U.S. patent number 3,640,638 [Application Number 05/050,100] was granted by the patent office on 1972-02-08 for axial flow compressor.
This patent grant is currently assigned to Rolls-Royce Limited. Invention is credited to Jack Britt, Alfred John Honey.
United States Patent |
3,640,638 |
Britt , et al. |
February 8, 1972 |
AXIAL FLOW COMPRESSOR
Abstract
An axial flow compressor comprises a rotor having a plurality of
axially spaced rings of rotor blades, a respective ring of stator
blades being disposed immediately downstream of each ring of rotor
blades, respective shroud surfaces being provided at the radially
inner ends of the working surfaces of the stator blades by
respective shroud rings carried by the rings of stator blades,
running clearances being provided between the shroud rings and the
rotor, and means whereby in operation a static pressure gradient is
maintained in the clearance between at least one shroud ring and
the rotor so that gas flow through the clearance is in a downstream
direction.
Inventors: |
Britt; Jack (Ambergate,
EN), Honey; Alfred John (Allestree, EN) |
Assignee: |
Rolls-Royce Limited (Derby,
Derbyshire, EN)
|
Family
ID: |
10352013 |
Appl.
No.: |
05/050,100 |
Filed: |
June 26, 1970 |
Foreign Application Priority Data
|
|
|
|
|
Jul 2, 1969 [GB] |
|
|
33,367/69 |
|
Current U.S.
Class: |
415/173.7;
415/199.5; 415/171.1; 415/914 |
Current CPC
Class: |
F04D
29/083 (20130101); F04D 29/681 (20130101); F04D
29/164 (20130101); F01D 11/001 (20130101); Y10S
415/914 (20130101) |
Current International
Class: |
F04D
29/66 (20060101); F01D 11/00 (20060101); F04D
29/68 (20060101); F04D 29/08 (20060101); F04D
29/16 (20060101); F01d 011/08 (); F04d
029/00 () |
Field of
Search: |
;415/121,168,DIG.1,171,172,170,199 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Raduazo; Henry F.
Claims
We claim:
1. An axial flow compressor comprising a rotor having a plurality
of axially spaced rings of rotor blades each having working
surfaces and a root, said roots defining a rotor surface, a
respective ring of stator blades being disposed immediately
downstream of each ring of rotor blades, each of said stator blades
having working surfaces and radially inner ends, each of said
radially inner ends carrying a shroud ring, said shroud rings
defining respective shroud surfaces, running clearances being
provided between said shroud rings and said rotor, a said rotor
surface being disposed immediately upstream of at least one of a
said shroud ring, means for maintaining a static pressure gradient
in the clearance between at least said one shroud ring and said
rotor whereby gas flow through said clearance will be in a
downstream direction, said means including said rotor surface
having a trailing edge and said shroud surface of said at least one
shroud ring having a leading and trailing edge, said trailing edge
of said rotor surface having a smaller diameter than said leading
edge of said shroud surface, said leading edge of the said shroud
surface projecting radially outwardly of the said trailing edge of
said rotor surface so that the boundary layer flow of gas passing
from said trailing edge of said rotor surface will be intercepted
by said leading edge of said shroud surface to increase the static
pressure of the boundary layer flow passing to said leading edge of
said shroud surface relative to the boundary layer static pressure
at said trailing edge of said shroud surface of the at least one
shroud ring.
2. A compressor as claimed in claim 1 wherein a ring of rotor
blades immediately downstream of the at least one shroud ring has a
further rotor surface at the roots of the working surfaces of the
blades, the leading edge of the rotor surface being of smaller
diameter than the trailing edge of the said shroud surface, said
trailing edge projecting radially outwardly of the said leading
edge of the further rotor surface.
3. A compressor as claimed in claim 1 wherein sealing means are
provided in said clearance between the at least one shroud ring and
the rotor, reducing the gas flow therethrough.
4. A compressor as claimed in claim 3 wherein the rotor has an
annular pocket between the ring of rotor blades immediately
upstream of the at least one shroud ring and a ring of rotor blades
immediately downstream of the at least one shroud ring, the at
least one shroud ring being received with said clearance in the
annular pocket, said sealing means being provided between a wall of
the pocket and the shroud ring.
5. A gas turbine engine having an axial flow compressor comprising
a rotor having a plurality of axially spaced rings of rotor blades
each having working surfaces and a root, said roots defining a
rotor surface, a respective ring of stator blades being disposed
immediately downstream of each ring of rotor blades, each of said
stator blades having working surfaces and radially inner ends, each
of said radially inner ends carrying a shroud ring, said shroud
rings defining respective shroud surfaces, running clearances being
provided between said shroud rings and said rotor, a said rotor
surface being disposed immediately upstream of at least one of a
said shroud ring, means for maintaining a static pressure gradient
in the clearance between at least said one shroud ring and said
rotor whereby gas flow through said clearance will be in a
downstream direction, said means including said rotor surface
having a trailing edge and said shroud surface of said at least one
shroud ring having a leading and trailing edge, said trailing edge
of said rotor surface having a smaller diameter than said leading
edge of said shroud surface, said leading edge of the said shroud
surface projecting radially outwardly of the said trailing edge of
said rotor surface so that the boundary layer flow of gas passing
from said trailing edge of said rotor surface will be intercepted
by said leading edge of said shroud surface to increase the static
pressure of the boundary layer flow passing to said leading edge of
said shroud surface relative to the boundary layer static pressure
at said trailing edge of said shroud surface of the at least one
shroud ring.
Description
This invention relates to axial flow compressors.
The compression ratio and therefore the delivery gas temperature
achieved in axial flow compressors is steadily increasing and is
now such that nickel base alloys have to be considered for the
compressor rotors. We have now discovered that this problem is
aggravated because the actual temperature reached by the structure
of a stage of an axial flow compressor is greater than the mean
temperature of the gas as it passes through that stage.
We believe that a reason for this anomaly is the recirculation of
gas via the running clearances between the stator shroud rings and
the rotor, so that the compressor stage is heated by hotter gas
recirculated from a downstream stage.
Therefore, in one aspect the invention provides an axial flow
compressor comprising a rotor having a plurality of axially spaced
rings of rotor blades, a respective ring of stator blades being
disposed immediately downstream of each ring of rotor blades,
respective shroud surfaces being provided at the radially inner
ends of the working surfaces of the stator blades by respective
shroud rings carried by the rings of stator blades, running
clearances being provided between the shroud rings and the rotor,
and means whereby in operation a static pressure gradient is
maintained in the clearance between at last one shroud ring and the
rotor so that gas flow through the clearance is in a downstream
direction.
The said means may be adapted to maintain the static pressure
gradient by increasing the static pressure of a boundary layer flow
passing to the leading edge of the shroud surface of the at least
one shroud ring from a rotor surface at the roots of the working
surfaces of the ring of rotor blades immediately upstream of the at
least one shroud ring, so that said boundary layer static pressure
is greater than the boundary layer static pressure at the trailing
edge of the shroud surface of the at least one shroud ring.
The said means may be adapted to effect said increase in boundary
layer static pressure by converting at least a portion of the
dynamic pressure of the boundary layer at the said upstream edge to
static pressure.
The trailing edge of the said rotor surface may be of smaller
diameter than the leading edge of the shroud surface of the at
least one shroud ring, so that in operation the leading edge of the
said shroud surface projects radially outwardly of the said
trailing edge of the rotor surface and intercepts the boundary
layer flow passing therefrom.
A trailing edge of the shroud surface of the at least one shroud
ring may be of greater diameter than that of a leading edge of a
further rotor surface at the roots of the working surfaces of a
ring of rotor blades immediately downstream of the at least one
shroud ring so that in operation the trailing edge of the said
shroud surface projects radially outwardly of the said leading edge
of the further rotor surface.
Sealing means may be provided in said clearance between the at
least one shroud ring and the rotor to reduce the gas flow
therethrough.
The at least one shroud ring may be received with said clearance in
an annular pocket in the rotor between the ring of rotor blades
immediately upstream thereof and a ring of rotor blades immediately
downstream thereof, said sealing means being provided between a
wall of the pocket and the shroud ring.
In another aspect, although not so restricted, the invention
provides a gas turbine engine having an axial flow compressor as
set forth above.
The invention will be described, merely by way of example, with
reference to the accompanying drawings, wherein:
FIG. 1 shows, partly in section, a gas turbine engine embodying a
compressor according to the invention, and
FIG. 2 shows a part of the structure of FIG. 1.
Referring to FIG. 1, a gas turbine engine comprises an inlet 10, an
axial flow compressor 12, a combustion section 14, a turbine
section 16 and an exhaust nozzle 18.
Attention is now directed to FIG. 2, which although showing a
compressor according to the invention will initially be used to
illustrate the above-mentioned problem as it occurs in conventional
compressors.
Conventionally, axial flow compressors comprise a rotor 20 having a
plurality of axially spaced rings of rotor blades two of which
rings are shown at 22, 24. A respective ring of stator blades is
disposed downstream of each ring of rotor blades. Such a ring of
stator blades is shown at 26. A shroud ring 28 is carried by the
ring of stator blades 26 at the radially inner ends of the working
surfaces of the stator blades. A running clearance is provided
between the shroud ring 28 and the rotor 20; in the case of shroud
rings 28 of intermediate stator blade rings such as 26, the running
clearance typically is provided by means of an annular pocket 30
between the rings of rotor blades 22, 24 respectively immediately
upstream and downstream of the ring of stator blades 26.
Surfaces (e.g., 32, 33) are provided on the rotor 20 at the roots
of the working surfaces of each ring of rotor blades 22, 24, and
the shroud rings 28 are provided with shroud surfaces 34, the
surfaces 32, 33, 34 defining the radially inner wall of the gas
flow duct through the compressor.
The function of stator blades in an axial flow compressor of course
is to convert part of the dynamic head of the gas flow leaving the
preceding rotor blades to a static head, whereby the pressure of
the gas is increased successively from stage to stage. By virtue of
the function of the stator blades, the static pressure of the gas
at the trailing edges 36 of the ring of stator blades 26 is greater
than the static head at the leading edges 38 thereof. The gas
temperature is also greater since the conversion to static pressure
of a dynamic head due to a gas velocity of V results in an increase
in gas temperature of V.sup.2 / 2C.sub.p J.sub.g where C.sub.p is
the specific heat at constant pressure of the gas, J is the
mechanical equivalent of heat and g is the acceleration due to
gravity.
We believe that the boundary layer flow across the surfaces 32, 34
suffers increases in temperature in the following way. The boundary
layer flow approaching the leading edge of the surface 32 from an
upstream stator shroud ring (not shown) has a velocity which is
markedly different from the free stream velocity for which the
blades of the rotor stage 22 are designed. Thus, the boundary layer
suffers a violent change in velocity as it reaches the surface 32,
resulting in an increase in temperature of the boundary layer.
The boundary layer flow then passes over the surface 32, and during
this time has a large circumferential component of velocity and a
small axial component of velocity. As the boundary layer flow
leaves the trailing edge of the surface 32 it passes to the leading
edge 40 of the stationary shroud surface 34, and since its velocity
is again markedly different from the free stream velocity for which
the stator blades 26 are designed, due to its small axial
component, the boundary layer suffers another violent change of
velocity, and its temperature is further increased.
Upon leaving the trailing edge 39 of the shroud surface 34, the
boundary layer flow impinges upon the leading edge of the surface
33, as described above in relation to the surface 32, and is thus
subjected to a yet further violent change in velocity and
consequent rise in temperature. The rise in boundary layer
temperature caused in this way may be equivalent to as many as
three dynamic pressures.
The static pressure difference between the trailing 36 and leading
38 edges of the ring of stator blades 26 diverts a portion of the
hot boundary layer leaving the trailing edge 39 through the
clearance 30 to the leading edge 40. This recirculated portion of
hot boundary layer flow is discharged at the leading edge 40 and
thus is again subjected to the violent velocity changes at the
edges 40, 39, and is further raised in temperature.
Some of the boundary layer flow even may be recirculated for a
second time, further raising its temperature.
Considering now the novel features of FIG. 2, we provide means
whereby in operation of the compressor a static pressure gradient
is maintained in the clearance 30 so that leakage gas flow is in a
downstream direction, as indicated by the arrows 42. The static
pressure gradient is maintained by increasing the static pressure
of the boundary layer flow passing from the trailing edge 44 of the
rotor surface 32 to the leading edge 40 of the shroud surface 34,
so that the boundary layer static pressure is greater at the
leading edge 40 than at the trailing edge 39 of the shroud surface
34.
To effect this increase in boundary layer static pressure, the
diameter of the rotor surface 32 at its trailing edge 44 is chosen
to be smaller than the diameter of the leading edge 40 of the
shroud surface 34, so that when the compressor is operating at its
normal speed the leading edge 40 projects radially outwardly of the
trailing edge 44 and intercepts the boundary layer flow passing
therefrom. The extent of the projection is indicated by the
dimension referenced 46. This projection 46 converts a sufficient
proportion of the dynamic pressure of the boundary layer flow to
static pressure to maintain the required pressure gradient.
Thus, the boundary layer flow leaving the surface 32 still is
subjected to a temperature rise due to the reduction in velocity
when it impinges upon the leading edge 40, but at least some of the
boundary layer flow then passes in a downstream direction through
the clearance 30. While passing through the clearance 30, the flow
tends to be entrained gradually by the rotating walls of the pocket
defining the clearance 30. Thus, the flow emerging from the
clearance 30 at the downstream end thereof has a substantial
circumferential component of velocity, and the velocity change when
the flow impinges on the leading edge of the surface 33 may be less
violent and may cause a smaller rise in temperature.
The novel construction of FIG. 2 may result in a reduction in the
rise in temperature of the boundary layer between the rings of
rotor blades 22, 24 equivalent to at least one dynamic
pressure.
A seal 48 is provided in the clearance 30 between the walls of the
pocket defining the clearance and the shroud ring 28, since it is
desirable to reduce the flow through the clearance 30. An
appreciable flow through the clearance 30 in the downstream
direction would disturb the gas velocities between the ring of
stator blades 26 and the next ring of rotor blades 24, reducing
compressor efficiency.
The trailing edge 39 of the shroud surface 34 is of greater
diameter than the leading edge 50 of the rotor surface 33 so that
in operation it projects radially outwardly thereof, as evidenced
by the dimension 52. The boundary layer gas flow leaving the
trailing edge 39 thus possibly may have an extractor of `jet pump`
effect on the gas in the clearance 30, thus augmenting the pressure
gradient in the clearance 30.
It will be appreciated that a compressor according to the invention
need not have a pressure gradient maintained to produce a
downstream flow in every running clearance between stator blade
shroud rings and the rotor; for example, it may be sufficient to
provide a pressure gradient only in one or more clearances, e.g.,
around the hottest stator blades shroud rings at the delivery end
of the compressor.
* * * * *