U.S. patent application number 10/717590 was filed with the patent office on 2005-01-06 for compressor for a gas turbine engine.
This patent application is currently assigned to ROLLS-ROYCE PLC. Invention is credited to Tonks, Robert C..
Application Number | 20050002781 10/717590 |
Document ID | / |
Family ID | 9948905 |
Filed Date | 2005-01-06 |
United States Patent
Application |
20050002781 |
Kind Code |
A1 |
Tonks, Robert C. |
January 6, 2005 |
Compressor for a gas turbine engine
Abstract
A compressor for a gas turbine engine comprising a casing, a
mixed flow stage including a first rotor and at least one further
mixed or centrifugal flow stage including a second rotor axially
spaced apart from the first rotor, and bearing means positioned
between the first and second rotors for rotatably mounting the
rotors with respect to the casing. The bearing means is positioned
closely adjacent the first rotor. The rotors are carried on a
common shaft and the bearing means rotatably mounts the shaft with
respect to a bearing support which extends radially with respect to
the shaft between the bearing and the casing. The bearing support
comprises a frusta-conical portion extending from the said bearing
means towards the said casing.
Inventors: |
Tonks, Robert C.;
(Bridgwater, GB) |
Correspondence
Address: |
OLIFF & BERRIDGE, PLC
P.O. BOX 19928
ALEXANDRIA
VA
22320
US
|
Assignee: |
ROLLS-ROYCE PLC
London
GB
|
Family ID: |
9948905 |
Appl. No.: |
10/717590 |
Filed: |
November 21, 2003 |
Current U.S.
Class: |
415/199.1 |
Current CPC
Class: |
F02C 3/04 20130101; F04D
17/122 20130101; F02C 7/06 20130101; F04D 29/056 20130101; F04D
29/057 20130101; F04D 29/0566 20130101; F04D 29/051 20130101; F04D
17/06 20130101 |
Class at
Publication: |
415/199.1 |
International
Class: |
F01D 001/02 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 3, 2002 |
GB |
0228050.1 |
Claims
1. A compressor for a gas turbine engine; the compressor comprising
a casing, a bearing means, a bearing support means, a mixed flow
stage including a first rotor and at least one further mixed or
centrifugal flow stage including a second rotor axially spaced
apart from the said first rotor; all of the said rotors being
carried on a common shaft rotatably mounted by the bearing means
with respect to the bearing support means, the said bearing means
being positioned between said first and second rotors for rotatably
mounting the said rotors with respect to the said casing.
2. A compressor as claimed in claim 1 wherein the said bearing
means is positioned closely adjacent the said first rotor.
3. A compressor as claimed in claim 1 wherein the said bearing
support means extends radially with respect to the said shaft
between the said bearing means and the said casing.
4. A compressor as claimed in claim 3 wherein the said bearing
support means comprises a frusta-conical portion extending from the
said bearing means towards the said casing.
5. A compressor as claimed in claim 1 further comprising at least
one diffuser section between the said rotors and wherein the said
bearing support means is connected to the said casing through the
said diffuser section.
6. A compressor as claimed in claim 5 wherein the said diffuser
section has an inlet section and an outlet section with a radially
inward flow section there between.
7. A compressor as claimed in claim 1 wherein the bearing means
comprises a journal bearing.
8. A compressor as claimed in claim 7 wherein the bearing means
additionally comprises a thrust bearing.
9. A compressor as claimed in claim 1 wherein the said compressor
is a two stage mixed flow compressor having two mixed flow
stages.
10. A compressor as claimed in claim 1 wherein the said compressor
stages have substantially the same pressure ratio.
11. A compressor as claimed in claim 10 wherein the pressure ratio
of each stage is greater than 4:1.
12. A gas turbine engine comprising a compressor according to claim
1.
13. A gas turbine engine as claimed in claim 12 wherein the said
origin is a single spool engine.
14. A gas turbine engine as claimed in claim 13 wherein the said
engine comprises a turbo-fan or a turbo-jet engine.
15. A gas turbine engine as claimed in claim 12 wherein the said
engine is for an aircraft.
16. A gas turbine engine as claimed in claim 13 wherein the said
engine is for an aircraft.
17. A gas turbine engine as claimed in claim 14 wherein the said
engine is for an aircraft.
Description
[0001] This invention relates to a compressor for a gas turbine
engine
[0002] In particular a two stage compressor with a mixed flow first
stage and at least one further stage for a single spool gas turbine
engine.
[0003] Compressors for gas turbine engines include axial flow and
radial flow types both of which have their particular advantages
depending upon the specific engine application. For example, it is
possible to reduce the number of compressor stages, and hence cost,
by the use of a single centrifugal flow stage instead of a greater
number of axial flow stages. Stage for stage much higher pressure
ratios are possible with centrifugal compressors than axial flow
machines and centrifugal flow compressors have greater resistance
to foreign object damage by ingestion of objects in the intake air
stream. Centrifugal flow compressors are more commonly found in
small turboshaft type engines where resistance to such damage can
be a highly relevant design consideration, for example in
helicopter applications. Axial flow compressors on the other hand
offer greater operational efficiency at the expense of more
compressor stages, and hence components and cost, for a given
pressure ratio. Axial flow compressors also have a lower frontal
area than centrifugal flow compressors which can be an important
consideration in aircraft applications. The increased frontal area,
or envelope, associated with radial flow compressors is due in part
to the dimensions of the radial flow impeller and also the
requirement to position an annular diffuser radially outwards of
the impeller.
[0004] The conflicting requirements of compressor operational
efficiency and reduced number of compressor stages has been partly
addressed by so-called mixed flow stages in which the flow through
the compressor stage has both an axial and radial component so that
the stage functions partly as a radial flow stage and partly as an
axial flow stage. Mixed flow compressors offer a combination of the
performance benefits of axial flow and radial flow compressors.
[0005] A two-stage compressor with a mixed flow first stage and a
centrifugal flow second stage is disclosed in International Patent
Application Number PCT/CA01/01335, in which the mixed flow rotor
and centrifugal rotor are joined together on a common shaft which
is supported with respect to the housing or compressor casing by
means of a bearing assembly positioned at the forward (upstream
end) of the two stage compressor. In this arrangement the
compressor rotor stages may be considered to be cantilevered from
the bearing at the forward end of the compressor and this can lead
to problems in particular with rotor blade tip clearances. This
arrangement also has the disadvantage that it requires a bearing
support structure to be provided in the inlet region of the
compressor resulting in a reduced flow cross-section, or a greater
diameter inlet, or compressor frontal area, for a specific flow
cross sectional area.
[0006] There is a requirement therefore for a compressor having at
least a first mixed flow stage and a second mixed or centrifugal
flow stage which exhibits improved tip clearance characteristics,
and also a requirement for such a compressor having a reduced
diameter inlet for a specific flow cross sectional area.
[0007] According to an aspect of the invention there is provided a
compressor for a gas turbine engine; the compressor comprising a
casing, a bearing means, a bearing support means, a mixed flow
stage including a first rotor and at least one further mixed or
centrifugal flow stage including a second rotor axially spaced
apart from the said first rotor; all of the said rotors being
carried on a common shaft rotatably mounted by the bearing means
with respect to the bearing support means, the said bearing means
being positioned between said first and second rotors for rotatably
mounting the said rotors with respect to the said casing.
[0008] This arrangement has the particular advantage that the
bearing and its associated support structure can be positioned away
from the compressor inlet, at the rear face, that is to say the
downstream side, of the first rotor so that the axial distance
between the bearing and the rotor blades, or vanes, is
significantly less than in arrangements where the rotors are
effectively cantilevered from the front of the compressor as
previously described. In this way it is readily possible to
maintain an appropriate tip clearance between the compressor rotor
vanes and the compressor casing or other ducting forming the
radially outer wall of the annular gas flow passage of the
compressor. This arrangement has the additional benefit of reducing
the compressor inlet diameter for a specific inlet flow area since
the radially inner flow boundary of the inlet may be positioned
closer to the rotation axis of the compressor. In this aspect of
the invention the bearing and associated support structure is
positioned remote from the compressor inlet and placed between the
first and second stage rotors, that is to say between the front
face of the first stage rotor and the rear face of the second stage
rotor.
[0009] By positioning the bearing between the first and second
stage rotors the bearing support can be positioned downstream of
the first compressor stage well away from the inlet section where
it would reduce the flow inlet area for a particular diameter of
inlet.
[0010] In preferred embodiments the bearing means is positioned
closely adjacent to the first rotor. This arrangement has the
additional benefit of readily enabling a frusto-conical bearing
support structure to be positioned between the first and second
stage rotors to support the bearing loads with respect to the
compressor casing. This arrangement has particular advantages where
the second stage comprises a centrifugal flow compressor stage
where radial tip clearances are of less significance than in the
first stage mixed flow impeller.
[0011] The bearing support means preferably extends radially with
respect to the compressor shaft between the bearing and the casing
so that the bearing loads can be readily transferred to the
compressor casing, and preferably the bearing support is
integrated, that is in the sense that it forms part of rather than
being integrally formed with, with other non rotating components
within the gas flow path of the compressor between the first and
second stage rotors.
[0012] In one embodiment the bearing support means comprises a
frusto-conical portion extending from the bearing towards the
casing. This is particularly advantageous when the bearing support
transfers axial thrust loads from the compressor shaft to the
casing. In particular embodiments an angle of 45.degree. may be an
appropriate angle for the conical portion.
[0013] Preferred embodiments comprise at least one diffuser section
between the first and second stage rotors. In arrangements where
such a diffuser section is provided the bearing support may be
connected to the compressor casing through the diffuser section,
that is to say through structural vane support struts extending
radially between the inner and outer radial peripheries of the
annular gas flow path of the compressor.
[0014] Preferably, the diffuser section has an inlet and an outlet
with a radially inward flow section therebetween. In this way it is
possible to redirect the flow at the exit of the first stage
radially inwards so that the inlet to the second stage rotor is at
a position radially inward of the first stage diffuser outlet. This
readily enables the diameter of the compressor to be reduced in the
region of the second stage at least, so that the frontal area of
the compressor, as determined by the compressor casing, can be
optimised, ie kept to a minimum.
[0015] In one particular embodiment the bearing comprises a journal
bearing and in other embodiments additionally or alternatively a
thrust bearing.
[0016] In one embodiment the compressor is a two staged mixed flow
compressor having two mixed flow stages. In such an embodiment the
compressor stages may have substantially equal pressure ratios
preferably greater than 4:1 so that the two stage compressor has an
overall pressure ratio of at least 16:1.
[0017] According to another aspect of the invention there is
provided a gas turbine engine comprising a compressor having a
mixed flow stage and at least one further mixed or centrifugal
stage according to the compressor as hereinbefore defined in
accordance with the above first mentioned aspect of the
invention.
[0018] In one preferred embodiment the gas turbine engine is a
single spool engine, for example a turbofan or turbojet single
spool engine. For the avoidance of doubt the term "single spool"
used herein refers to an engine configuration in which all the
turbine and compressor rotor stages are mounted on a common engine
shaft. Such an engine configuration is particularly suitable for
applications where a low cost engine is required, for example where
production and maintenance costs are a more significant
consideration than operating costs, for example where the engine
has a limited operational life. The gas turbine engine may be an
aero engine for a manned or unmanned aircraft as required.
[0019] An embodiment of the invention will now be more particularly
described, by way of example, with reference to the accompanying
drawing, which shows an axial cross-section view through part of a
single spool gas turbine engine comprising a compressor according
to one embodiment of the present invention.
[0020] Referring to the drawing a single spool gas turbine engine
comprises a two stage mixed flow compressor 10 including a first
stage 12 and a second stage 14 mounted within a generally
cylindrical compressor housing 16. The compressor 10 comprises part
of a single spool gas turbine engine, the other components of which
are not shown in the drawing but briefly comprise a combustor
downstream of the compressor and a turbine downstream of the
combustor and rotatably connected to the compressor rotor stages by
a common shaft 18. The shaft 18 is rotatably mounted with respect
to the compressor housing or casing 16 by a first bearing assembly
20 located in the compressor, and additionally a second bearing
assembly (not shown) downstream of the compressor, for example in
the turbine section of the engine. The compressor 10 has an inlet
22 through which air is inducted into the first stage 12, and an
outlet 24 through which high pressure air is delivered to the
engine combustor section (not shown) downstream of the
compressor.
[0021] The first compressor stage 12 comprises a rotor 26 which is
connected to the forward end of the shaft 18 and a diffuser 28. The
rotor 26 includes a disc part 30 which carries a plurality of
circumferentially spaced vanes 32 which pressurise the inducted air
in the annular gas flow passage defined between the rim 34 of the
disc and an interior wall part 36 of the gas flow passage. The
mixed flow rotor turns the inducted inlet air so that the air which
has a generally axial flow direction at the inlet has an outlet
flow direction which includes a radial as well as an axial
component. As it is well understood in the art of gas turbine
engine design the diffuser 28 functions to reduce the outlet
velocity of the gas flow exiting the rotor vanes 32 so that the
flow may be delivered at an appropriate velocity to a downstream
compressor stage. In the embodiment shown in the drawing the
diffuser 28 turns the gas flow exiting the rotor vanes 32 so that
it exits the diffuser at 38 having a substantially axial flow
direction.
[0022] The second mixed flow compressor stage 14 is substantially
the same as the first compressor stage in the sense that it
comprises a rotor 42 including a disc 44 and vanes 46 and a
diffuser section 48 immediately downstream of the rotor vanes 46.
The two mixed flow stages 12 and 14 are spaced apart along the
compressor axis 50 with the outlet 38 to the first stage diffuser
28 being connected to the second stage rotor inlet 52 by a further
duct 40 which extends between the two compressor stages, and to a
bypass duct 66 radially outwards of the duct 40. The inner and
outer annular surfaces of the duct 40 comprise inflection curves of
revolution such that the duct 40 turns the flow radially inwards
from the diffuser outlet 38 to the second stage inlet 52 which has
a smaller mean radius with respect to the compressor arcs 50 than
the outlet 38. The duct 40 preferably comprises an array of support
vanes 54 which extend radially through the gas flow passage of the
duct to provide a structural support between the casing 16 and a
bearing support assembly 56 which extends from the vane structure
54 to the bearing assembly 20. The support vanes 54 may extend only
partially along the duct 40 as shown in the drawing, or
alternatively extend along the full length of the duct between
first stage outlet 38 and second stage inlet 52. In the embodiment
shown in the drawing the vanes 54 are connected to the compressor
casing through the bypass duct 66 by corresponding radially
extending bypass support vanes 62.
[0023] The bearing support 56 includes a frusto-conical part 58 and
a cylindrical part 60 and two radial parts 62 and 64. The annular
structure defined by the bearing support parts 58, 60, 62 and 64
readily supports the shaft and rotor stages connected thereto and
provides for the transfer of engine loads from the rotatable
components to the compressor casing 16. The radial parts 62 and 64
schematically illustrate the axial extent of the vanes 54 when they
extend the full length of the duct 40.
[0024] In the drawing the bearing assembly 20 is positioned between
the compressor rotors 26 and 42 and in this particular embodiment
it is positioned towards the rear face, that is to say the
downstream side, of the first stage rotor 26. In other embodiments
the bearing assembly 20 may be positioned anywhere between the
rotors 26 and 42 without significantly affecting the tip clearance
control of the rotor vanes 32 with respect to the outer wall 36 of
the annular gas flow passage of the compressor.
[0025] Although aspects of the invention have been described with
reference to the embodiments shown in the accompanying drawing, it
is to be understood that the invention is not limited to this
precise embodiment and that various changes and modifications may
be effected without further inventive skill and effort. For
example, the second compressor stage 14 may comprise a centrifugal
flow compressor instead of a second mixed flow compressor. In
addition the gas turbine engine may comprise a turbofan or turbojet
engine. In alternative embodiments the bearing support structure 56
may be supported by the compressor casing 16 through the diffuser
section 28 rather than the duct 40. Moreover, one or more
additional bearing assemblies may be provided between the
compressor rotors 26 and 42 if additional support is required. The
invention also contemplates embodiments where further mixed flow or
centrifugal flow compressor stages are provided, that is to say
three, four, or more compressor stages.
* * * * *