TURBOCOMPRESSEUR A ECOULEMENT AXIAL

AXIAL FLOW TURBO COMPRESSOR

Abrégé français

L'invention concerne un turbocompresseur à écoulement axial comprenant une ou plusieurs parties stator espacées axialement (14) ayant chacune un réseau circonférentiel d'aubes directrices (15), et une ou plusieurs parties rotor (12, 16) ayant chacune un réseau circonférentiel d'ailettes de rotor (13, 17; A, B). Entre les ailettes de rotor successives (A, B), on trouve des chemins d'écoulement, tandis qu'entre les parties stator et rotor, on trouve des espaces axiaux (22, 23, 24) dont la configuration diverge, de façon à former dans chaque chemin d'écoulement, une zone de diffusion (C) qui s'étend de la section transversale la plus étroite (a¿2?) du chemin d'écoulement situé à une certaine distance en amont du bord d'attaque d'une seconde ailette de rotor (B) à une section transversale plus large (a¿3?) située approximativement sur le bord d'attaque de la seconde ailette de rotor (B), chaque chemin d'écoulement comprenant une zone de transition (D) d'une zone de section transversale n'augmentant sensiblement pas qui s'étend de la section transversale plus large (a¿3?) aux bords arrière des ailettes de rotor (A, B).

Abrégé anglais

An axial flow compressor including one or more axially spaced stator sections
(14) each having a circumferential array of guide vanes (15), and one or more
rotor section (12, 16) each having a circumferential array of rotor blades
(13, 17; A, B), and between successive rotor blades (A, B) there are formed
flow paths and between the stator and rotor sections there are provided axial
gaps (22, 23, 24) of a diverging configuration so as to form in each flow path
a diffusion region (C) which extends from a narrowest cross section (a2) of
the flow path located at a certain distance upstream of the leading edge of a
second rotor blade (B) to a wider cross section (a3) located approximately at
the leading edge of the second rotor blade (B), wherein each flow path has a
transition region (D) of a substantially non-increasing cross sectional area
extending from the wider cross section (a3) to the trailing edges of the rotor
blades (A, B).

Revendications

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Claims
1. Axial flow turbine compressor, comprising a stator
with at least one axial section including a circumferential
array of flow directing guide vanes (13,17), a rotor with
at least one axial section including a circumferential
array of rotor blades (15; A, B), an inner peripheral wall
(28), and an outer peripheral wall (29), wherein a flow
path is formed between every two successive rotor blades
(A, B) in the direction of rotor rotation (.omega.) and between
said inner and outer peripheral walls (28,29),
characterized in that each flow path comprises
.cndot. a narrowest cross sectional area at a first cross
section (a2) located at a certain distance upstream of
the leading edge of a second rotor blade (B) in the
direction of rotor rotation (.omega.),
.cndot. a diffusion region (C) having a successively increasing
cross sectional area in the flow direction and extending
from said first cross section (a2) to a second cross
section (a3) located approximately at the leading edge
of said second rotor blade (B), and
.cndot. a transition region (D) extending in the flow direction
from said second cross section (a3) to the trailing edge
of said second rotor blade (B), said transition region
(D) has a substantially non-increasing cross sectional
area in the flow direction throughout its length.
2. Turbine compressor according to claim 1, wherein
the transverse distance between said first and second rotor
blades (A, B) increases throughout said transition region
(D), whereas the radial distance between said inner
peripheral wall (28) and said outer peripheral wall (29)
decreases such that the cross sectional area of the flow
path does not increase throughout said transition region
(D).

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3. Turbine compressor according to claim 1 or 2,
wherein the flow through said diffusion region (C) is
substantially laminar.
4. Turbine compressor according to anyone of claims
1-3, wherein each flow path upstream of said diffusion
region (C) has a substantially constant cross sectional
area.
5. Axial flow type turbine compressor, comprising a
stator with at least one circumferential row of flow
directing guide vanes (13,17), and a rotor with at least
one circumferential row of rotor blades (15; A,B), wherein
between said guide vanes (13,17) and said rotor blades (15;
A,B) there are formed a number of parallel flow paths, and
between successive rotor blades (A, B) there are formed
rotor flow passages through which said flow paths extend,
characterized in that each flow path has a
diffusion region (C) extending from a location
approximately at the leading edge of a first one (A) of
said successive rotor blades (A, B) to a location
approximately at the leading edge of a second rotor blade
(B) in the rotation direction (.omega.) of said rotor, wherein
each one of said rotor flow passages downstream of said
diffusion region (C) has a substantially non-increasing
cross sectional area throughout its length.
6. Turbine compressor according to claim 5, wherein
said flow paths as well as said rotor flow passages are
partly defined by an outer peripheral wall (29) and an
inner peripheral wall (28).
7. Turbine compressor according to claim 5 or 6,
wherein each one of said rotor flow passages is defined by
a first rotor blade (A), and a second rotor blade (B)
succeeding said first rotor blade (A) in the direction of

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rotor rotation (.omega.), said diffusion region (D) extends in
the flow direction from a location approximately at the
leading edge of said first rotor blade (A) to a location
approximately at the leading edge of said second rotor
blade (B), and each one of said flow paths has a narrowest
cross sectional area (a2) at the upstream end of said
diffusion region (C).
8. Turbine compressor according to claim 1, wherein
said guide vanes (13,17) are arranged in two or more
axially spaced stator sections, and said rotor blades (15)
are arranged in two or more axially spaced rotor sections,
.cndot. said rotor sections and said stator sections are
arranged with an axial gap (22,23,24) between them,
.cndot. said axial gap (22,23,24) having an axial width of at
least 30% of the chord length of the preceding guide
vane (13,17) or rotor blade (15), and
.cndot. said axial gap (22,23,24) forms a flow passage region
with a radially diverging shape in the axial direction.
9. Turbine compressor according to claim 8, wherein
said flow passage region has a radial extent (h2) at its
downstream end that is at least 10% larger than the radial
extent (h1) at its upstream end.
10. Turbine compressor according to claim 8, wherein
said flow passage region has an axial width of at least 50%
of the chord length of the preceding guide vane (13,17) or
rotor blade (15).

Description

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The invention relates to a turbo compressor of the axial
flow type comprising a stator with at least one axial
section including a circumferential array of flow directing
guide vanes and a rotor with at least one axial section
including a Circumferential array of rotor blades, wherein
between the guide vanes and the rotor blades and between an
inner peripheral wall and an outer peripheral wall there
are formed parallel flow paths, and between successive
rotor blades there are formed rotor flow passages through
which the flow paths extend.
In prior art compressors of the above type there is a
problem of obtaining an increased pressure ratio across
each compressor stage and/or an increased efficiency. A
factor which is limiting and Crucial for these objectives
is the mean velocity of the air flow through the
compressor. It is a well known fact that an increased air
flow velocity would give a higher pressure ratio across
each Compressor stage and/or an increased efficiency. In
prior art compressors, however, the flow velocity is kept
well below the sonic velocity, i.e. a Mach number below
1,0, usually around 0,7, because at super sonic velocity
there arise shock waves in the air flow which are difficult
to avoid and which are detrimental to the pressure ratio
and the Compressor efficiency. By keeping the Mach number
around 0,7 there is ensured that the Mach number 1,0 will
not be reached and that no shock waves will arise.
The reason for using such a large "safety" margin of 0,7 -
1,0 in Mach number is that in prior art compressors the air
flow velocity normally increases locally at the downstream
ends of the flow paths of the stator or rotor sections. The
reason for such velocity increase is that when departing
from a flow path between two guide vanes or two drive
blades the air flow is subjected to a tangential
SUBSTITUTE SHEET (RULE 26)

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contraction due to a change in flow direction. Such an
increase of the flow velocity might bring the air flow
velocity up to a Mach number around 1,0, and undesired
shock waves might arise in the air flow. In order to make
sure that sonic velocity is not reached at any location in
the compressor, the air flow velocity is kept down to the
"safe" Mach number 0,7.
There are transonic compressors working at velocities
exceeding Mach number 1,0 and by which special arrangements
have been made to avoid the negative influence of shock
waves. However, that type of compressor would also benefit
from lower air flow losses in the stator and rotor flow
paths in accordance with the invention.
The main object of the invention is to accomplish a
compressor of the above type working at subsonic air flow
velocities and where the air flow passages through the
compressor are improved in such a way that the mean air
flow velocity through the stator and rotor sections may be
increased considerably without risking the Mach number
reaching the 1,0 level.
Further objects and advantages of the invention will appear
from the following specification and claims.
A preferred embodiment of the invention is hereinafter
described in detail with reference to the accompanying
drawing.
On the drawing
Fig. 1 shows the geometry of the flow path through a rotor
blade passage.
Fig. 2 shows a side view of the rotor blade passage of Fig.
1.
Fig. 3 shows a fractional longitudinal section through a
compressor according to the invention.

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Fig. 4 shows a spread-out view of the rotor blades and
stator guide vanes of the compressor illustrated in Fig. 1.
In Fig. 1 there is illustrated a flow path relative to the
rotor which extends through a passage between two
successive rotor blades A and B. Before entering the
passage between the rotor blades A,B, the medium flow is
deflected in a direction opposite the movement direction w
of the rotor blades by an angle dal , which is the
difference between the original flow path angle al and the
new flow path angle &1. This deflection of the flow causes
a sidewise contraction of the flow path and follows a
curvature which may be theoretically calculated via a well
recognised method, see for instance: Eckert/Schnell "Axial-
und Radialkompressoren", 2:nd edition, page 264, or
"Dubbel Taschenbuch fur den Maschinenbau", 1974, page 334.
The curvature has a shape which is close to a circle line
with a radius R.
When reaching a section a2 at a certain distance upstream
of the leading edge of the second rotor blade B or slightly
downstream of the leading edge of the first rotor blade A
the flow path passes through a diffusor region C which
extends in the flow direction to a section a3 approximately
at the leading edge of the second rotor blade B.
Accordingly, the diffusor region C has an entrance section
a~ and an exit section a3 , wherein the entrance section a2
has a cross sectional area which is smaller than that of
the exit section a3 . The entrance section a~ of the
diffusor region C is also the narrowest cross section of
the entire flow path between aland a4.
Downstream of the diffusor region C, the flow path extends
through a transition region D which has a substantially
non-increasing or slightlyodecreasing cross sectional area
all the way from section a3 to an exit section a4 . To

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compensate for a downstream increasing distance between
the rotor blades A and B the radial extent of the rotor
blades, i.e. the radial distance between the inner
peripheral wall 28 and the outer peripheral wall 29, has to
be reduced so as to keep the cross sectional area
substantially constant throughout the transition region D.
See Fig. 2. In some cases it might be advantageous to have
a slight acceleration of the flow through the transition
region D.
Upstream of the diffusor region C, the flow path has a
substantially constant cross sectional area, from an
initial section a1 to the diffusor region entrance section
a~ so as to generate~a non-increasing flow velocity. As
illustrated in Fig. 2, this is accomplished by forming the
inner and/or the outer peripheral walls 28,29 of the rotor
and the stator, respectively, with diverging surfaces F and
G. These diverging surfaces F, G compensate for the
sidewise contraction of the flow path, as described above,
and serves to keep down the Mach-number of the flow
velocity and prevent shock waves to occur in the medium
flow.
By locating the diffusor region C,of each flow path
upstream of the flow deflecting transition region D between
two successive rotor blades A, B there is accomplished a
reduction in flow velocity and, hence, a reduction of the
flow losses during the flow path deflection between the
rotor blades A, B. This means an improved efficiency of the
compressor.
In order to ensure a good efficiency of the compressor the
flow velocity shall be equally high over the entire radial
extent of each rotor blade. This is accomplished by
employing a guide vane configuration in the initial
compressor stage such that each guide vane 10 has a
different flow deflection angle at its top end compared to

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its bottom end. See Fig. 4. Thereby, there is obtained
optimum flow directions for generating an equal flow
velocity at all radial locations on each rotor blade in the
initial compressor stage.
In Figs. 3 and 4, there is illustrated a preferred
embodiment of the invention including the flow path
characteristics illustrated in Fig. 1.
In Fig. 3 there is shown a sectional view of an inlet
nozzle for the initial stage of the compressor including
guide vanes 10 rigidly mounted in a housing 11. Downstream
of the nozzle 10 there is a rotor section 12 with a rotor
blade 13 followed by a stator section 14 having a guide
vane 15 secured to the housing 11, and another rotor
section 16 with a rotor blade 17. Rotor flow paths 20
extend between two adjacent rotor blades 13,17, and stator
flow paths 21 are formed between two adjacent guide vanes
15. The flow paths 20, 21 are also defined by an inner
peripheral surface 28 and an outer peripheral surface 29.
Between the stator sections and the rotor sections there
are provided axial gaps which form annular air flow
passages 22, 23 and 24.
The main character of the air flow passage through the
compressor is successively converging from the inlet nozzle
end toward the outlet end. As illustrated in Fig. 3, the
cross sectional area of the air passage decreases step-
..A..
wise. In the air flow paths 20 between the rotor blades 13
as well as the flow paths between the guide vanes 15 the
radial extent of the flow passage decreases, whereas in the
flow passages 22, 23 and 24 located between the stator
sections 14 and the rotor sections 12 the radial extent of
the flow passage increases.

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A characteristic feature of the invention is the provision
of the axial gaps between the stator and rotor sections
forming the flow passages 22, 23 and 24. The reason for
introducing these axially extended and radially diverging
passages 22, 23 and 24 is to accomplish a velocity reducing
diffusor region with the purpose to reduce flow losses and
increase the compressor efficiency.
As illustrated in Fig. 4 an air flow approaching the rotor
flow path 20 between two rotor blades 13,17 has a
converging shape, because depending on a difference in
direction between the incoming air flow and the direction
of the rotor blades 13,17, the air flow has to change
direction. As illustrated in Fig. 4, the direction of the
incoming air flow path forms an angle to a radial plane and
is denoted [31 . This angle is larger than the angle of the
rotor blades 13,17, which is denoted (3'1. Due to this
change in flow direction, the air flow path is subjected to
a tangential contraction, which causes an increased flow
velocity. This is illustrated by the numerals b1 and b2,
where b2 illustrates a narrower flow path cross section
than the incoming flow in section b1 . The acceleration of
the air flow is disadvantageous since it results in an
increased frictional losses in the flow paths.
This undesirable acceleration of the air flow is omitted by
increasing the available cross sectional area in the flow
passage, i.e. by the introduction of the intermediate and
radially diverging flow passages 22, 23 and 24. By
increasing the radial extent of these passages by at least
10% there is obtained an improved compressor efficiency.
For obtaining a substantial increase in the compressor
efficiency the increase in the radial extent of the flow
passages 22, 23 and 24 should be at least 200. In the
illustrated example, the radial extent of the passages
increases from hl at the entrance to h2 at the exit end.

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For obtaining a favourable shape of the air flow path
through the compressor, the increase in radial extent of
the intermediate passages 22, 23 and 24 has to be
accomplished over a certain passage length. Therefore, the
passages 22, 23 and 24 should have an axial length
exceeding 30% of the rotor blade and guide vane length,
respectively. Depending on the radial extent of the blades
and vanes the passage length could be 500 or more of the
length of the blades and vanes, respectively.

Dessin représentatif