U.S. patent number 3,904,308 [Application Number 05/469,500] was granted by the patent office on 1975-09-09 for supersonic centrifugal compressors.
This patent grant is currently assigned to Office Nationale d'Etudes et de Recherches Aerospatiales (O.N.E.R.A.). Invention is credited to Yves Ribaud.
United States Patent |
3,904,308 |
Ribaud |
September 9, 1975 |
Supersonic centrifugal compressors
Abstract
A supersonic centrifugal compressor comprises a rotor located in
a housing aving a fluid intake eye. The fluid (air for instance)
successively travels through an intake region wherein the rotor has
a small number of blades which deflect the fluid tangentially by a
small amount only, then through a compression region wherein the
rotor has a higher number of blades producing tangential and
meridian flow deflection. Last, the fluid flows substantially
radially with respect to the rotor into a stationary diffuser.
Inventors: |
Ribaud; Yves (Palaiseau,
FR) |
Assignee: |
Office Nationale d'Etudes et de
Recherches Aerospatiales (O.N.E.R.A.) (Chatillon-sous-Bagneux,
FR)
|
Family
ID: |
9119415 |
Appl.
No.: |
05/469,500 |
Filed: |
May 13, 1974 |
Foreign Application Priority Data
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|
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May 16, 1973 [FR] |
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73.17730 |
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Current U.S.
Class: |
415/143; 415/181;
415/208.3; 416/183; 416/185; 416/188 |
Current CPC
Class: |
F04D
21/00 (20130101); F04D 29/30 (20130101); F04D
29/284 (20130101) |
Current International
Class: |
F04D
29/28 (20060101); F04D 21/00 (20060101); F04D
021/00 (); F04D 029/44 () |
Field of
Search: |
;415/181,143,213,211,219A,163 ;416/183,188 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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|
|
|
|
|
459,043 |
|
Jul 1953 |
|
IT |
|
1,188,110 |
|
Mar 1959 |
|
FR |
|
830,542 |
|
Feb 1952 |
|
DT |
|
594,537 |
|
Nov 1947 |
|
GB |
|
Primary Examiner: Raduazo; Henry F.
Attorney, Agent or Firm: Larson, Taylor & Hinds
Claims
I claim:
1. A supersonic centrifugal compressor comprising a rotor provided
with at least first blades and second blades and with a stationary
diffuser surrounding the rotor and carried by a housing having at
least one axial fluid inlet, wherein said rotor and housing limit
an intake region, a compression region and a radial flow region
through which the fluid flows successively and wherein said blades
each have a tip portion and a root portion and extend continuously
from a leading edge to a trailing edge, the trailing edges of all
said blades being located in said radial flow region,
said intake region extending from the leading edges of said first
blades to the leading edges of said second blades, said second
blades being shorter than the first blades, said first blades
having a stag variation in the intake region which is such as to
produce a tangential fluid deflection which is small as compared
with the tangential flow deflection in the compression region, and
having a substantial stag at their leading edges,
said first and second blades being so shaped in the compression
region as to produce the major portion of the tangential flow
deflection and azimuthal flow deflection,
and said first and second blades having trailing end portions
directed radially with respect to the rotor in said radial flow
region.
2. A compressor according to claim 1, wherein the deflection along
the upper surfaces of the blades between the intake and the
orthogonal projection of the leading edge of the next blade is
about 7.degree. at the tip of the blade, and the divergence between
the tips of adjacent blades is about 7.degree..
3. A compressor according to claim 1, wherein the diffuser
comprises a vaneless diffuser surrounded by a vaned diffuser and
the ratio of the outer diameter of the vaneless diffuser to the
rotor outlet diameter is between 1.15 and 1.30.
4. A compressor according to claim 1, wherein in the radial flow
region, the rotor has third blades whose leading edges are located
at the outlet of the compression region.
5. A compressor according to claim 4, wherein the thickness of the
first blades increases from their leading edges up to a
cross-section level with the leading edges of the second
blades.
6. A compressor according to claim 4, wherein the rotor has from
six to 12 first blades, from 24 to 32 blades in the compression
region, and at least 36 blades in the radial flow region.
7. A compressor according to claim 1, wherein the absolute outlet
angle of the fluid from the rotor is between 6.degree. and
12.degree. with respect to the tangential direction.
8. A compressor according to claim 7, wherein the diffuser has
vanes and the angular position of the diffuser vanes can be
adjusted so as to alter the cross-section of throats of said
diffuser.
9. A compressor according to claim 1, wherein the hub ratio in the
compressor intake cross-section is of the order of 2 to 2.2.
10. A compressor according to claim 1, wherein the stag angle of
the leading edges of the rotor blades at the tip of the blades is
from 60.degree. to 65.degree..
11. A compressor according to claim 4, wherein the rotor consists
of an upstream wheel bearing a front fraction of the first blades,
and a wheel which carries the balance of the blades.
Description
BACKGROUND AND SUMMARY OF THE INVENTION
The invention relates to supersonic centrifugal compressors and
more particularly to compressors which are adapted to provide a
large flow rate which is large as compared with the front
dimensions of the rotor.
A conventional centrifugal compressor comprises a rotor consisting
of a disc secured to a shaft and provided with blades and mounted
in a housing comprising a diffuser. Fluid enters through an axial
aperture in the housing (or two apertures in the case of a
double-flow compressor) and is accelerated by the blades, after
which its pressure increases owing to slowing down with respect to
the blades and owing to centrifugation. On leaving the rotor, the
fluid has considerable kinetic energy which is recovered in the
form of pressure in the diffuser. There are numerous known kinds of
diffusion systems, the most widely-used being the vaneless diffuser
and the vaned diffuser when the intake flow is supersonic, and the
scroll.
If a centrifugal compressor is to provide a high compression rate,
the rotor peripheral speed must be high, typically about 600 m/s
for compression rate of about 10 if the fluid is air. Under these
conditions, the gas flow leaving the rotor has a supersonic
absolute velocity. In the case, as before, of a centrifugal
compressor providing a highly specific flow rate (the ratio of the
mass flow rate to the frontal cross-section of the rotor) and a
high compression rate (e.g. 10 as hereinbefore mentioned) the
absolute velocity at the blade tip at the rotor intake will usually
also be supersonic. When the ratio between the peripheral radius
and the intake radius of the rotor is 1.5, the relative velocity at
the blade tip at the rotor intake is considerably greater than that
of sound, typically of about MACH 1.3. Since the flow is supersonic
both in the rotor intake regions near the blade tips and in the
diffuser, recompression shock waves will necessarily occur. The
compressor efficiency is closely dependent on the manner in which
the flow is organized in the shock wave regions and on the
stabilities thereof. Furthermore, in most prior-art compressors,
the tangential and axial velocities of the fluid at the rotor
outlet are not uniform, thus reducing the efficiency of a vaneless
diffuser surrounding the rotor so that it is necessary, in many
cases, to provide the rotor with guide means, such as described in
French Pat. No. 7,219,200 of the assignee of the present
application, while such guide means are satisfactory, they have the
disadvantage of making the compressor more complicated.
Other known centrifugal compressors (French Pat. application No.
2,023,770) comprise an upstream wheel forming an axial compressor
followed by a wheel keyed to the same shaft and having blades
separated from those of the outer wheel by an axial gap for flow
stabilization. The latter approach increases the axial length and
weight of the compressor, which is objectionable in aeronautics,
and renders it necessary to machine two sets of blades haaving
different characteristics, thereby rendering machining more
intricate and costly.
It is an object of the invention to provide a supersonic
centrifugal compressor which is improved with respect to prior-art
supersonic compressors. It is a more specific object to provide a
compressor whose efficiency, more particularly that of the rotor,
is improved, and in which a relatively uniform tangential and axial
flow is achieved at the rotor outlet, i.e. in the axial direction
of the rotor.
To this end, in a supersonic centrifugal compressor comprising a
bladed rotor and a stationary diffuser surrounding the rotor and
borne by a housing having at least one axial fluid intake, the
fluid to be compressed successively travels through an intake
region wherein the rotor has a small number of blades producing a
slight tangential deflection in the fluid and a slight divergence
in order that the recompression shock be stabilized near the
leading edge on the compressing surface of the blades and
breakdowns in the flow and a compression region. In the compression
region, the rotor has a higher number of blades simultaneously
producing tangential deflection of the fluid and deflection in the
meridian plane. Last, the fluid flows through a region where the
flow is substantially radial with respect to the rotor.
The tangential deflection between two successive points along a
same flow line is the variation in the stag between the two points.
The stag is the angle formed by the tangent to the flow line at a
given point and the meridian line of the plane which is tangent to
the flow surface at the same point.
The intake region is also designed to stabilize the recompression
shock wave. In that same intake region, the deflection along the
upper surface of the blades and the divergence may advantageously
be of about 7.degree.. Usually, the number of blades in the intake
region is between six and twelve. In the compression region,
between 24 and 32 blades can be provided, the number being higher
if the front size of the rotor is greater. As an example,
intermediate blades, e.g. 3 or 4 intermediate blades per main
blade, can be provided between the blades extending all the way
along the rotor, starting from the intake. In the outlet region,
additional short substantially radial blades are provided and
adapted to prevent breakdowns or stall in the flow on the upper
surface. In practice, at least 32 blades are usually required in
the outlet region. In many cases, it may be sufficient to provide
an additional guide blade in the outlet region in the middle of
each duct bounded by two longer blades.
The compressor according to the invention is suitable for use in a
number of technical fields. More particularly, it may be used in
aeronautics as a turbo-jet intake component. It can also be used in
industry, inter alia when heavy gases have to be compressed and
when it is important to obtain high efficiency and/or a high
specific flow rate, e.g. for compressing heavy gases, e.g. uranium
hexafluoride used in isotopic enrichment plants.
The invention will be more clearly understood from the following
description of a particular embodiment of a supersonic compressor
given by way of example.
SHORT DESCRIPTION OF THE DRAWINGS
FIG. 1 is a diagrammatic half sectional view of a supersonic
compressor, showing the essential components thereof, along an
axial plane;
FIG. 2 is a perspective view of the compressor rotor;
FIG. 3 is a simplified graph showing part of the intake of the
rotor blades and the corresponding velocities diagram; and
FIG. 4 is a simplified view along IV--IV of FIG. 1, showing the
rotor outlet and part of the diffuser.
DESCRIPTION OF A PREFERRED EMBODIMENT
Referring to FIG. 1, there is shown a centrifugal compressor for
use where maximum reduction of the front size of the rotor is not a
primary requirement. Such compressors may be used mainly in
stationary installations. Since the rotor diameter is relatively
large, the total number of vanes can be substantially higher than
in the case of a compressor for use in aeronautics, when the radial
dimensions have to be smaller.
The compressor comprises a single flow rotor 1 carried by a driving
shaft 2 and provided with blades. A vaneless diffuser 5 and a vaned
diffuser 6 are provided around the rotor in a housing 3 which
surrounds the rotor and limits an intake 4. The rotor 1, which is
provided with blades, draws the fluid to be compressed (e.g. a
heavy gas) through intake 4 and forces it into the vaneless
diffuser, whence it travels into the vaned diffuser 6 and finally
into an annular scroll (not shown) surrounding the vaned
diffuser.
During its travel from intake 4 to diffuser 5, it can be considered
that the fluid travels through three successive regions.
The intake region begins at the leading edge of blades 7 (FIGS 1, 2
and 3). As already stated, the flow will be considerably above
sonic speed at the blade tip at point A (e.g. the relative speed of
the fluid with respect to the blades will be MACH 1.3) and will be
only slightly below sonic speed at the root of the blade at point B
(e.g. MACH 0.9) Consequently, under started flow conditions, e.g.
during normal operating conditions, a recompression shock wave 8a
will form and will be connected to an oblique shock wave coming
from the leading edge. The shock wave 8a must be in a region where
the fluid is only very slightly deflected by the blades and where
the flow is only very slightly divergent in order to prevent
breakdown or stall therein.
In practice, in spite of the increased thickness of the blades from
the leading edge onwards, there is a divergence owing to the blade
curvature, and the latter must be moderate (more particularly on
the upper surface of the blade). More particularly, the diffuser
throat cross-section must be selected so that the shock wave 8a is
kept in the immediate neighbourhood of the leading edge at the
blade tip, on the compress-surface side (i.e. near the line AC). It
is known that, if the throat section is decreased, the shock wave
tends to move towards the leading edge. Even if the shock wave is
substantially level with AC, there is no appreciable risk of surge,
since the relative flow speed remains subsonic at the root of the
blade.
In order to obtain satisfactory efficiency, the recompression shock
wave 8a must be in a region in which the fluid is only very
slightly deflected by the blades, since this minimises overspeeds
along the upper surface, i.e. the intensity of shocks, and prevents
considerable breakdown in the flow or stall. In practice, 7.degree.
may be considered as near the optimum value for the angle of
deflection between the leading edge A of the blade and the point C
where the blade intersects the perpendicular from the leading edge
A of the next blade. A divergence of 7.degree. is likewise
acceptable.
To obtain a large mass flow rate, only a small number (about 8) of
blades 7 must be provided in the intake region, so that the total
thickness of the blades does not excessively reduce the most
restricted cross-area left for fluid flow. Advantageously, the
blades are narrow and have a high hub ratio (ratio between the
radii of the blade tip and the blade root) in the intake
cross-section.
That ratio may be from 2 to 2.2. The thickness of the blade, in an
embodiment which will again be referred to later, is e.g. 1.5 mm at
the blade root and 0.75 mm at the blade tip. The stag angle of the
leading edge of the blade tips is e.g. 60.degree. to 65.degree.
with respect to the axis.
Since there is a small number of blades, the limit flow rate which
can be provided by the rotor can be determined relatively
accurately from the triangle of velocities given in FIG. 3. The
rate of flow velocity (or meridian velocity C.sub.m) can be
calculated by constructing the triangle of velocities, since the
rotating speed .omega..sub.r and the direction of the fluid flow
relative to the blades (substantially in the direction of the
blades in the intake region AC) are known.
On leaving the intake region, the fluid travels into the deflection
region, in which the blades azimutally (or tangentially) deflect
the flow to an angle which is usually of about 58.degree. if the
total deflection at the blade tip is of the order of 65.degree. and
if, as already stated, the deflection in the intake region does not
exceed 7.degree..
The following steps may be taken in order to obtain a large
deflection together with satisfactory uniformity of flow at the
rotor outlet, where the blades are substantially radial.
First, the blades in the compression region are shaped so that they
deflect the flow simultaneously in the tangential direction and in
the meridian direction. The center line 9a (FIG. 1) is given a
radius of curvature defined by the following conditions: the radius
of curvature of the central flow meridian (or centroid) line 9a is
such that the pressure gradient is zero along a direction at right
angles to the centroid line. Starting from this line, the blade
skeleton is defined by a straight line which bears on the rotor
axis and the central current line 9a while remaining perpendicular
to the centroid current surface (which is a surface of revolution,
while the meridian plane has an angular position which changes
along the blade).
The corresponding equations can be written as follows: ##EQU1##
In addition to these equations, the following conditions in respect
of the central meridian or centroid line (the line whose distance
at each point from the axis is such that it divides the flow into
two parts having the same cross-section) should be fulfilled:
-- the rate of variation in the meridian velocity, in dependance on
the curvilinear abscissa along the line should be substantially
linear or proportional,
-- the rate of variation of the tangential deflection in dependence
on the curvilinear abscissa s along the meridian line is
s.sup..alpha., with .alpha. between 1.5 and 2.
Equation (1) defines the average flow line, in conjunction with the
above-mentioned conditions. In the equations, F.sub.n denotes the
force component at right angles to the blade, i.e. in the direction
of the line .DELTA. in FIG. 1; the term ##EQU2## corresponds to the
force produced by the curvature of the flow surface; and the third
term corresponds to the centrifugal force, r denoting the radius at
the point in question and .epsilon. denoting the angle between the
axis and the tangent to the meridian flow line. The third equation,
in which C.sub.m denotes the meridian velocity and v denotes the
absolute tangential velocity, defines the radius of curvature R.
The radius of curvature R of the central meridian line is infinite
at the intake at the outlet, .epsilon. being equal to 90.degree. at
the outlet and v (the tangential velocity) being substantially zero
at the intake.
For a same curvilinear abscissa, the deflections of the blade tip
and the blade root as determined by the previously-defined method
of generation are different from the deflection along the central
meridian line. The coefficient .alpha. should be made low enough to
ensure that, starting from constructive data (e.g. the hub ratio
and the intake angle), a tip deflection is computed which does not
detrimentally affect the mechanical stresses at the blade root to
an excessive extent. If the peripheral velocity (at the blade tip)
is too high, the skeleton may have to be generated from straight
lines which are not perpendicular to the central meridian line any
longer.
Another step is to increase the number of blades, in order to
ensure satisfactory efficiency and uniform velocities in the
compression region. The blades 7 extending all along the rotor are
supplemented with blades 8 such that the total number of blades in
the compression region is between 24 and 32. To this end, in the
example shown, three additional blades 8 are inserted between each
two blades 7. The baldes 8 have a shape which is identical with
that of that part of blades 7 in the same annular portion. The
thickness of the leading edge will be greater than that along A-B,
since the speed is sub-sonic. In the example considered
hereinbefore, the thickness and diameter of the leading edge may be
1.5 mm at the tip and 3 mm at the root of blade 8.
In order to simplify manufacture, the rotor can be made in two
components. A first component, comprising the upstream wheel, bears
those portions of the blades 7 which are upstream of the leading
edge of blades 8. A rear part, forming the wheel proper, bears the
rest of the blades. The rotor can be machined after the wheel and
the outer wheel have been assembled.
Last, after substantially complete tangential deflection, the fluid
flows through a region of substantially radial flow which comprises
that part of the blades 7 and 8 which extends until the rotor
outlet and supplemental short blades 10 adapted to guide the flow
and prevent breakdown from occuring therein on the upper surface of
the blades.
The inner and outer walls of the rotor illustrated by way of
example in FIG. 1 are designed so that the meeridian velocity in
each cross-section taken at right angles to the flow surface
corresponding to 9a increases in direct proportion of the
curvilinear abscissa along the central meridian line; this result
is achieved by providing the radial flow region with convergent
surfaces 11 and 12. If, for example, a convergent angle of
12.degree. is selected, the surfaces of the radial portion will be
inclined at 6.degree. to the radial direction. Of course, an
asymmetrical arrangement is possible or even necessary in a double
flow compressor.
The flow speed at the rotor outlet is uniform in both the axial and
the tangential direction. Since the flow is axially uniform, a
vaneless diffuser 5 may be used which is known to have a low
efficiency if there is marked non-uniformity. A vaneless diffuser
has the advantage of reducing the absolute supersonic velocity of
the flow at the rotor outlet, without producing a shock.
An industrial compressor according to the invention can include a
large vaneless diffuser 5; the ratio between its outer and inner
diameter can be up to 1.30. Space requirements frequently render a
lower ratio -- at most 1.15-- advisable. In an aeronautical
compressor, having a compression rate of 5 to 6 and for M .about.
1.5 at the intake, the ratio is frequently limited to 1.05 in order
to reduce bulk.
A vaned diffuser 6 is disposed downstream of the vaneless diffuser
5 and has an axial depth equal to that of the wheel (i.e.
corresponding to the axial length of the trailing edge of the
blades). Advantageously, the number of ducts is large, so that the
bulk can be reduced. The leading angle can be low, e.g. between
6.degree. and 12 .degree., with respect to the tangential
direction, but increasing the diffusion length for a given radial
size. The diameter of the leading edge of each blade of diffuser 13
is typically about 5 to 10% of the throat width. The inlet of each
actual duct forms a throat having a length which is substantially
equal to half its depth, and adapted to stabilize the recompression
shock waves. Downstream, the blades define a duct having an angle
of divergence of approximately 6.degree. to 7.degree.. The
uncovered part of the upper surface of each blade, i.e. the intake
region from the leading edge which does not bound a duct, is
preferably in the form of a spiral. The ratio between the outlet
cross-section of the diffuser and the intake cross-section
perpendicular to the flow may be e.g. between 3 and 3.5 for a
divergence of 6.degree..
The blades can be angularly adjustable around a shaft 14 so that
the throat width can be adjusted by moving the blades from the
broken-line position to the continuous-line position in FIG. 4, in
order to determine the position for optimum efficiency.
The vaneless diffuser 5 having parallel surfaces could be replaced
by a convergent diffuser extending the rotor ducts, as illustrated
in chain-dotted lines in FIG. 1.
In the example given, the diffuser had 23 channels, and this has
given satisfactory results. A centrifugal compressor of that type
can be constructed which provides a per stage efficiency greater
than 80% for a compression rate between 6 and 8, if the fluid is
air.
* * * * *