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Our invention relates to improvements in the diffusers used with gas
turbine engines, particularly high performance engines as are employed in
the propulsion of aircraft. In such engines, the diffuser and associated
compressor are essential components for pressurizing air as a preliminary
step in the generation of a high energy, hot gas stream.
Rotary type compressors are often used for this purpose and comprise an
impeller, or rotor, which imparts energy to the air, primarily in the form
of increased velocity. The high velocites of the air discharged from the
exit side of the impeller are too great for practical utilization in
supporting combustion of fuel. Therefore, it is accepted practice to
provide a diffuser immediately downstream of the impeller. The diffuser
decelerates the discharge air to relatively low velocities and converts a
major portion of the velocity energy to static pressure energy. In most
compressors, the impeller, or rotor, has projecting blades over which the
air flows in discrete paths as it is accelerated thereby. Likewise, the
diffuser, or stator, has vanes which split the high velocity discharge air
into discrete flow paths.
A major problem in the operation of compressors is the phenomenon known as
surge. When this condition occurs, flow of air through the compressor is
throttled, either locally or completely, and in some cases reverse air
flow can occur. The result of compressor surge is a reduction in power in
all cases and frequently a flameout of the combustor, in which case there
is a complete loss of power.
Surge will occur, at a given engine speed, when the aerodynamic loading on
the blades or vanes exceed a given limit, causing separation of the air
from the flow passageway surfaces and a condition of high turbulence. This
limit varies between different compressor designs and is established for
each compressor design by way of what is known as a compressor map.
Knowing the characteristics of a given design, it is then possible to
control the operation of the engine, primarily through the rate of fuel
flow to the combustor, so that there is a margin of safety in both steady
state and transient operation.
Several different approaches have been used to solve the surge problem.
Conrad in German Pat. No. 1,938,132 and British Pat. No. 1,043,168 show
implementations wherein pressure is bled from a higher of a lower level to
prevent build up of shockwaves in the diffuser passageway throats. The
pressure bleed off is achieved by means of connecting pipes which either
recirculate the fluid to a lower pressure point in the system or vent it.
O'Connor in U.S. Pat. No. 3,768,919, shows a pipe diffuser with an
aerodynamically variable throat area. A series of ports are provided in
the throat region of the diffuser passages to momentarily inject
pressurized exit air to aerodynamically vary the throat flow
characteristics and prevent surge during operation of the stage above its
normal surge line.
Sobey in U.S. Pat. No. 3,006,145 shows an antisurge control system which
makes use of a compressor bleed system. He uses a bleed valve which is
responsive to both compressor rotor speed and acceleration of the
compressor rotor.
Our invention differs from the above in that we provide slots in the
sidewalls of the throat section of each vane of the diffuser. These slots
communicate through cavities in each vane with a closed manifold. The
benefits achieved by the use of a closed manifold have been verified by
means of test instrumentation. Data taken from operating diffusers show
that shock waves tend to build up in the throat areas of some passageways
before they do in others. This may be due to imperfections in the vanes or
can be caused by the shadow effects of strut vanes in the compressor
stages. Use of a closed manifold in communication with slots in the
passageway wall alleviated the problem in that tendencies for pressure
surges in one or more passageways was quickly equalized across all passges
through flow into and out of the connecting manifold. This phenomenon was
never mentioned in any lof the cited patents.
SUMMARY OF THE INVENTION
While relating to compressor assemblies generally, this invention will be
described as it rlates to a compressor stage having a bladed radial flow
impeller and an annular radial flow diffuser having its inner periphery
closely surrounding the discharge end of the impeller. The inlet of the
diffuser includes a vaneless entrance space for receiving fluid discharged
from the impeller.
The entrance space is formed by spaced apart walls which are coextensive
with the impeller shroud. Between the spaced apart walls of the diffuser
are a multiplicity of wedge-shaped vanes. These vanes are symmetrically
disposed, adjacent vanes forming therebetween a plurality of intersecting
passageways which extend outwardly from the annular entrance space in a
direction that is tangential with the inner periphery of the diffuser.
Each passageway has a convergent entrance portion immediately adjacent the
vaneless entrance. This is followed by a throat section of constant cross
section. Downstream of the throat section, each passage opens into an area
of expanding cross section wherein fluid velocity is exchanged for an
increase in pressure. The divergent section of each passageway terminates
in an exhaust manifold.
Our invention pertains to the incorporation of flow equalization for
preventing surge and stablizing fluid flow through the diffuser
passageways. Flow energization was achieved by forming slots in the inward
facing wall of each wedge-shaped vane. Each slot communicated with a
cavity inside each vane. Openings made through one of the spaced apart
walls of the diffuser allowed the multiplicity of cavities to communicate
with a closed common manifold. Several slot locations and configurations
were tried as will be described later. However, the preferred approach
involved forming transverse slots in the throat section of each
passageway.
Inclusion of a common manifold in communication with slots in the low
pressure side of each vane allowed fluid to flow into and out of the
manifold via the multiplicity of cavities within the vanes, thereby
serving to equalize the pressure in all of the passageway throat sections.
This greatly improved surge margin performance.
Shaping of the slots can affect performance. Several configurations were
tried and embodiments which function best are delineated. It is the
concept of pressure equalization by means of a closed common manifold in
combination with cavities and slots which communicate with each of the
diffuser passaageways that is the heart of our invention.
Previous proposals for so increasing the surge or operating range of a
given compressor design have either involved undue performance penalties
in terms of efficiency or have been of limited effectiveness, or both.
Accordingly, the primary object of the present invention is to increase
the surge range of rotary compressors for pressurizing compressible
fluids.
Another object of the present invention is to increase such surge range
with a minimum adverse effect on compressor or engine cycle efficiency, if
not, in fact, obtaining an increase in such efficiency.
Another object of the present invention is to minimize the occurence of
surge in both the rotating and stationary components of compressors,
whether the radial flow or axial flow type.
In the broader aspects of the invention, these ends are attained by a
compressor comprising a rotor component and a relatively stationary
diffuser component, which together form a compressor stage. At least one
of these components comprises a plurality of flow passageways divided by
spaced vanes. The sidewalls of the diffuser vanes produce passageways
which together form a throat section downstream of the leading edges of
the vanes. Slots are provided in the throat section walls. The slot
openings connect with cavities in each vane. The cavities are then
interconnected with a closed manifold encircling the outside wall of the
diffuser. By thus interconnecting the flow passageways, surge causing
conditions are equalized between the several flow passageways. Where flow
conditions might have caused surge in a given passageway which could build
up and propagate to all passageways, the manifold interconnection relieves
such conditions to the end that individual passageways are not
aerodynamically overloaded and the surge range and operating range are
appreciably increased. The increased surge range enable operation at
higher pressure ratios with a resultant increase in compressor and engine
cycle efficiency, while the increase in operating range gives a greater
margin of safety in engine operation.
The slots are preferably disposed along a line of equal pressure within
each passageway throat section. The slots may be advantageously located on
the vane suction surfaces. Slots may also be employed on more than one
surface of the flow passageways. In axial flow compressor rotors, the slot
means are preferably located at the tip end portions of the vanes which
define the flow passageways thereof. The slots are also effective in
so-called pipe diffusers.
The interconnecting manifold may also be bled to a lower pressure during
critical portions of engine operation, such as acceleration, to
temporarily provide an even greater increase in the operating range of the
compressor.
The above and other related objects and features of the invention will be
apparent from a reading of the following description of the disclosure,
with reference to the accompanying drawings, and the novelty thereof
pointed out in the appended claims.
BRIEF DESCRIPTION OF THE DRAWINGS
In the drawings:
FIG. 1 is a simplified, longitudinal, half section of a gas turbine engine
of the type in which the improved compressor of the present invention may
be advantageously incorporated;
FIG. 2 is a view, on an enlarged scale, taken generally on line 2--2 in
FIG. 1;
FIG. 3 is a view, on a further enlarged scale, of a portion of the diffuser
seen in FIG. 2, more particularly illustrating the invention;
FIG. 4 is a section taken on line 4--4 in FIg. 3;
FIG. 5 is a section taken generally on line 5--5 in FIG. 3;
FIG. 6 is a view similar to that of FIG. 3, illustrating another embodiment
of the invention;
FIG. 7 is a section taken on line 7--7 in FIG. 6; FIG. 8 is a view similar
to that of FIG. 3, illustrating another embodiment of the invention;
FIG. 9 is a section taken on line 9--9 in FIG. 8;
FIG. 10 is a view similar to that of FIG. 3, illustrating another
embodiment of the invention;
FIG. 11 is a section taken on line 11--11 in FIG. 10;
FIG. 12 is a section similar to that of FIG. 3, and taken on line 12--12 in
FIG. 13, illustrating the invention in a different type of radial flow
diffuser, known as a pipe diffuser;
FIG. 13 is a section taken on line 13--13 in FIG. 12;
FIG. 14 is a section taken generally on line 14--14 in FIG. 12;
FIG. 15 is a schematic view of the invention incorporated into an engine
control system;
FIG. 16 is a longitudinal section of a portion of an axial flow compressor
in which the present invention is embodied;
FIG. 17 illustrates a flow cascade of rotor blades seen in FIG. 16;
FIG. 18 is a section, on an enlarged scale, taken on line 18--18 in FIG.
16; and
FIG. 19 is a plot of compressor operating parameters known as a compressor
map.
DESCRIPTION OF THE PREFERRED EMBODIMENT
Reference will first be made to FIG. 1 for a description of a gas turbine
engine of the type in which the present invention finds particular
utility. Such engines are well known to those skilled in the art and FIG.
1 is therefore greatly simplified, omitting structural details.
The gas turbine engine, indicated generally by reference character 10,
comprises, as basic units, a radial flow compressor 12, a combustor 14,
and a turbine 16, which are sometimes collectively referred to as a gas
generator.
Air is induced into the compressor 12 through an inlet 18 which turns it
into an axial direction for entrance into the compressor 12. The latter
comprises an impeller 20 having a hub 22 and blades 24. The hub 22 and a
surrounding shroud 26 define an annular flow path which curves from the
axially facing entrance to a circumferential, radial exit, with the flow
path being progressively reduced in area towards the radial exit. As the
impeller rotates, the blades 24, which are in close clearance relationship
with the shroud 26, propel the air at increasing velocities and discharge
it circumferentially of the radial exit at substantially increased total
pressures.
The impeller discharge air then enters a radial flow diffuser 28 from which
it is turned to an axial direction and enters an axial diffuser, or guide
vanes, 30 which properly direct the pressurized air to the combustor 14,
which is of the reverse flow type. The pressurized air flows into an
annular combustion chamber 32 where it supports combustion of fuel
discharged from fuel nozzles 34, in the generation of a high energy, hot
gas stream. This hot gas stream is then turned inwardly through an angle
of approximately 180.degree. to the nozzle diaphragm 36 of the turbine 16.
The hot gas stream is then directed through a bladed turbine rotor 38
which is directly coupled to the compressor impeller 20. The turbine
extracts a portion of the energy of the hot gas stream in thus driving the
compressor impeller of the gas generator.
The majority of the remaining energy of the hot gas stream is then
converted to a useful output, as by being discharged through a propulsion
nozzle, or, as herein illustrated, by driving a power turbine 40. The
latter comprises a nozzle diaphragm 42, mounted on a frame member 44,
which directs the hot gas stream through a bladed turbine rotor 46. The
power turbine rotor 46 is mounted on a forwardly extending shaft 48,
which, generally speaking, has a rate of rotation too great to be directly
coupled to a driven unit. Therefore it is usual practice to provide a gear
box 50 on the front end of the engine 10. The input to this gear box, from
shaft 48, is reduced in speed to a much lower rotational rate and motive
power then derived from an output shaft (not shown) of the gear box.
Reference will also be had to FIG. 2 for a more detailed description of the
compressor 12. The impeller blades 24 are preferably formed integrally
with the hub 22 and have their trailing edges at the periphery of the hub.
The impeller discharge exit thus extends circumferentially of the impeller
with a width, in an axial direction, from the hub side to the shroud side
of the blades 24, generally between parallel planes normal to the impeller
axis.
The exit velocities at the impeller discharge are very high and in advanced
compressor designs are usually supersonic. The diffuser 28 is therefore
provided to efficiently convert a major portion of the velocity energy of
the impeller discharge air to static pressure energy as the flow rate is
reduced to a much lower velocity, compactible with the operational
characteristics of the combustor 14. The diffuser 28 may be of
conventional design in having a plurality of tangentially extending flow
passageways, or channels, 52 which are defined by wedge shaped vanes 54
disposed between a front wall 56 and a rear wall 58 (see also FIGS. 4 and
5). The front diffuser wall 56 may be formed as an extension of the shroud
26 and is generally aligned with the shroud sides of the trailing edges of
the impeller blades 24. The rear diffuser wall 58 may be formed by a frame
member 60 and is generally aligned with the hub sides of the blades 24.
The circumferential, air discharge from the impeller 20 is split into
discrete flow paths by the leading edges 62 of the vanes 54 to enter the
channels 52, which are of rectangular cross section. Each channel 52 has a
slighly convergent entrance portion leading to a throat section the (FIGS.
3 and 4) downstream of which the cross sectional area increases in a
controlled fashion to obtain a maximum reduction of velocity and recovery
of static pressure in a minimum of flow path length.
The vanes 54, or at least the upstream portion thereof, function as
airfoils having suction surfaces 64 and pressure surfaces 66. Nominally
there is a zero degrees incidence angle of the air impinging on the
suction surfaces 64. Variations in static pressure gradient (related to
flow velocity) and incidence angle beyond certain limits will result in
flow separation of the air and cause an increase in the thickness of the
boundary layer of air along the suction surfaces. Beyond certain limits
such increases tend to reduce the mass flow rate of the air for a given
engine speed, until a turbulent separation of the air from the channel
surfaces, particularly the suction surfaces, occurs. This can then result
in a surge condition. The net effect of surge is to throttle or block air
flow and in the some cases, due to the dynamics of the compressible fluid,
i.e. air, there will be reverse flow through the compressor. Surge is
usually initiated in one or a few flow channels and then, due to the
resultant pressure and flow perturbations, propagates to adjacent channels
until surge exits in the entire compressor. While isolated pockets of flow
separation, or stall can exist for a period of time, it is usual for a
surge condition to propagate rapidly, if not instantaneously, causing a
flameout in the combustor and complete loss of engine power. This result
in the propulsion of an aircraft can be quite serious, or even
catastrophic.
The basic flow parameters of velocity and incidence angle are
proportionate, at any engine operating speed, to the pressure ratio across
the compressor and the mass flow of air therethrough. These latter
parameters can be measured directly or indirectly to control engine
operation, usually by means of the rate of fuel flow to the combustor, so
as to avoid conditions which will initiate surge. These relationships,
which vary between different compressor designs, are commonly represented
by what is known as a compressor map, a typical compressor map being shown
in FIG. 19. This map depicts the relationship between the referred weight
flow, or mass flow rate, and the pressure ratio across the compressor at
three engine speeds (N) of 50%, 80% and 100%, by the thin lines on the
map. It will be noted that mass flow remains constant, at a given engine
speed, as the pressure ratio increases through a choke flow range c and
then decreases until surge occurs at the point indicated on the thin surge
line on the map. The surge line is a plot of an infinite number of engine
speeds at which surge occurs.
In order to avoid conditions which would result in surge, normal engine
operation is maintained at a pressure ratio approximately at the upper end
of the choke flow range at any given speed. A plot of an infinite number
of such operating points produces the thin broken operating line for a
typical conventional compressor. The margin between the surge line and the
operating line for steady state operation protects against abnormal
conditions which might affect air flow or pressure ratio and also provides
for safe and rapid surge free engine acceleration.
A measure of compressor performance is its operating range, a preferred
definition of which is
##EQU1##
By increasing the operating range of the compressor, increased performance
is available without the danger of surge.
The means now to be described increase the operating range and raise the
surge range of the typical compressor whose performance has been reflected
by the thin lines in FIG. 19.
Referencing again, the preferred embodiment system shown in FIGS. 2-5, a
slot 68 extends along the height of each channel suction surface 64 at the
throat section th. The slot 68 extends into the vane 54 to a cavity 70
which opens into a passageway 72 formed in the overlying front wall 56.
The passageways 72, in turn, open into a manifold 74 which is mounted on
the front wall 56. All of the slots 68 are thus placed in fluid
communication with each other by way of the cavities 70 and the common
manifold 74.
The effect of these interconnected slots on compressor performance is
illustrated in FIG. 19 by the thick speed lines (N) showing that higher
pressure ratios are attained before surge occurs at the thick surge line
on this compressor map. With the surge range thus increased the operating
line of the compressor can also be raised, as indicated by the thick
broken line on the map, enabling normal operation at higher pressure
ratios. Compared with a base compressor configuration, the performance of
which is indicted by the thin lines in FIG. 19, the described slotted
configuration, the performance of which is indicated by the thick lines in
FIG. 19, increases the operating range at all speeds and, at least at
speeds of N=80% to N=100% provides increased peak efficiences, as well as
increased pressure ratios on both the operating line and the surge line.
For example at N=80% the operating range is 20.5% compared to a base of
11.0% and at N=100% the operating range is 10.7% compared to a base of
7.5%.
The underlying reasons for the improved results obtained are believed to be
twofold. It is a known fact that surge generally initiates in one or a few
channels, or flow passageways, due to manufacturing tolerance variations
between the several channels, or because of transient variations in air
flow or because of conditions affecting flow which are unique to one or a
few channels. These factors cause the vanes of such channels to be
aerodynamically overloaded and surge results. Initial overloading is first
relieved by the plenum effect of the cavities 70 in the transient
initiation of surge. The fluid communication provided by the manifold 74
then provides a steady state equalization of pressures to the end that a
critical channel or channels continue to have favorable vane loadings up
to the point where essentially the entire stage becomes overloaded and
surge occurs simultaneously in all channels, but at a higher pressure
ratio than would have otherwise been obtainable. It would be added that
the manifold itself, in certain configurations, could provide the plenum
effect for transient pressure perturbations.
Another embodiment of the invention is illustrated in FIGS. 6 and 7. The
compressor components are the same as in the previous embodiment (and are
identified by the same reference characters) except that a slot 80 is
provided in the suction surface 64 upstream of the throat th,
approximately half way towards the vane leading edge 62. It will also be
seen that the slot extends only along about one half of the height of the
suction surface 64. The slot 80 opens into an elongated cavity 82 which
extends into registration with the front wall passageway 72. All of the
vanes 54 are provided with slots 80 and cavities 82 thus placing the
several channels 52 in fluid communication with each other through the
plenum 74.
This embodiment of the invention illustrates, at least for radial flow
diffusers, the approximate minimum length of slots that are effective for
the purposes of the present invention. It also illustrates that the slots
can be effectively disposed upstream of the throat section of the flow
channel.
Another embodiment of the invention is illustrated in FIGS. 8 and 9. Again
like reference characters identify the basic component of the compressor
which are unchanged except as regards the slot means and manifold. In this
embodiment the slot means comprise an elongated slot 90 in the rear wall
58 of each flow channel 52. Each slot 90 is disposed upstream of the
throat section th and is angled relative thereto to lie on a line of
approximately equal pressure of the air flowing into the channel. The
slots 90 open directly into an annular manifold 92 formed in the frame
member 60. This provides for fluid communication between the several
channels, as well as providing the plenum effect which was provided by the
cavities 70 and 82 in the previous embodiment. The separate manifold 74
has been eliminated by the internal manifold 92.
This embodiment illustrates that the slot means may be effectively disposed
on other than the suction surfaces of the vanes. It also illustrates that
the slots would lie on lines of essentially equal pressure in the air flow
path. This was, in fact, the case in the previous embodiments where the
slots disposed on the suction surfaces were parallel to the channel throat
sections.
Another embodiment of the invention is illustrated in FIGS. 10 and 11.
Again the basic components of the compressor are unchanged, except for the
slot means and are identified by like reference characters. In this
embodiment there are two slots in each flow channel 52. A slot 100 extends
across the full height of the suction surface 64, as in the first
embodiment. In addition a slot 102 extends across the major portion of the
rear diffuser wall 58, also at the throat section th. The slots 100 and
102, respectively, open into interconnecting cavities 104, 106. The cavity
104 is registered with the passageway 72 formed in the front wall 56, thus
providing a fluid interconnection between the several channels, 52,
through the manifold 74, as before.
This embodiment of the invention illustrates that slot means may be
effectively provided in more than one wall of the flow passageways of the
diffuser to work in combination.
Another embodiment of the invention is shown in FIGS. 12-14. A diffuser
28', commonly known as a pipe diffuser, surrounds an impeller which may be
the same as the impeller 20 previously described. The diffuser 28'
comprises a plurality of flow channels 110 formed in a frame member 112
and extending tangentially of the impeller 20. The channels 110 are
circular in cross section and have cylindrical inlet portions which extend
from a curved groove 114, surrounding the impeller 20, to a throat section
th. Downstream of the throat section th, each channel 110 is divergently
conical to provide the diffusion function. The intersection of the
circular channels 110 with the curved groove 114 results in a swept effect
on the leading edges 116 of the vane portions 118, of the frame member
112, which separate the channels 110. These vane portions likewise
function as airfoils in splitting the air flow into the discrete flow
passageways of the channels 110. Each vane portion 118 has a suction
surface portion 120 and a pressure surface portion 122 leading to the
throat sections of adjacent channels 110. This configuration of diffuser
has been found particularly effective in minimizing losses where the
impeller discharge air is at supersonic velocities.
A slot 124 is formed at the throat section th of each channel 110. The
slots 124 extend around approximately one half of the peripheries of the
respective channels 110, being centered on the suction surface portions
120. Each slot 124 enters a cavity 126. A plate 128 overlies the frame
member 112 and has openings 130 which are registered with the several
cavities 126. A manifold 132 is mounted on the plate 128 and is registered
with the openings 130 to again provide a fluid interconnection between the
several flow channels 110 of the diffuser.
This embodiment illustrates the use of the slot means of the present
invention in a pipe diffuser.
Another embodiment of the invention is illustrated in FIGS. 16-18. FIG. 16
shows, in simplified fashion, a portion of an axial flow compressor 140.
This type of compressor is well known to those skilled in the art and is
employed for the same basic function in gas turbine engines as the radial
flow compressors which have been previously described. The configuration
of gas turbine engines incorporating axial flow compressors is also well
known to those skilled in the art.
The compressor 140 comprises a rotor 142 having a circumferential row of
blades 144 projecting generally radially relative to its axis of rotation.
Immediately downstream of the blades 144 is a circumferential row of
stator vanes 146 which function as a diffuser and together with the blades
144 form a compressor stage. A second row of blades 148, mounted on the
rotor 142, and stator vanes 150 form a second compressor stage, it being
usual that axial flow compressors comprise several stages. The air flow
path through the compressor 140 is annular and generally concentric of the
axis of rotation of the rotor 142. This flow path is defined, at its outer
bounds, by a composite casing 152, with its inner bounds being defined by
an inlet conical member 154, platforms 156 at the bases of the blades 144,
a liner 158 at the inner ends of the vanes 146 and the platforms and inner
liners of subsequent stages. The blades 144 and vanes 146 function as
airfoils, defining separated flow passageways and are aerodynamically
equivalent to the impeller blades 24 and diffuser vanes 54 an imparting
velocity energy to the air and then recovering static pressure energy.
FIG. 17 illustrates a flow cascade of the blades 144 indicating that each
flow passageway is defined by blade suction surfaces 160 and pressure
surfaces 162 leading to a throat section th. Surge problems are equivalent
in that when the blades 144 are overloaded, separation of the air flow
occurs and surge results.
Again interconnected slot means are provided. Each blade 144 has a slot 164
extending along its suction surface 160. Since surge is usually a problem
where peripheral speeds are greatest, the slots 164 are provided in the
tip end portions of the blades 144, where they extend in a generally
radial direction along the throat section th. The slots 164 open into
cavities 166 which extend inwardly to passageways 168 which connect with a
manifold chamber 170. The manifold chamber 170 is defined by a seal member
172 overlies a groove formed annularly in the rotor 142. The seal member
cooperates with sealing grooves on the liner 158 to provide a fluid seal
between the first and second compressor stages.
The manifold chamber 170 provides a fluid interconnection between the
several flow passageways defined by the blades 144 and will likewise
relieve overloaded blade surfaces to deter initiation of surge.
Slot means may also be provided in the diffuser vanes 146 as is indicated
in FIG. 16. Again these slots are provided in the regions of highest
velocity. Slots 176 extend along the suction surfaces of the vanes 146 at
their outer end portions and at their throat sections of their location of
highest loading if it is not at the throat section. The slots 176 open
into cavities 174 formed in the vanes 146 and extending through an outer
liner 178 which is a part of the composite casing 152. The outer liner 178
has an annular groove 180 which defines a manifold 182. This again places
all of the slots 176 in fluid communication with each other.
This embodiment illustrates that the invention is applicable to axial flow
compressors, as well as radial flow compressors and also that it may be
employed on the rotating, accelerating component of the compressor where
surge may also be a problem, particularly in axial flow machines.
The benefits of employing interconnected slot means were described in
detail in connection with the first embodiment of FIGS. 2-5, with
reference to FIG. 19. The other embodiments of the invention also provide
such benefits in raising the operating line and surge line to permit safe
operation at increased pressure ratios with an increased operating range.
Another benefit of the interconnected slot means is that unexpectedly large
increases in the operating range have been obtained by bleeding the
interconnecting manifold during acceleration. FIG. 15 schematically
illustrates a system for attaining these added benefits. The manifold 74,
of the compressor 12, is connected to a valve 190 by a conduit 192. The
valve 190 may be mechanically controlled through a connection 194 to a
function generator 196. The latter may have a mechanical input 198 from a
throttle lever 200 which is normally provided and controls flow of fuel in
the operation of the engine 10. When the throttle lever 200 is displaced,
the mechanical connections 198, 194 open the valve 190 to bleed air from
the manifold 74. Upon completion of the acceleration mode, or after the
rate of acceleration is reduced below a given level, the function
generator 196, acting through the mechanical connection 194, causes the
valve 190 to close, returning the interconnected slot means to the mode of
operation previously described.
The result is to provide an improvement on the showing of Conrad (German
Pat. No. 1,938,132) and O'Connor (U.S. Pat. No. 3,768,919) in that for
most operating conditions the closed manifold alone will prevent
initiation of surge. However, during emergency acceleration of the engine,
pressure surges can be bled off and stable operation achieved. By limiting
bleed to the relatively short duration required for acceleration, there is
a minimal effect on overall compressor efficiency. Of greater importance
is the fact that relatively small amounts of bleed flow produce very
significant increases in the operating range and thus provide a greatly
increased margin of safety at a time when surge is most likely to occur.
In the preceding description reference has been made to specific forms of
compressors employed in gas turbine engines for pressurizing air in the
generation of a high energy, hot gas stream. The broader aspects of the
invention are not so limited, but are applicable to any form of rotary
compressor for compressible fluids wherein the flow therethrough is
divided by vanes or blades, herein generically denominated airfoils, from
which the fluid flow may separate in a surge condition.
The spirit and scope of the present inventive concepts is, therefore, to be
derived solely from the following claims.
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