/20 20130101; B64D 2033/024 20130101; F05D 2220/323 20130101;
F02K 3/075 20130101; F05D 2260/213 20130101; B64D 33/10
20130101 |
International
Class: |
F02K 3/075 20060101
F02K003/075; F02K 3/06 20060101 F02K003/06; B64D 33/10 20060101
B64D033/10; B64D 27/20 20060101 B64D027/20 |
Foreign Application Data
Date |
Code |
Application Number |
Oct 9, 2020 |
GB |
2016000.8 |
Jan 26, 2021 |
GB |
2101017.8 |
Claims
1. An aircraft comprising a machine body, the machine body
enclosing a turbofan gas turbine engine, the turbofan gas turbine
engine comprising, in axial flow sequence, a heat exchanger module,
a fan assembly, a compressor module, a turbine module, and an
exhaust module, the heat exchanger module being in fluid
communication with the fan assembly by an inlet duct, the heat
exchanger module comprising a plurality of first heat transfer
elements, the first heat transfer elements being configured for the
transfer of heat energy from a first fluid contained within the
first heat transfer elements to an airflow passing over a surface
of the first heat transfer elements prior to entry of the airflow
into an inlet to the fan assembly, the first fluid contained within
the first heat transfer elements having a temperature, the airflow
passing over the surface of the first heat transfer elements having
a temperature, the turbofan gas turbine engine further comprising
at least one second heat transfer element, the or each second heat
transfer element being configured for the transfer of heat energy
from the first fluid to a second fluid, wherein, in use, the
aircraft can maintain a sustained airspeed, and when the airflow
temperature is less than the first fluid temperature, the first
fluid passes through the first heat transfer elements, and when the
airflow temperature is equal to or greater than the first fluid
temperature, the first fluid passes through the second heat
transfer elements.
2. The aircraft as claimed in claim 1, wherein the fan assembly
comprises a plurality of fan blades defining a fan diameter, and
the fan diameter is within the range of 0.3 m to 2.0 m.
3. The aircraft as claimed in claim 1, wherein the heat exchanger
module has a fluid path diameter, wherein the fluid path diameter
is greater than the fan diameter.
4. The aircraft as claimed in claim 1, wherein the turbofan gas
turbine engine further comprises an outer housing, the outer
housing enclosing the sequential arrangement of heat exchanger
module, fan assembly, compressor module, and turbine module, an
annular bypass duct being defined between the outer housing and the
sequential arrangement of modules, a bypass ratio being defined as
a ratio of a mass air flow rate through the bypass duct to a mass
air flow rate through the sequential arrangement of modules, and
wherein the bypass ratio is less than 4.0.
5. The aircraft as claimed in claim 1, wherein the fan assembly
comprises two or more fan stages, at least one of the fan stages
comprising a plurality of fan blades defining the fan diameter.
6. A method of operating an aircraft comprising a machine body, the
machine body enclosing a turbofan gas turbine engine, the gas
turbine engine comprising, in axial flow sequence, a heat exchanger
module, an inlet duct, a fan assembly, a compressor module, a
turbine module, and an exhaust module, and wherein the method
comprises the steps of: (i) providing the fan assembly, the
compressor module, the turbine module, and the exhaust module; (ii)
providing the heat exchanger module with a plurality of first heat
transfer elements for transfer of heat from a first fluid contained
within the heat transfer elements to an airflow passing over a
surface of the heat transfer elements prior to entry of the airflow
into the fan assembly, the first fluid having a temperature; (iii)
providing the heat exchanger module with at least one second heat
transfer element for the transfer of heat from the first fluid to a
second fluid (iv) positioning the heat exchanger module in fluid
communication with the fan assembly by the inlet duct; and (v)
operating the aircraft such that the aircraft maintains a sustained
airspeed V and the airflow passing over a surface of the heat
transfer elements has a temperature; (vi) if the airflow
temperature is less than the first fluid temperature, directing the
first fluid through the first heat transfer elements; and (vii) if
the airflow temperature is equal to or greater than the first fluid
temperature, directing the first fluid through the second heat
transfer elements.
Description
[0001] This disclosure claims the benefit of UK Patent Application
No. GB 2016000.8, filed on 9 Oct. 2020, and UK Patent Application
No. GB 2101017.8, filed on 26 Jan. 2021, each of which is hereby
incorporated herein in its entirety.
FIELD OF THE DISCLOSURE
[0002] The present disclosure relates to an aircraft having a
single engine air intake aperture and a single engine exhaust
aperture for each engine and particularly, but not exclusively, to
an aircraft having a single engine air intake aperture and a single
engine exhaust aperture with an airframe heat exchanger positioned
in the air intake aperture.
BACKGROUND TO THE DISCLOSURE
[0003] A conventional turbofan gas turbine engine uses heat
exchangers to cool a variety of fluids including inter alia air,
fuel and oil. Typically, such heat exchangers use bypass air or an
air offtake from the compressor as the cooling medium. The heat
exchanger itself may be positioned in the bypass duct or externally
to the engine with the corresponding ducting.
[0004] The use of bypass air or a compressor offtake stream as the
cooling medium in a heat exchanger will adversely affect the
performance of the engine, for example by reducing specific thrust
or increasing specific fuel consumption. Alternatively, or
additionally, such offtakes can adversely affect engine
performance, for example by reducing surge margin.
[0005] In a further alternative conventional arrangement, an
airflow to provide the cooling medium in a heat exchanger may be
drawn separately from the airflow through the gas turbine engine.
For example, in an airframe application the airflow providing the
cooling medium may be drawn from an air intake or duct separate
from the engine.
[0006] As used herein, a range "from value X to value Y" or
"between value X and value Y", or the likes, denotes an inclusive
range; including the bounding values of X and Y. As used herein,
the term "axial plane" denotes a plane extending along the length
of an engine, parallel to and containing an axial centreline of the
engine, and the term "radial plane" denotes a plane extending
perpendicular to the axial centreline of the engine, so including
all radial lines at the axial position of the radial plane. Axial
planes may also be referred to as longitudinal planes, as they
extend along the length of the engine. A radial distance or an
axial distance is therefore a distance in a radial or axial plane,
respectively.
STATEMENTS OF DISCLOSURE
[0007] According to a first aspect of the present disclosure, there
is provided an aircraft comprising a machine body, the machine body
enclosing a turbofan gas turbine engine, the turbofan gas turbine
engine comprising, in axial flow sequence, a heat exchanger module,
a fan assembly, a compressor module, a turbine module, and an
exhaust module, the heat exchanger module being in fluid
communication with the fan assembly by an inlet duct, the heat
exchanger module comprising a plurality of first heat transfer
elements, the first heat transfer elements being configured for the
transfer of heat energy from a first fluid contained within the
first heat transfer elements to an airflow passing over a surface
of the first heat transfer elements prior to entry of the airflow
into an inlet to the fan assembly, the first fluid contained within
the first heat transfer elements having a temperature T.sub.F, the
airflow passing over the surface of the first heat transfer
elements having a temperature T.sub.A, the turbofan gas turbine
engine further comprising at least one second heat transfer
element, the or each second heat transfer element being configured
for the transfer of heat energy from the first fluid to a second
fluid,
[0008] wherein, in use, the aircraft can maintain a sustained
airspeed V (M), and when the airflow temperature T.sub.A is less
than the first fluid temperature T.sub.F, the first fluid passes
through the first heat transfer elements, and when the airflow
temperature T.sub.A is equal to or greater than the first fluid
temperature T.sub.F, the first fluid passes through the second heat
transfer elements.
[0009] When the airflow temperature T.sub.A is less than the first
fluid temperature T.sub.F the first fluid is circulated through the
first heat transfer elements in the heat exchanger module and heat
energy from the first fluid is rejected to the airflow passing
through the heat exchanger module and passing over the surface of
the first heat transfer elements.
[0010] When the airflow temperature T.sub.A is less than the first
fluid temperature T.sub.F, the waste heat energy can be effectively
transferred from the first fluid to the air flow in the first heat
transfer elements. For example, at a sustained airspeed of M 0.95
the incoming airflow may have a temperature of approximately
65.degree. C., while the first fluid may typically have a
temperature of approximately 150.degree. C.
[0011] However, when the sustained airspeed increases, for example
to a supersonic level, the temperature of the air flow entering the
heat exchanger increases above the ambient air temperature. This is
because the air flow must be slowed to a subsonic level as it
enters the inlet to the turbofan engine. In the process of slowing
the air flow its temperature increases because of the viscous
losses caused in slowing the air flow.
[0012] The temperature of the airflow may therefore be greater than
the temperature of the first fluid, which makes it impossible to
reject waste heat energy from the first fluid to the airflow
passing through the heat exchanger module.
[0013] Consequently, at a high sustained airspeed, for example M1.0
or greater, the first fluid passes through the second heat transfer
elements and heat energy from the first fluid is rejected to the
second fluid in passing through the second heat transfer elements.
The second fluid temperature is greater than the temperature of the
ambient air that would enter the turbofan engine at a low sustained
airspeed, for example less than M1.0, but is less than the
temperature of the air flow entering the turbofan engine at a high
sustained airspeed, for example greater than M1.0.
[0014] Optionally, the fan assembly comprises a plurality of fan
blades defining a fan diameter (D), and the fan diameter D is
within the range of 0.3 m to 2.0 m, preferably within the range 0.4
m to 1.5 m, and more preferably in the range of 0.7 m to 1.0 m. In
one embodiment of the disclosure, the fan diameter is 0.9 m.
[0015] Consequently, for the same heat energy loading rejected to
the air flow through the heat exchanger, the loss in propulsive
efficiency of the turbofan engine is proportionately smaller for a
large diameter (for example, approximately 1.5 to 2.0 m in
diameter) turbofan engine than for a small diameter turbofan
engine.
[0016] The fan tip diameter, measured across a centreline of the
engine and between an outermost tip of opposing fan blades at their
leading edge, may be in the range from 95 cm to 200 cm, for example
in the range from 110 cm to 150 cm, or alternatively in the range
from 155 cm to 200 cm. The fan tip diameter may be greater than any
of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm,
150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190
cm or 195 cm. The fan tip diameter may be around 110 cm, 115 cm,
120 cm, 125 cm, 130 cm, 135 cm, 140 cm, 145 cm, 150 cm, 155 cm, 160
cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The
fan tip diameter may be greater than 160 cm.
[0017] The fan tip diameter may be in the range from 95 cm to 150
cm, optionally in the range from 110 cm to 150 cm, optionally in
the range of from 110 cm to 145 cm, and further optionally in the
range from 120 cm to 140 cm.
[0018] The fan tip diameter may be in the range from 155 cm to 200
cm, optionally in the range from 160 cm to 200 cm, and further
optionally in the range from 165 cm to 190 cm.
[0019] Optionally, the heat exchanger module has a flow area
A.sub.HEX and the fan module has a flow area A.sub.FAN, and a ratio
of A.sub.FAN to .sub.AHEX being in the range of 0.3 to 1.0.
[0020] The flow area is to be understood to mean a cross-sectional
area of the air flow taken perpendicularly to a central axis of the
flow in the flow direction. In other words, for the heat exchanger
module the flow area A.sub.HEX corresponds to the cross-sectional
area of the heat exchanger module through which the flow passes.
Likewise, for the fan assembly the flow area A.sub.FAN corresponds
to the cross-sectional area of the fan assembly through which the
flow passes.
[0021] Optionally, the heat exchanger module has a fluid path
diameter E, wherein the fluid path diameter E is greater than the
fan diameter D.
[0022] In one embodiment, the heat exchanger module has a fluid
path diameter E that is greater than the fan diameter D. In this
embodiment, the inlet duct that connects the heat exchanger module
to the fan assembly has a diameter than converges from an exit from
the heat exchanger module to an entrance to the fan assembly.
[0023] Optionally, the turbofan gas turbine engine further
comprises an outer housing, the outer housing enclosing the
sequential arrangement of heat exchanger module, fan assembly,
compressor module, and turbine module, an annular bypass duct being
defined between the outer housing and the sequential arrangement of
modules, a bypass ratio being defined as a ratio of a mass air flow
rate through the bypass duct to a mass air flow rate through the
sequential arrangement of modules, and wherein the bypass ratio is
less than 4.0.
[0024] A turbofan engine having a bypass ratio (BPR) of less than
approximately 4.0 will have a generally smaller bypass duct (the
annular duct surrounding the core gas turbine engine) than a
turbofan engine having a BPR greater than approximately 4.0. For a
turbofan engine with a BPR greater than, say, 4.0, the
correspondingly larger bypass duct volume provides more scope for
positioning a heat exchanger within the bypass duct than would be
the case for a low BPR turbofan engine.
[0025] Optionally, the fan assembly comprises two or more fan
stages, at least one of the fan stages comprising a plurality of
fan blades defining the fan diameter D.
[0026] In one arrangement, the fan assembly has two fan stages with
both fan stages comprising a plurality of fan blades defining the
same fan diameter. Alternatively, each of the fan stages may have
different fan diameters.
[0027] According to a further aspect of the present disclosure,
there is provided a method of operating an aircraft comprising a
machine body, the machine body enclosing a turbofan gas turbine
engine, the gas turbine engine comprising, in axial flow sequence,
a heat exchanger module, an inlet duct, a fan assembly, a
compressor module, a turbine module, and an exhaust module, and
wherein the method comprises the steps of: [0028] (i) providing the
fan assembly, the compressor module, the turbine module, and the
exhaust module; [0029] (ii) providing the heat exchanger module
with a plurality of first heat transfer elements for transfer of
heat from a first fluid contained within the heat transfer elements
to an airflow passing over a surface of the heat transfer elements
prior to entry of the airflow into the fan assembly, the first
fluid having a temperature T.sub.F; [0030] (iii) providing the heat
exchanger module with at least one second heat transfer element for
the transfer of heat from the first fluid to a second fluid [0031]
(iv) positioning the heat exchanger module in fluid communication
with the fan assembly by the inlet duct; and [0032] (v) operating
the aircraft such that the aircraft maintains a sustained airspeed
V and the airflow passing over a surface of the heat transfer
elements has a temperature T.sub.A; [0033] (vi) if temperature
T.sub.A is less than temperature T.sub.F, directing the first fluid
through the first heat transfer elements; and [0034] (vii) if
temperature T.sub.A is equal to or greater than temperature
T.sub.F, directing the first fluid through the second heat transfer
elements.
[0035] When the airflow temperature T.sub.A is less than the first
fluid temperature T.sub.F, the first fluid is circulated through
the first heat transfer elements in the heat exchanger module and
heat energy from the first fluid is rejected to the airflow passing
through the heat exchanger module and passing over the surface of
the first heat transfer elements.
[0036] When the airflow temperature T.sub.A is less than the first
fluid temperature T.sub.F, the waste heat energy can be effectively
transferred from the first fluid to the air flow in the first heat
transfer elements. For example, at a sustained airspeed of M 0.95
the incoming airflow may have a temperature of approximately
65.degree. C., while the first fluid may typically have a
temperature of approximately 150.degree. C.
[0037] However, when the sustained airspeed increases, for example
to a supersonic level, the temperature of the air flow entering the
heat exchanger increases above the ambient air temperature. This is
because as the air flow must be slowed to a subsonic level as it
enters the inlet to the turbofan engine. In the process of slowing
the air flow its temperature increases because of the viscous
losses caused in slowing the air flow.
[0038] The temperature of the air flow may therefore be greater
than the temperature of the first fluid, which makes it impossible
to reject waste heat energy from the first fluid to the airflow
passing through the heat exchanger module.
[0039] Consequently, at a high sustained airspeed, for example M1.0
or greater, the first fluid passes through the second heat transfer
elements and heat energy from the first fluid is rejected to the
second fluid in passing through the second heat transfer elements.
The second fluid temperature is greater than the temperature of the
ambient air that would enter the turbofan engine at a low sustained
airspeed, for example less than M1.0, but is less than the
temperature of the air flow entering the turbofan engine at a high
sustained airspeed, for example greater than M1.0.
[0040] According to a further aspect of the present disclosure,
there is provided an aircraft comprising a machine body, the
machine body enclosing a turbofan gas turbine engine and a
plurality of ancillary systems, the turbofan gas turbine engine
comprising, in axial flow sequence, a heat exchanger module, a fan
assembly, a compressor module, a turbine module, and an exhaust
module; [0041] wherein the machine body comprises a single fluid
inlet aperture, the fluid inlet aperture being configured to allow
a fluid flow to enter the machine body and to pass through the heat
exchanger module, the heat exchanger module being configured to
transfer a waste heat load from the gas turbine engine and the
ancillary systems to the fluid flow prior to an entry of the fluid
flow into the fan module, and thence the compressor module, the
turbine module, and the exhaust module.
[0042] By performing all aircraft cooling through the inlet air
stream that feeds the turbofan gas turbine engine, other air
intakes and scoops to provide cooling air flows to the airframe can
be eliminated. Minimising the quantity of apertures on the airframe
will reduce the aerodynamic drag on the airframe and so improve the
aerodynamic efficiency of the airframe.
[0043] The use of `cool` inlet air to absorb waste heat load
maximises the heat exchanger efficiency. Reducing the temperature
of the heat energy absorbing fluid in the heat exchanger will
increase the efficiency of the heat exchanger by increasing the
temperature differential between the primary and secondary fluids
in the heat exchanger.
[0044] In this arrangement, all the active heat rejection (from the
gas turbine engine and the ancillary systems) is rejected into the
fluid flow entering the gas turbine engine. It will be understood
that there may be some passive heat rejection through the surface
of the machine body.
[0045] In contrast, the engine offtake flow used in a conventional
turbofan engine (for example from a compressor stage or from the
bypass flow) for providing a cooling feed flow to a heat exchanger
will have a considerably higher temperature than the intake air
flow entering the turbofan engine. This in turn limits the
temperature at which heat can be rejected to the heat exchanger to
temperatures that are greater than that of the corresponding engine
offtake flow. Such a prior art arrangement will also require a
larger and less efficient heat exchanger than that of the present
invention for the same quantity of heat rejection. In addition,
such conventional arrangements result in a loss of propulsive flow
and a consequent reduction in overall engine efficiency.
[0046] In an aircraft according to the present disclosure, the
location of the heat exchanger in an inlet duct upstream of the fan
assembly means that the engine is able to use all of the intake air
that is used by the heat exchanger module to dissipate waste energy
to subsequently provide propulsive thrust. This means that an
aircraft according to the present disclosure can be more efficient
than a conventional aircraft while providing the same level of heat
exchanger capacity.
[0047] The positioning of the heat exchanger upstream of the fan
assembly means that the cooling air flow entering the heat
exchanger has a temperature that is lower than any engine offtake
flow, for example a bleed flow from a compressor stage or the
bypass flow. Consequently, in a turbofan engine according to the
present disclosure it becomes possible to reject heat to the heat
exchanger flow at a lower temperature than can be achieved with any
prior art arrangement described above. This makes the turbofan
engine according to the present disclosure more versatile than
conventional turbofan engines.
[0048] Optionally, the machine body comprises two fluid inlet
apertures, each of the fluid inlet apertures being configured to
allow a respective fluid flow to enter the machine body, the two
fluid flows then being blended together prior to entry of the
blended fluid flow into the heat exchanger module.
[0049] In an alternative arrangement, the machine body may comprise
two fluid inlet apertures. Each of the two fluid inlet apertures is
configured to allow a respective intake air flow to enter the
machine body. The two intake air flows are then blended together
within the machine body.
[0050] The single blended intake air flow to the turbofan gas
turbine engine passes through the heat exchanger module prior to
entry into the fan module and thence to the compressor, combustor,
turbine and exhaust modules.
[0051] In an alternative arrangement, a heat exchanger module is
provided in each of the two fluid inlet flows between the
respective inlet aperture and the point at which the two intake air
flows are blended together. These two heat exchanger modules
replace the single heat exchanger module upstream of the fan that
has been previously described.
[0052] Optionally, the machine body further comprises a single
fluid exhaust aperture, the fluid exhaust aperture being configured
to channel the fluid flow from the or each exhaust module out of
the machine body.
[0053] The use of a single exhaust aperture minimises the
aerodynamic drag associated with an exhaust stream from the
turbofan gas turbine engine (or engines) leaving the aircraft.
[0054] In an arrangement with two turbofan gas turbine engines, the
exhaust flow from the exhaust modules of each of the turbofan gas
turbine engines exits the airframe through a single exhaust
aperture. The exhaust aperture may be bifurcated with each
bifurcated portion carrying an exhaust flow from a respective one
of the turbofan gas turbine engines.
[0055] According to a further aspect of the present disclosure,
there is provided a method of operating an aircraft, the aircraft
comprising a machine body, the machine body enclosing a turbofan
gas turbine engine and a plurality of ancillary systems, the
turbofan gas turbine engine comprising, in axial flow sequence, a
heat exchanger module, a fan assembly, a compressor module, a
combustor module, a turbine module, and an exhaust module; [0056]
wherein the method comprises the steps of: [0057] (i) providing the
machine body; [0058] (ii) arranging the fan assembly, the
compressor module, the combustor module, the turbine module, and
the exhaust module within the machine body; [0059] (iii) providing
the machine body with a single fluid inlet aperture, the fluid
inlet aperture being configured to allow a fluid flow to enter the
machine body and to pass through the heat exchanger module; [0060]
(iv) configuring the heat exchanger module to transfer a waste heat
load from the gas turbine engine and the ancillary systems to the
fluid flow prior to an entry of the fluid flow into the fan module;
and [0061] (v) operating the engine such that the waste heat load
from the gas turbine engine and the ancillary systems is
transferred to the fluid flow prior to the entry of the fluid flow
into the fan module.
[0062] By performing all aircraft cooling through the inlet air
stream that feeds the turbofan gas turbine engine, other air
intakes and scoops to provide cooling air flows to the airframe can
be eliminated. Minimising the quantity of apertures on the airframe
will reduce the aerodynamic drag on the airframe and so improve the
aerodynamic efficiency of the airframe.
[0063] The use of `cool` inlet air to absorb waste heat load
maximises the heat exchanger efficiency. Reducing the temperature
of the heat energy absorbing fluid in the heat exchanger will
increase the efficiency of the heat exchanger by increasing the
temperature differential between the primary and secondary fluids
in the heat exchanger.
[0064] In contrast, the engine offtake flow used in a conventional
turbofan engine (for example from a compressor stage or from the
bypass flow) for providing a cooling feed flow to a heat exchanger
will have a considerably higher temperature than the intake air
flow entering the turbofan engine. This in turn limits the
temperature at which heat can be rejected to the heat exchanger to
temperatures that are greater than that of the corresponding engine
offtake flow. In addition, such conventional arrangements result in
a loss of propulsive flow and a consequent reduction in overall
engine efficiency.
[0065] In an aircraft according to the present disclosure, the
location of the heat exchanger in an inlet duct upstream of the fan
assembly means that the engine is able to use all of the intake air
that is used by the heat exchanger module to dissipate waste energy
to subsequently provide propulsive thrust. This means that an
aircraft according to the present disclosure can be more efficient
than a conventional aircraft while providing the same level of heat
exchanger capacity.
[0066] The positioning of the heat exchanger upstream of the fan
assembly means that the cooling air flow entering the heat
exchanger has a temperature that is lower than any engine offtake
flow, for example a bleed flow from a compressor stage or the
bypass flow. Consequently, in a turbofan engine according to the
present disclosure it becomes possible to reject heat to the heat
exchanger flow at a lower temperature than can be achieved with any
prior art arrangement described above. This makes the turbofan
engine according to the present disclosure more versatile than
conventional turbofan engines.
[0067] Optionally, step (iii) comprises the steps of: [0068]
(iii-a) providing the machine body with two fluid inlet apertures,
each of the two fluid inlet apertures being configured to allow a
respective fluid flow to enter the machine body; [0069] (iii-b)
blending together the two fluid flows prior to entry of the blended
fluid flow into the heat exchanger module.
[0070] In an alternative arrangement, the machine body may comprise
two fluid inlet apertures. Each of the two fluid inlet apertures is
configured to allow a respective intake air flow to enter the
machine body. The two intake air flows are then blended together
within the machine body.
[0071] The single blended intake air flow to the turbofan gas
turbine engine passes through a heat exchange module prior to entry
into the fan module and thence to the compressor, combustor,
turbine and exhaust modules.
[0072] Optionally, step (iii) comprises the additional following
step of: [0073] (iii)' providing the machine body with a single
fluid exhaust aperture, the fluid exhaust aperture being configured
to channel the fluid flow from the exhaust module out of the
machine body.
[0074] The use of a single exhaust aperture minimises the
aerodynamic drag associated with an exhaust stream from the
turbofan gas turbine engine (or engines) leaving the aircraft.
[0075] In an arrangement with two turbofan gas turbine engines, the
exhaust flow from the exhaust modules of each of the turbofan gas
turbine engines exits the airframe through a single exhaust
aperture. The exhaust aperture may be bifurcated with each
bifurcated portion carrying an exhaust flow from a respective one
of the turbofan gas turbine engines.
[0076] According to a further aspect of the present disclosure,
there is provided an aircraft comprising a machine body, the
machine body enclosing a turbofan gas turbine engine and a
plurality of ancillary systems, the turbofan gas turbine engine
comprising, in axial flow sequence, a heat exchanger module, a fan
assembly, a compressor module, a combustor module, a turbine
module, and an exhaust module; [0077] wherein the machine body
comprises a single fluid inlet aperture, the fluid inlet aperture
being configured to allow a fluid cooling flow to enter the machine
body and to pass through the heat exchanger module, the heat
exchanger module being configured to transfer a waste heat load
from the gas turbine engine and the ancillary systems to the fluid
cooling flow prior to an entry of the fluid cooling flow into the
fan module.
[0078] In an aircraft according to this arrangement, all of the air
that is used to cool both the engine and any ancillary systems of
the aircraft is ingested by the turbofan gas turbine engine. This
means that the entirety of the cooling air flow used by the
aircraft passes through the turbofan gas turbine engine and so
contributes to providing thrust to the aircraft.
[0079] In contrast, a conventional aircraft comprises external
cooling inlets (such as scoops or ducts) on the airframe that
provide cooling air to separate heat exchangers that cool the
engine system and/or airframe ancillary systems. In such
arrangements, the cooling air flows make no contribution to
aircraft thrust and indeed cause parasitic drag on the
aircraft.
[0080] By arranging for all of the cooling air to pass through the
turbofan engine, the overall efficiency of the aircraft can be
increased relative to conventional arrangements as described in the
previous paragraph.
[0081] Optionally, a proportion B.sub.COMB of the fluid cooling
flow passes sequentially through the compressor, combustor, and
turbine modules, [0082] wherein the B.sub.COMB parameter is defined
in the range of 0.20 to 0.71
[0083] In one arrangement of the engine with a bypass ratio of 0.4,
the proportion of the fluid cooling flow that passes the engine
core, namely the B.sub.COMB parameter, is 0.71.
[0084] Optionally, the B.sub.COMB parameter is defined in the range
of 0.29 to 0.71.
[0085] In another arrangement of the engine, this time with a
bypass ratio of 1.0, the proportion of the fluid cooling flow that
passes the engine core, namely the B.sub.COMB parameter, is
0.50.
[0086] According to a further aspect of the present disclosure,
there is provided a method of operating an aircraft, the aircraft
comprising a machine body, the machine body enclosing a turbofan
gas turbine engine and a plurality of ancillary systems, the
turbofan gas turbine engine comprising, in axial flow sequence, a
heat exchanger module, a fan assembly, a compressor module, a
combustor module, a turbine module, and an exhaust module; [0087]
wherein the method comprises the steps of: [0088] (i) providing the
machine body; [0089] (ii) arranging the fan assembly, the
compressor module, the combustor module, the turbine module, and
the exhaust module within the machine body; [0090] (iii) providing
the machine body with a single fluid inlet aperture, the fluid
inlet aperture being configured to allow a fluid flow to enter the
machine body and to pass through the heat exchanger module; [0091]
(iv) configuring the heat exchanger module to transfer a waste heat
load from the gas turbine engine and the ancillary systems to the
fluid flow prior to an entry of the fluid flow into the fan module;
and [0092] (v) operating the engine such that the entire fluid flow
enters the fan module.
[0093] In an aircraft according to this arrangement, all of the air
that is used to cool both the engine and any ancillary systems of
the aircraft is ingested by the turbofan gas turbine engine. This
means that the entirety of the cooling air flow used by the
aircraft passes through the turbofan gas turbine engine and so
contributes to providing thrust to the aircraft.
[0094] In contrast, a conventional aircraft comprises external
cooling inlets (such as scoops or ducts) on the airframe that
provide cooling air to separate heat exchangers that cool the
engine system and/or airframe ancillary systems. In such
arrangements, the cooling air flows make no contribution to
aircraft thrust and indeed cause parasitic drag on the
aircraft.
[0095] By arranging for all of the cooling air to pass through the
turbofan engine, the overall efficiency of the aircraft can be
increased relative to conventional arrangements as described in the
previous paragraph.
[0096] Optionally, step (v) comprises the step of: [0097] (v)'
operating the engine such that the entire fluid flow enters the fan
module, and a proportion B.sub.COMB of the fluid flow passes
sequentially through the compressor, combustor, and turbine
modules, and the B.sub.COMB parameter is defined in the range of
0.20 to 0.71.
[0098] In one arrangement of the engine with a bypass ratio of 0.4,
the proportion of the fluid cooling flow that passes the engine
core, namely the B.sub.COMB parameter, is 0.71.
[0099] The skilled person will appreciate that a feature described
above in relation to any one of the aspects may be applied, mutatis
mutandis, to any other aspect of the invention. For example, in
various embodiments any two or more of the conditions for ratios as
defined above, and optionally all specified ratio ranges, may apply
to any given aspect or embodiment. All aspects may apply to an
engine of some embodiments. Furthermore, any feature described
below may apply to any aspect and/or may apply in combination with
any one of the claims.
[0100] As noted elsewhere herein, the present disclosure may relate
to a turbofan gas turbine engine. Such a gas turbine engine may
comprise an engine core comprising a turbine, a combustor, a
compressor, and a core shaft connecting the turbine to the
compressor. Such a gas turbine engine may comprise a fan (having
fan blades) located upstream of the engine core. The fan may
comprise any number of stages, for example multiple stages. Each
fan stage may comprise a row of fan blades and a row of stator
vanes. The stator vanes may be variable stator vanes (in that their
angle of incidence may be variable).
[0101] The turbofan gas turbine engine as described and/or claimed
herein may have any suitable general architecture. For example, the
gas turbine engine may have any desired number of shafts that
connect turbines and compressors, for example one, two or three
shafts. Purely by way of example, the turbine connected to the core
shaft may be a first turbine, the compressor connected to the core
shaft may be a first compressor, and the core shaft may be a first
core shaft. The engine core may further comprise a second turbine,
a second compressor, and a second core shaft connecting the second
turbine to the second compressor. The second turbine, second
compressor, and second core shaft may be arranged to rotate at a
higher rotational speed than the first core shaft.
[0102] In such an arrangement, the second compressor may be
positioned axially downstream of the first compressor. The second
compressor may be arranged to receive (for example directly
receive, for example via a generally annular duct) flow from the
first compressor.
[0103] In any turbofan gas turbine engine as described and/or
claimed herein, a combustor may be provided axially downstream of
the fan and compressor(s). For example, the combustor may be
directly downstream of (for example at the exit of) the second
compressor, where a second compressor is provided. By way of
further example, the flow at the exit to the combustor may be
provided to the inlet of the second turbine, where a second turbine
is provided. The combustor may be provided upstream of the
turbine(s).
[0104] The or each compressor (for example the first compressor and
second compressor as described above) may comprise any number of
compressor stages, for example multiple stages. Each compressor
stage may comprise a row of rotor blades and a row of stator vanes.
The stator vanes may be variable stator vanes (in that their angle
of incidence may be variable). The row of rotor blades and the row
of stator vanes may be axially offset from each other.
[0105] The or each turbine (for example the first turbine and
second turbine as described above) may comprise any number of
turbine stages, for example multiple stages. Each turbine stage may
comprise a row of rotor blades and a row of stator vanes. The row
of rotor blades and the row of stator vanes may be axially offset
from each other.
[0106] Each fan blade may be defined as having a radial span
extending from a root (or hub) at a radially inner gas-washed
location, or 0% span position, to a tip at a 100% span position.
The ratio of the radius of the fan blade at the hub to the radius
of the fan blade at the tip may be less than (or on the order of)
any of: 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31,
0.30, 0.29, 0.28, 0.27 or 0.26. The ratio of the radius of the fan
blade at the hub to the radius of the fan blade at the tip may be
in an inclusive range bounded by any two of the values in the
previous sentence (i.e. the values may form upper or lower bounds),
for example in the range of from 0.28 to 0.32. These ratios may
commonly be referred to as the hub-to-tip ratio. The radius at the
hub and the radius at the tip may both be measured at the leading
edge (or axially forwardmost) part of the blade. The hub-to-tip
ratio refers, of course, to the gas-washed portion of the fan
blade, i.e. the portion radially outside any platform.
[0107] The diameter of the fan may be measured across the engine
centreline and between the tips of opposing fan blades at their
leading edge. The fan diameter may be greater than (or on the order
of) any of: 50 cm, 60 cm, 70 cm (around 27.5 inches), 80 cm (around
31.5 inches), 90 cm, 100 cm (around 39 inches), 110 cm (around 43
inches), 120 cm (around 47 inches), 130 cm (around 51 inches), 140
cm (around 55 inches), 150 cm (around 59 inches), or 160 cm (around
130 inches). The fan diameter may be in an inclusive range bounded
by any two of the values in the previous sentence (i.e. the values
may form upper or lower bounds), for example in the range of from
50 cm to 70 cm or 90 cm to 130 cm.
[0108] The fan face area may be equal to .pi. multiplied by the
square of the fan tip radius.
[0109] The rotational speed of the fan may vary in use. Generally,
the rotational speed is lower for fans with a higher diameter.
Purely by way of non-limitative example, the rotational speed of
the fan in steady-state flight conditions may be less than 10000
rpm, for example less than 9000 rpm. Purely by way of further
non-limitative example, the rotational speed of the fan in
steady-state flight conditions for an engine having a fan diameter
in the range of from 50 cm to 90 cm (for example 60 cm to 80 cm or
65 cm to 75 cm) may be in the range of from 7000 rpm to 22000 rpm,
for example in the range of from 7000 rpm to 16000 rpm, for example
in the range of from 7500 rpm to 14000 rpm. Purely by way of
further non-limitative example, the rotational speed of the fan in
steady-state flight conditions for an engine having a fan diameter
in the range of from 90 cm to 150 cm may be in the range of from
4500 rpm to 12500 rpm, for example in the range of from 4500 rpm to
10000 rpm, for example in the range of from 6000 rpm to 10000
rpm.
[0110] In use of the turbofan gas turbine engine, the fan (with
associated fan blades) rotates about a rotational axis. This
rotation results in the tip of the fan blade moving with a velocity
U.sub.tip. The work done by the fan blades 13 on the flow results
in an enthalpy rise dH of the flow. A fan tip loading may be
defined as dH/U.sub.tip.sup.2, where dH is the enthalpy rise (for
example the 1-D average enthalpy rise) across the fan and U.sub.tip
is the (translational) velocity of the fan tip, for example at the
leading edge of the tip (which may be defined as fan tip radius at
leading edge multiplied by angular speed). The fan tip loading at
cruise conditions may be greater than (or on the order of) any of:
0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.30, 0.31, 0.32,
0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.40 (all values being
dimensionless). The fan tip loading may be in an inclusive range
bounded by any two of the values in the previous sentence (i.e. the
values may form upper or lower bounds), for example in the range of
from 0.28 to 0.31, or 0.29 to 0.30.
[0111] Turbofan gas turbine engines in accordance with the present
disclosure may have any desired bypass ratio, where the bypass
ratio is defined as the ratio of the mass flow rate of the flow
through the bypass duct to the mass flow rate of the flow through
the core at cruise conditions. In some arrangements the bypass
ratio may be greater than (or on the order of) any of the
following: 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1.0, 1.1, 1.2, 1.3, 1.4,
1.5, 1.6, 1.7, 1.8, 1.9, 2.0, 2.4, 2.8, 3.2, 3.6, or 4.0. The
bypass ratio may be in an inclusive range bounded by any two of the
values in the previous sentence (i.e. the values may form upper or
lower bounds), for example in the range of form of 0.4 to 1.0, 0.5
to 0.9, or 0.6 to 0.9. Alternatively, the bypass ratio may be in a
bounded range in the form of 1.0 to 4.0, 1.8 to 3.6, or 2.4 to 3.6.
The bypass duct may be substantially annular. The bypass duct may
be radially outside the core engine. The radially outer surface of
the bypass duct may be defined by a nacelle and/or a fan case.
[0112] The overall pressure ratio of a turbofan gas turbine engine
as described and/or claimed herein may be defined as the ratio of
the stagnation pressure upstream of the fan to the stagnation
pressure at the exit of the highest-pressure compressor (before
entry into the combustor). By way of non-limitative example, the
overall pressure ratio of a gas turbine engine as described and/or
claimed herein in steady-state flight may be greater than (or on
the order of) any of the following: 10, 15, 20, 25, 30, 35 or 40.
The overall pressure ratio may be in an inclusive range bounded by
any two of the values in the previous sentence (i.e. the values may
form upper or lower bounds), for example in the range of from 20 to
35.
[0113] A turbofan gas turbine engine as described and/or claimed
herein may have any desired maximum thrust. Purely by way of
non-limitative example, a gas turbine as described and/or claimed
herein may be capable of producing a maximum thrust of at least (or
on the order of) any of the following: 20 kN, 40 kN, 60 kN, 80 kN,
100 kN, 120 kN, 140 kN, 160 kN, 180 kN, or 200 kN. The maximum
thrust may be in an inclusive range bounded by any two of the
values in the previous sentence (i.e. the values may form upper or
lower bounds). Purely by way of example, a gas turbine as described
and/or claimed herein may be capable of producing a maximum thrust
in the range of from 60 kN to 160 kN, for example 70 kN to 120 kN.
The thrust referred to above may be the maximum net thrust at
standard atmospheric conditions at sea level plus 15 degrees C.
(ambient pressure 101.3 kPa, temperature 30 degrees C.), with the
engine static.
[0114] A fan blade and/or aerofoil portion of a fan blade described
and/or claimed herein may be manufactured from any suitable
material or combination of materials. For example, at least a part
of the fan blade and/or aerofoil may be manufactured at least in
part from a composite, for example a metal matrix composite and/or
an organic matrix composite, such as carbon fibre. By way of
further example at least a part of the fan blade and/or aerofoil
may be manufactured at least in part from a metal, such as a
titanium-based metal or an aluminium-based material (such as an
aluminium-lithium alloy) or a steel-based material. The fan blade
may comprise at least two regions manufactured using different
materials. For example, the fan blade may have a protective leading
edge, which may be manufactured using a material that is better
able to resist impact (for example from birds, ice or other
material) than the rest of the blade. Such a leading edge may, for
example, be manufactured using titanium or a titanium-based alloy.
Thus, purely by way of example, the fan blade may have a
carbon-fibre or aluminium based body (such as an aluminium lithium
alloy) with a titanium leading edge.
[0115] A fan as described and/or claimed herein may comprise a
central portion, from which the fan blades may extend, for example
in a radial direction. The fan blades may be attached to the
central portion in any desired manner. For example, each fan blade
may comprise a fixture which may engage a corresponding slot in the
hub (or disc). Purely by way of example, such a fixture may be in
the form of a dovetail that may slot into and/or engage a
corresponding slot in the hub/disc in order to fix the fan blade to
the hub/disc. By way of further example, the fan blades maybe
formed integrally with a central portion. Such an arrangement may
be referred to as a bladed disc or a bladed ring. Any suitable
method may be used to manufacture such a bladed disc or bladed
ring. For example, at least a part of the fan blades may be
machined from a block and/or at least part of the fan blades may be
attached to the hub/disc by welding, such as linear friction
welding.
[0116] The turbofan gas turbine engines described and/or claimed
herein may or may not be provided with a variable area nozzle
(VAN). Such a variable area nozzle may allow the exit area of the
bypass duct to be varied in use. The general principles of the
present disclosure may apply to engines with or without a VAN.
[0117] The fan stage of a turbofan gas turbine engine as described
and/or claimed herein may have any desired number of fan blades,
for example 12, 14, 16, 18, 20, 22, 24, 26, 28, 30, 32, 34, or 36
fan blades.
[0118] According to an aspect of the disclosure, there is provided
an aircraft comprising a turbofan gas turbine engine as described
and/or claimed herein. The aircraft according to this aspect is the
aircraft for which the gas turbine engine has been designed to be
attached. Accordingly, the steady-state flight conditions according
to this aspect correspond to the steady-state flight conditions of
the aircraft, as defined elsewhere herein.
[0119] According to an aspect of the disclosure, there is provided
a method of operating a turbofan gas turbine engine as described
and/or claimed herein. The operation may be at the steady-state
flight conditions as defined elsewhere herein (for example in terms
of the thrust, atmospheric conditions and Mach Number).
[0120] According to an aspect of the disclosure, there is provided
a method of operating an aircraft comprising a turbofan gas turbine
engine as described and/or claimed herein. The operation according
to this aspect may include (or may be) operation at the
steady-state flight conditions of the aircraft, as defined
elsewhere herein.
[0121] The skilled person will appreciate that except where
mutually exclusive, a feature or parameter described in relation to
any one of the above aspects may be applied to any other aspect.
Furthermore, except where mutually exclusive, any feature or
parameter described herein may be applied to any aspect and/or
combined with any other feature or parameter described herein.
[0122] Other aspects of the disclosure provide devices, methods and
systems which include and/or implement some or all of the actions
described herein. The illustrative aspects of the disclosure are
designed to solve one or more of the problems herein described
and/or one or more other problems not discussed.
BRIEF DESCRIPTION OF THE DRAWINGS
[0123] There now follows a description of an embodiment of the
disclosure, by way of non-limiting example, with reference being
made to the accompanying drawings in which:
[0124] FIG. 1 shows a schematic perspective view of an aircraft
according to a first embodiment of the disclosure;
[0125] FIG. 2 shows a schematic sectional view of the aircraft of
FIG. 1;
[0126] FIG. 3 shows a schematic view of the heat exchanger module
and fan module of the turbofan gas turbine engine of the aircraft
of FIG. 1;
[0127] FIG. 4 shows a schematic sectional view of the heat
exchanger module and fan module of the turbofan gas turbine engine
of the aircraft of FIG. 1; and
[0128] FIG. 5 shows an alternative arrangement of the heat
exchanger module and fan module shown in FIG. 4.
[0129] It is noted that the drawings may not be to scale. The
drawings are intended to depict only typical aspects of the
disclosure, and therefore should not be considered as limiting the
scope of the disclosure. In the drawings, like numbering represents
like elements between the drawings.
DETAILED DESCRIPTION
[0130] Referring to FIGS. 1 and 2, an aircraft according to a first
embodiment of the disclosure is designated generally by the
reference numeral 100. The aircraft 100 comprises a machine body
102 in the form of a fuselage with wings and a tail plane. The
machine body 102 encloses a turbofan gas turbine engine 110,
together with a plurality of ancillary systems 104.
[0131] Within the machine body 102 there is a cockpit volume 106, a
payload volume 108, and a plurality of ancillary systems 104.
[0132] The turbofan gas turbine engine 110 comprises, in axial flow
sequence, a heat exchanger module 120, a fan module 130, a
compressor module 140, a combustor module 150, a turbine module
160, and an exhaust module 170. The turbofan gas turbine engine 110
further comprises a second heat transfer element 124. The second
heat transfer element 124 takes the form of a heat exchanger that
uses the engine's fuel as a cooling medium.
[0133] The fan module 130, compressor module 140, combustor module
150, turbine module 160, and exhaust module 170, together forming
the core engine, are enclosed within an outer casing 180. An
annular bypass duct 182 is defined between the core engine and the
outer casing 180.
[0134] The heat exchanger module 120 comprises a plurality of first
heat exchanger elements 122. In the present arrangement,
illustrated in FIG. 3, the first heat exchanger elements 122 are
arranged as a circumferential array of radially extending vanes
122. The inlet air flow 101 passes over the surface of the first
heat exchanger elements 122 as the air flow passes through the heat
exchanger module 120.
[0135] The heat exchanger module 120 has a total heat rejection
capacity. The total heat rejection capacity is the amount of waste
heat energy that can be dissipated into an air flow passing through
the heat exchanger module 120.
[0136] The machine body 102 comprises only one fluid inlet aperture
112. The fluid inlet aperture 112 is configured to allow an intake
air flow 101 to enter the machine body 102. In other words, there
is only one inlet aperture 112 on the machine body 102 through
which an air flow 101 can enter the machine body 102.
[0137] The intake air flow 101 passes through the heat exchanger
module 120 and subsequently passes through the fan module 130. Once
through the fan module, the air flow divides into a first flow (not
shown) and a second flow (not shown). The first flow (the `core`
flow) passes sequentially through the core engine, i.e.
sequentially through compressor module 140, the combustor module
150, the turbine module 160, and the exhaust module 170. The second
flow (the `bypass` flow) exits the fan module 130 and passes
through the annular bypass duct 182 to the exhaust module 170.
[0138] The machine body 102 further comprises only one fluid
exhaust aperture 104. The air flow from the exhaust module 170
exits the machine body 102 through the single fluid exhaust
aperture 104. In other words, there is only one exhaust aperture
104 in the machine body 102 through which an air flow 101 can exit
the machine body 102.
[0139] As outlined above, the machine body 102 of present
disclosure includes only two apertures 112,114 in its outer
surface; an inlet aperture 112 allowing an air flow into the
machine body and an exhaust aperture 114 allowing the air flow to
exhaust from the machine body. The presence of apertures in the
machine body 102 causes parasitic aerodynamic drag on the machine
body 102. As in the present arrangement, the use of only two
apertures 112,114 in the machine body 102 minimises this parasitic
aerodynamic drag.
[0140] In the present arrangement, the fan assembly 130 comprises
two fan stages (not shown), with each fan stage comprising a
plurality of fan blades (not shown). In the present arrangement
each fan stage has the same fan diameter 132, with the respective
plurality of fan blades defining a fan diameter of 0.9 m. In an
alternative arrangement, the two fan stages may have different fan
diameters 132 each defined by the corresponding plurality of fan
blades. As previously mentioned, the fan diameter (D) 132 is
defined by a circle circumscribed by the leading edges of the
respective plurality of fan blades.
[0141] In use, both the turbofan gas turbine engine 110 and the
ancillary systems 104 generate waste heat energy that is required
to be dissipated to ensure the safe operation of the turbofan
engine 110 and the ancillary systems 104.
[0142] As shown in FIG. 3, the heat exchanger module 120 has a flow
area (A.sub.HEX) 126. The heat exchanger module flow area 126 is
the cross-sectional area of the heat exchanger module 120 through
which an inlet air flow 101 passes before being ingested by the fan
module 130. In the present arrangement, the heat exchanger module
flow area 126 has an annular cross-section and corresponds directly
to the shape of the air flow passing through the heat exchanger
module 120.
[0143] The fan module 130 has a corresponding flow area (A.sub.FAN)
134. The fan module flow area 134 is the cross-sectional area of
the fan module 130 through which an inlet air flow 101 passes
before separating into the core engine flow and the bypass flow.
The fan assembly flow area 134 has an annular shape since it
corresponds to the annular area swept by the fan blades 131.
[0144] In the present arrangement (illustrated in FIG. 4) the heat
exchanger module flow area 126 is equal to the fan module flow area
134, and the corresponding ratio of A.sub.HEX/A.sub.FAN is equal to
1.0.
[0145] The heat exchanger module 120 has a flow diameter (E) 128,
which is the diameter of the air flow passing through the heat
exchanger module 120. In the present arrangement, shown in FIG. 4,
the heat exchanger module flow diameter 128 is equal to the fan
diameter 132. In an alternative arrangement (see FIG. 5) the heat
exchanger module flow diameter 228 is greater than the fan diameter
132.
[0146] The heat exchanger module 120 is configured to transfer a
waste heat load from the gas turbine engine 110 and the ancillary
systems 104 to the fluid flow 101 prior to the entry of the fluid
flow 101 into the fan module 130. A first fluid 116, which in this
embodiment is a synthetic oil, is circulated through hot portions
of the turbofan engine 110 and the ancillary systems 104 to collect
waste heat energy.
[0147] As outlined above, once the air flow 101 has passed through
the fan module 130, the air flow 101 divides into two flow
portions, a first portion being the so-called `core flow`, and a
second portion being the so-called `bypass flow`. The core flow
enters the compressor module 140 and continues sequentially through
the combustor module 150, turbine module 160, and exhaust module
170. The bypass flow passes through the annular bypass duct 182
into the exhaust module 170. The core flow and the bypass flow join
in the exhaust module 170 and are exhausted from the machine body
102 through the exhaust aperture 114.
[0148] The core flow can be characterised by the parameter
B.sub.COMB which represents the proportion of the fluid flow 101
entering the machine body 102 that subsequently passes sequentially
through the compressor, combustor, turbine and exhaust modules
140,150,160,170. In the present arrangement, the turbofan engine
110 has a bypass ratio of 2. In this arrangement, the turbofan
engine 110 can be characterised by a B.sub.COMB parameter of
0.29.
[0149] In use, at a cruise condition the aircraft 100 is capable of
maintaining a sustained airspeed V (in metres per second, m/s). A
definition for the cruise condition has been provided earlier in
the disclosure. At this sustained airspeed V, the heat exchanger
module 120 transfers a total waste heat energy load H (in Watts, W)
to the fluid flow 101.
[0150] In use, when the sustained airspeed V of the aircraft 100 is
less than Mach 1.0 (i.e. the aircraft 100 is in subsonic flight)
the first fluid 116 is circulated through the first heat exchanger
elements 122 to dissipate the waste heat energy contained in the
first fluid 116 to the inlet fluid flow 101.
[0151] When the sustained airspeed V of the aircraft 100 exceeds,
for example, Mach 1.0 (i.e. supersonic flight conditions) the
temperature of the inlet fluid flow (T.sub.A) 101 increases. This
temperature increase will significantly reduce the efficiency of
the transfer of the waste heat energy from the first fluid 116 to
the inlet fluid flow 101. Continued increase in the sustained
airspeed V of the aircraft 100 will cause a continued rise in the
temperature of the inlet fluid flow 101. Once this temperature
T.sub.A reaches the temperature of the first fluid (T.sub.F) 116 it
will not be possible to dissipate waste heat energy to the inlet
fluid flow 101 via the first heat transfer elements 122.
[0152] Consequently, in the arrangement of the present invention,
when the airflow temperature T.sub.A is equal to or greater than
the first fluid temperature T.sub.F, the flow of the first fluid
116 is routed through the second heat transfer element 124. The
second heat transfer element 124 uses the fuel supply to the
turbofan engine 110 as the cooling medium.
[0153] While the temperature of the inlet fluid flow T.sub.A 101
will increase at and above a sustained airspeed of, for example,
M1.0, the temperature of the engine's fuel will remain
substantially constant. By routing the first fluid 116 through the
second heat transfer element 124 it becomes possible to continue to
dissipate the waste heat energy from the turbofan engine 110 and
the ancillary systems 104 even when the temperature of the inlet
fluid flow 101 is greater than the temperature of the first fluid
116.
[0154] Note that the terms "low-pressure turbine" and "low-pressure
compressor" as used herein may be taken to mean the lowest pressure
turbine stages and lowest pressure compressor stages (i.e. not
including the fan 23) respectively and/or the turbine and
compressor stages that are connected together by the
interconnecting shaft 26 with the lowest rotational speed in the
engine. In some literature, the "low-pressure turbine" and
"low-pressure compressor" referred to herein may alternatively be
known as the "intermediate-pressure turbine" and
"intermediate-pressure compressor". Where such alternative
nomenclature is used, the fan 23 may be referred to as a first, or
lowest pressure, compression stage.
[0155] Other gas turbine engines to which the present disclosure
may be applied may have alternative configurations. For example,
such engines may have an alternative number of compressors and/or
turbines and/or an alternative number of interconnecting shafts.
Whilst the described example relates to a turbofan engine, the
disclosure may apply, for example, to any type of gas turbine
engine, such as an open rotor (in which the fan stage is not
surrounded by a nacelle) or turboprop engine, for example.
[0156] The geometry of the gas turbine engine 10, and components
thereof, is defined by a conventional axis system, comprising an
axial direction (which is aligned with the rotational axis 9), a
radial direction (in the bottom-to-top direction in FIG. 1), and a
circumferential direction (perpendicular to the page in the FIG. 1
view). The axial, radial and circumferential directions are
mutually perpendicular.
[0157] It will be understood that the invention is not limited to
the embodiments above-described and various modifications and
improvements can be made without departing from the concepts
described herein. Except where mutually exclusive, any of the
features may be employed separately or in combination with any
other features and the disclosure extends to and includes all
combinations and sub-combinations of one or more features described
herein.
[0158] The invention includes methods that may be performed using
the subject devices. The methods may comprise the act of providing
such a suitable device. Such provision may be performed by the end
user. In other words, the "providing" act merely requires the end
user obtain, access, approach, position, set-up, activate, power-up
or otherwise act to provide the requisite device in the subject
method. Methods recited herein may be carried out in any order of
the recited events which is logically possible, as well as in the
recited order of events.
[0159] In addition, where a range of values is provided, it is
understood that every intervening value, between the upper and
lower limit of that range and any other stated or intervening value
in that stated range, is encompassed within the invention.
[0160] Except where mutually exclusive, any of the features may be
employed separately or in combination with any other features and
the disclosure extends to and includes all combinations and
sub-combinations of one or more features described herein.
* * * * *