CA2048829C - High pressure compressor flowpath bleed valve extraction slot - Google Patents
High pressure compressor flowpath bleed valve extraction slot Download PDFInfo
- Publication number
- CA2048829C CA2048829C CA002048829A CA2048829A CA2048829C CA 2048829 C CA2048829 C CA 2048829C CA 002048829 A CA002048829 A CA 002048829A CA 2048829 A CA2048829 A CA 2048829A CA 2048829 C CA2048829 C CA 2048829C
- Authority
- CA
- Canada
- Prior art keywords
- compressor
- air
- slot
- pressure
- bleed
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Fee Related
Links
- 238000000605 extraction Methods 0.000 title abstract description 9
- 230000003068 static effect Effects 0.000 claims description 3
- 238000006243 chemical reaction Methods 0.000 abstract description 2
- 230000000740 bleeding effect Effects 0.000 description 6
- 238000001816 cooling Methods 0.000 description 5
- 238000002485 combustion reaction Methods 0.000 description 2
- 238000009792 diffusion process Methods 0.000 description 2
- 239000000446 fuel Substances 0.000 description 2
- 230000006978 adaptation Effects 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 230000002427 irreversible effect Effects 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000001141 propulsive effect Effects 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
- 238000011144 upstream manufacturing Methods 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/02—Surge control
- F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
- F04D27/023—Details or means for fluid extraction
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
A high pressure, compressor flowpath bleed extraction slot is provided wherein an articulated portion of the outer band of the compressor forms a slot for providing efficient conversion of core air to diffuser air while retaining substantial pressure and velocity energies. The diffuser slot can be applied to an earlier compressor stage with improved performance.
Description
U~
HTGH PRESSURE OOMPRESSOIt FLOWPATki BLEED VALVE EXTRAG"fION SIFT
This invention relates to simplified bleed extraction slots for, gas turbine engines and, more particularly, to a specially configured bleed extraction slot for efficiently converting core air to bleed air with a minimum loss in bleed air velocity and pressure.
BACItGROITPTD OF THE INVENTION
It is often desirable that an aircraft gas turbine engine include within its compressor, a structure which permits bleeding or diversion of high pressure air from a stage, such as the 5th stage of the compressor to provide high pressure air for cooling purposes and for operation of airframe accessories, engine accessories, or engine or aircraft de-icing systems. In other cases, i~t is desirable to include a struoture which pezmits the bleeding of even higher pressure air from the discharge of the compressor to provide pressurized air far cooling downstream turbine components. Both interstage bleed and the compressor discharge bleeding are accomplished by flowpath mechanisms which interfere with the normal airflow patterns in the compressor. Further, the casing or bleed structure adds complexity to the assembly of such an engine.
sy 9~. ! t r~ s G.~ Z;~ r~ ~ (.3 13DV°10107 °2°
The axial location or stage at which air is bled from the compressor is determined by the pressure required to drive the specific system intended to be serviced by that air. In most instances, it is desirable to achieve the highest possible source pressure to also ensure a high delivery pressure. For this reason, prior systems have extracted air from the latter stages of the compressor and more particularly, engines having these systems have been designed to extract high pressure air from the 5th stage of the compressor for low pressure turbine cooling and turbine thermal clearance control. However, bleeding air from the earliest possible stage of the compressor generally increases compressor efficiency by reducing the amount of work invested in the extracted air.
Therefore, it is desirable to achieve the highest possible system supply pressure from the earliest and lowest pressure stage of the compressor. The resulting temperature of the cooling air is also lower and hence more effective.
Known examples of bleed openings or ports can be found in U.S. Patent 4,711,084 to Brockett for an ejector°assisted compressor bleed which discloses a bleed aperture 17 in Fig. 2 having rounded hole edges. U.S. Patent 3,108,767 to Eltis, et al., for a bypass gas turbine engine with an air bleed means in Fig. 3 discloses a duct 19 which is attached to the compressor through a series of chopped holes. U.S.
Patent 3,898,799 to Pollert, et al., for a device for bleeding off compressor air in a turbine jet engine, in Fig. 2 discloses a compressor orifice marked with the arrow K. U.S. Patent 3,777,489 to 0'ohnston, et al., discloses a combustor casing having a concentric air bleed structure which includes a series of conical ~~ ~ ~~ C
:'tUC),.a~
arms 62, 64, and 66 situated in the low velocity area of the diffuser with the bled air structure making a turn of approximately 180°. U.S. Patent 4,344,282 to Anders is directed to a compressor bleed system which includes a locking strap 12 which seals a series of bleed ports 8. U.S. Patent 4,827,'713 to Peterson, et al., for a stator valve assembly for rotory machine which includes a passage 30 in the compressor bleed system 28. The structure disclosed in each of these patents significantly reduces the pressure or velocity of the extracted air and thus reduces the energy level of the diffuser air. These documents fail to teach or suggest a pressure efficient diffuser slot which maintains the energy and pressure level of the diffused air to allow the extraction of air from an earlier compressor stage yet having a pressure and energy level equivalent to air previously extracted from a later stage.
SUB~ARX OF THE II~1VBNTI()h1 It is therefore desirable to provide a bleed air structure capable of efficiently extracting compressor discharge air with a minimum energy loss and delivering the extracted air to external systems with little pressure loss and at as high a pressure as possible.
Briefly stated, the above and similarly related objects are obtained by providing a gas turbine engine which includes an axial flow, multistage compressor, a combustor, and a turbine. A high pressure compressor bleed air extraction slot structure is provided comprising a compressor outer band having a bleed air portion positioned proximate a rearward and preferably interstage section of a compressor. A diffusing slot can be disposed in the compressor outer casing and can comprise an articulated or punched-out portion of the outer band articulated at an angle approximately 10-20 degrees from a band baseline whereby the diffusion coefficient of the bleed valve is improved. In a s preferred embodiment, the articulated angle is 15 degrees and the exit velocity V2 is less than the baseline velocity V1 while the exit pressure P2 is greater than the baseline pressure P1.
BRIEF DESCRIPTION OF THE DRAWINGS
io While the specification concludes with a series of claims which particularly point out and distinctly claim the subject matter which applicants consider to be their invention, a more complete understanding of the invention will be gained from the following detailed i5 description which is given in connection with the accompanying drawings, in which FIGURE 1 is a greatly simplified schematic view taken in cross section of a turbofan engine having a previously proposed bleed valve;
2o FIGURE 2 is a greatly simplified schematic view taken partially in section of a turbofan engine incorporating a bleed valve in accordance with the present invention;
FIGURE 3a is an enlarged schematic illustration of 2s typical bleed valves in accordance with FIGURE 1 illustrating the theoretical airflows associated therewith;
FIGURE 3b is an enlarged schematic illustration of the bleed valve in accordance with FIGURE 2 of the 3o present invention; and, FIGURE 3c is a sectional view taken at section l3DV-10107 - 4a -line 3C-3C of FIGURE 3b.
DETAILED DESCRIPTION OF THE DRAWINGS
Referring to the drawings wherein the numerals correspond to like elements throughout, attention is s directed initially to FIGURE 1 wherein a gas turbine engine 10 is shown in major cross section to include a fan rotor 12, and a core engine rotor 14. The fan rotor 12 includes a plurality of fan blades 16 mounted -. J (] ,. a _5_ for rotation on a disk 20. The fan rotor 12 also includes a low pressure or fan turbine 22 which drives the fan disk 20 in a well known manner. The core engine rotor 14 includes a compressor 24 and a high 5 power or high pressure turbine 26 which drives the compressor 24.' The core engine also includes a combustion system 28.
In operation, air enters the gas turbine 10 through an inlet 30 provided by means of a suitable cowling 32 which surrounds the fan rotor 12 and core engine rotor 14 and provides the external casing for the engine. Air entering the inlet 30 is compressed by means of the rotation of fan blades 16 and thereafter is split into two flow streams, a bypass 15 stream 34 flowing in a bypass passageway 35, and a core engine stream 36 flowing in a core passageway 37.
The pressurized air which enters the core engine passageway 37 is f~xrther pressurized by means of the compressor 24 and is thereafter mixed and ignited 20 along with high energy fuel in the combustion system 28. This highly energized gas stream then flows through the high pressure turbine 26 to drive the compressor 24 and thereafter through the low pressure turbine 22 to drive the fan rotor l2 and disk 20. The 25 pressurized air flowing through the bypass passageway 35 is either mixed with the core engine exhaust system stream by means of a suitable mixer (not shown) or is allowed to exhaust to ambient conditions as a relatively low velocity, low pressure stream 30 surrounding the core engine exhaust. In either case, the core engine stream 36 exhaust and fan bypass stream 34 exhaust provide a propulsive force for an aircraft powered by the turbofan engine l0.
It should be noted that although the present 35 description is limited to an aircraft gas turbine ,~ ~.~ m.
engine, the present invention may be applicable to any gas turbine engine powerplant such as those utili2ed for marine or industrial usage. A description of the engine shown in Fig. 1 is thus merely illustrative of the type of engine to which the present invention is applicable.
As shown in Figs. 1 and 3A, a diffusing port ar hole 40 comprises an orifice 42 located in line with an outer band 44 of the engine cowling or casing 32.
The compressor casing structure 44 provides an annular orifice 42 immediately upstream of one of the intermediate stages of the rotor blades 38 for bleeding inner stage air from the interior of the compressor 24.
Referring now to Figs. 2 and 3B, the details of the inventive diffusion slot and bleed air structure in accordance with the present invention are shown in an enlarged cross-sectional view of the compressor 24. As shown therein, the compressor 24 includes a rotor 14 having a number of rotor stages 40 which carry a plurality of rotor blades 38. The cpmpressor 24 further includes a casing structure 32 which defines the outer bounds of the compressor flowpath and inoludes mounting provisions for a plurality of stator vanes 46 aligned in individual stages between each stage of rotor blades 38.
In accordance with a preferred embodiment, shown in Fig. 3, the outer band 44 includes a diffuser slot 62 comprising a punched-out and articulated portion 64 articulated at an angle of between to and 20 degrees and preferably 15 degrees measured from a baseline 60 of the outer band 44.
Referring now t~ Figs. 3A and 3B in combination, a comparison of the prior annular 5th stage orifice 42 ~3DV-a.ozo~
is shown in Fig. 3A in relation to the present articulated 4th stage diffuser slot 62 in accordance with the present invention, shown in Fig. 3B. More particularly, as illustrated in Fig. 3A, the annular orifice 42 induces a swirling airflow 50 which . substantially restricts the opening of orifice 42 and reduces the discharge coefficient Cd associated with the orifice, Moreover, the annular orifice 42 requires the exiting air to alter its velocity by approximately ~0 degrees with a concommitant energy reduction.
In contrast, a diffusing slot 62 in accordance with the present invention, which is shown in Fig. 3B, includes an articulate portion 64 which expands the volume of a lateral cavity 54 of the compressor vane to cause the cavity to immediately capture diffuser air and minimally change the velocity and energy level of the captured air. The volume of the lateral cavity 54 is considered to be the volume between the casing baseline 60 and the articulated member 64. As is illustrated, the swirl pattern established by this slot 62 occurs closely adjacent the slot°s surfaces 44 and 64 and thus introduces a minimal obstruction to the air flowpath. Accordingly, the pressure drop associated with the slot 62 is minimized, the discharge coefficient, Cd, associated with this slot is maximized and the energy level of air passing through the diffuser is maintained. The efficient energy conversion achieved by this slot produces air at a higher pressure than that previously achieved.
Accordingly, the slot 62 can b~ applied to an earlier or lower pressure stage of the compressor and yet still supply air of a pressure equivalent to that previously derived from a later stage. The bleed slot ~~x.UU
v 62 of the present invention provides a means to recaver and convert a portion of the gas steam dynamic pressure into a manifold static pressure rise. The angled recessed surface of the articulated portion 64 5 acts as a diffuser to decelerate the air as it passes through the outer band opening thereby reducing the irreversible losses in energy.
The discharge coefficient Cd is defined as the ratio of actual mass flow to ideal mass flow through a restriction and can be expressed by the equation Cd = M1/M2. The invention can be characterized based on test data which shows clearly that a higher Cd is achieved for the diffusing slot 62 compared to a standard orif.i.ce 42 each having the same 15 cross-sectional area. More particularly, in a typical 9-stage compressor, the prior orifice 42 when applied to the 5th stage could achieve a discharge pressure of 132 Asia (16/in2) at temperature of 1207°R (Rankine).
In contrast, the present invention, when applied to 20 the 4th stage of the same compressor, can achieve a discharge pressure of 118 Asia at temperature of 1089°R; thus, improving the efficiency of the engine.
Accordingly, the diffuser extraction slot 62 of the present invention allows a portion of the gas 25 flowpath velocity pressure to be recovered as usable manifold static pressure: This higher pressurized flow allows the bleed extraction point to be relocated at least one stage forward in the compressor and represents an overall increase.in efficiency and 30 engine performance which can be reflected in lower specific fuel consumption. In addition, the extraction of air earlier in the compressor provides a lower temperature source for turbine cooling systems.
Although the invention has been shown and 35 described in detail with respect to the preferred Zr ~ U ~ Ll -9_ embodiments thereof, it should be recognized by those skilled in the art that various changes in the form and detail thereof may be made without departing from the true spirit arid scope of the present invention.
Accordingly, the size and location of the diffuser slot can be changed to reflect the pressure drop and flow requirements of the systems) that the bleed slots supplies. Further, the shape of the orifice can be changed such that the pressure gradient across the opening can be minimized to insure a high pressure flow. Accordingly, the bleed diffuser slot construction of the present invention can be adapted to fit a number of gas turbine engines as described herein.
It will be readily understood by those skilled in the art that the present invention is not limited to the specific embodiments described and illustrated herein. Different embodiments and adaptations besides those shown herein and described, as well as many variations, modifications and equivalent arrangements will now be apparent or will be reasonably suggested by the foregoing specification and drawings, without departing from the substance or scope of the invention. While the present invention has been described herein in detail in relation to its preferred embodiments, it is to be understood that this disclosure is only illustrative and exemplary of the present invention and is made merely for purposes of providing a full and enabling disclosure of the invention: Accordingly, it is intended that the invention be limited only by the spirit and scape of the claims appended hereto.
HTGH PRESSURE OOMPRESSOIt FLOWPATki BLEED VALVE EXTRAG"fION SIFT
This invention relates to simplified bleed extraction slots for, gas turbine engines and, more particularly, to a specially configured bleed extraction slot for efficiently converting core air to bleed air with a minimum loss in bleed air velocity and pressure.
BACItGROITPTD OF THE INVENTION
It is often desirable that an aircraft gas turbine engine include within its compressor, a structure which permits bleeding or diversion of high pressure air from a stage, such as the 5th stage of the compressor to provide high pressure air for cooling purposes and for operation of airframe accessories, engine accessories, or engine or aircraft de-icing systems. In other cases, i~t is desirable to include a struoture which pezmits the bleeding of even higher pressure air from the discharge of the compressor to provide pressurized air far cooling downstream turbine components. Both interstage bleed and the compressor discharge bleeding are accomplished by flowpath mechanisms which interfere with the normal airflow patterns in the compressor. Further, the casing or bleed structure adds complexity to the assembly of such an engine.
sy 9~. ! t r~ s G.~ Z;~ r~ ~ (.3 13DV°10107 °2°
The axial location or stage at which air is bled from the compressor is determined by the pressure required to drive the specific system intended to be serviced by that air. In most instances, it is desirable to achieve the highest possible source pressure to also ensure a high delivery pressure. For this reason, prior systems have extracted air from the latter stages of the compressor and more particularly, engines having these systems have been designed to extract high pressure air from the 5th stage of the compressor for low pressure turbine cooling and turbine thermal clearance control. However, bleeding air from the earliest possible stage of the compressor generally increases compressor efficiency by reducing the amount of work invested in the extracted air.
Therefore, it is desirable to achieve the highest possible system supply pressure from the earliest and lowest pressure stage of the compressor. The resulting temperature of the cooling air is also lower and hence more effective.
Known examples of bleed openings or ports can be found in U.S. Patent 4,711,084 to Brockett for an ejector°assisted compressor bleed which discloses a bleed aperture 17 in Fig. 2 having rounded hole edges. U.S. Patent 3,108,767 to Eltis, et al., for a bypass gas turbine engine with an air bleed means in Fig. 3 discloses a duct 19 which is attached to the compressor through a series of chopped holes. U.S.
Patent 3,898,799 to Pollert, et al., for a device for bleeding off compressor air in a turbine jet engine, in Fig. 2 discloses a compressor orifice marked with the arrow K. U.S. Patent 3,777,489 to 0'ohnston, et al., discloses a combustor casing having a concentric air bleed structure which includes a series of conical ~~ ~ ~~ C
:'tUC),.a~
arms 62, 64, and 66 situated in the low velocity area of the diffuser with the bled air structure making a turn of approximately 180°. U.S. Patent 4,344,282 to Anders is directed to a compressor bleed system which includes a locking strap 12 which seals a series of bleed ports 8. U.S. Patent 4,827,'713 to Peterson, et al., for a stator valve assembly for rotory machine which includes a passage 30 in the compressor bleed system 28. The structure disclosed in each of these patents significantly reduces the pressure or velocity of the extracted air and thus reduces the energy level of the diffuser air. These documents fail to teach or suggest a pressure efficient diffuser slot which maintains the energy and pressure level of the diffused air to allow the extraction of air from an earlier compressor stage yet having a pressure and energy level equivalent to air previously extracted from a later stage.
SUB~ARX OF THE II~1VBNTI()h1 It is therefore desirable to provide a bleed air structure capable of efficiently extracting compressor discharge air with a minimum energy loss and delivering the extracted air to external systems with little pressure loss and at as high a pressure as possible.
Briefly stated, the above and similarly related objects are obtained by providing a gas turbine engine which includes an axial flow, multistage compressor, a combustor, and a turbine. A high pressure compressor bleed air extraction slot structure is provided comprising a compressor outer band having a bleed air portion positioned proximate a rearward and preferably interstage section of a compressor. A diffusing slot can be disposed in the compressor outer casing and can comprise an articulated or punched-out portion of the outer band articulated at an angle approximately 10-20 degrees from a band baseline whereby the diffusion coefficient of the bleed valve is improved. In a s preferred embodiment, the articulated angle is 15 degrees and the exit velocity V2 is less than the baseline velocity V1 while the exit pressure P2 is greater than the baseline pressure P1.
BRIEF DESCRIPTION OF THE DRAWINGS
io While the specification concludes with a series of claims which particularly point out and distinctly claim the subject matter which applicants consider to be their invention, a more complete understanding of the invention will be gained from the following detailed i5 description which is given in connection with the accompanying drawings, in which FIGURE 1 is a greatly simplified schematic view taken in cross section of a turbofan engine having a previously proposed bleed valve;
2o FIGURE 2 is a greatly simplified schematic view taken partially in section of a turbofan engine incorporating a bleed valve in accordance with the present invention;
FIGURE 3a is an enlarged schematic illustration of 2s typical bleed valves in accordance with FIGURE 1 illustrating the theoretical airflows associated therewith;
FIGURE 3b is an enlarged schematic illustration of the bleed valve in accordance with FIGURE 2 of the 3o present invention; and, FIGURE 3c is a sectional view taken at section l3DV-10107 - 4a -line 3C-3C of FIGURE 3b.
DETAILED DESCRIPTION OF THE DRAWINGS
Referring to the drawings wherein the numerals correspond to like elements throughout, attention is s directed initially to FIGURE 1 wherein a gas turbine engine 10 is shown in major cross section to include a fan rotor 12, and a core engine rotor 14. The fan rotor 12 includes a plurality of fan blades 16 mounted -. J (] ,. a _5_ for rotation on a disk 20. The fan rotor 12 also includes a low pressure or fan turbine 22 which drives the fan disk 20 in a well known manner. The core engine rotor 14 includes a compressor 24 and a high 5 power or high pressure turbine 26 which drives the compressor 24.' The core engine also includes a combustion system 28.
In operation, air enters the gas turbine 10 through an inlet 30 provided by means of a suitable cowling 32 which surrounds the fan rotor 12 and core engine rotor 14 and provides the external casing for the engine. Air entering the inlet 30 is compressed by means of the rotation of fan blades 16 and thereafter is split into two flow streams, a bypass 15 stream 34 flowing in a bypass passageway 35, and a core engine stream 36 flowing in a core passageway 37.
The pressurized air which enters the core engine passageway 37 is f~xrther pressurized by means of the compressor 24 and is thereafter mixed and ignited 20 along with high energy fuel in the combustion system 28. This highly energized gas stream then flows through the high pressure turbine 26 to drive the compressor 24 and thereafter through the low pressure turbine 22 to drive the fan rotor l2 and disk 20. The 25 pressurized air flowing through the bypass passageway 35 is either mixed with the core engine exhaust system stream by means of a suitable mixer (not shown) or is allowed to exhaust to ambient conditions as a relatively low velocity, low pressure stream 30 surrounding the core engine exhaust. In either case, the core engine stream 36 exhaust and fan bypass stream 34 exhaust provide a propulsive force for an aircraft powered by the turbofan engine l0.
It should be noted that although the present 35 description is limited to an aircraft gas turbine ,~ ~.~ m.
engine, the present invention may be applicable to any gas turbine engine powerplant such as those utili2ed for marine or industrial usage. A description of the engine shown in Fig. 1 is thus merely illustrative of the type of engine to which the present invention is applicable.
As shown in Figs. 1 and 3A, a diffusing port ar hole 40 comprises an orifice 42 located in line with an outer band 44 of the engine cowling or casing 32.
The compressor casing structure 44 provides an annular orifice 42 immediately upstream of one of the intermediate stages of the rotor blades 38 for bleeding inner stage air from the interior of the compressor 24.
Referring now to Figs. 2 and 3B, the details of the inventive diffusion slot and bleed air structure in accordance with the present invention are shown in an enlarged cross-sectional view of the compressor 24. As shown therein, the compressor 24 includes a rotor 14 having a number of rotor stages 40 which carry a plurality of rotor blades 38. The cpmpressor 24 further includes a casing structure 32 which defines the outer bounds of the compressor flowpath and inoludes mounting provisions for a plurality of stator vanes 46 aligned in individual stages between each stage of rotor blades 38.
In accordance with a preferred embodiment, shown in Fig. 3, the outer band 44 includes a diffuser slot 62 comprising a punched-out and articulated portion 64 articulated at an angle of between to and 20 degrees and preferably 15 degrees measured from a baseline 60 of the outer band 44.
Referring now t~ Figs. 3A and 3B in combination, a comparison of the prior annular 5th stage orifice 42 ~3DV-a.ozo~
is shown in Fig. 3A in relation to the present articulated 4th stage diffuser slot 62 in accordance with the present invention, shown in Fig. 3B. More particularly, as illustrated in Fig. 3A, the annular orifice 42 induces a swirling airflow 50 which . substantially restricts the opening of orifice 42 and reduces the discharge coefficient Cd associated with the orifice, Moreover, the annular orifice 42 requires the exiting air to alter its velocity by approximately ~0 degrees with a concommitant energy reduction.
In contrast, a diffusing slot 62 in accordance with the present invention, which is shown in Fig. 3B, includes an articulate portion 64 which expands the volume of a lateral cavity 54 of the compressor vane to cause the cavity to immediately capture diffuser air and minimally change the velocity and energy level of the captured air. The volume of the lateral cavity 54 is considered to be the volume between the casing baseline 60 and the articulated member 64. As is illustrated, the swirl pattern established by this slot 62 occurs closely adjacent the slot°s surfaces 44 and 64 and thus introduces a minimal obstruction to the air flowpath. Accordingly, the pressure drop associated with the slot 62 is minimized, the discharge coefficient, Cd, associated with this slot is maximized and the energy level of air passing through the diffuser is maintained. The efficient energy conversion achieved by this slot produces air at a higher pressure than that previously achieved.
Accordingly, the slot 62 can b~ applied to an earlier or lower pressure stage of the compressor and yet still supply air of a pressure equivalent to that previously derived from a later stage. The bleed slot ~~x.UU
v 62 of the present invention provides a means to recaver and convert a portion of the gas steam dynamic pressure into a manifold static pressure rise. The angled recessed surface of the articulated portion 64 5 acts as a diffuser to decelerate the air as it passes through the outer band opening thereby reducing the irreversible losses in energy.
The discharge coefficient Cd is defined as the ratio of actual mass flow to ideal mass flow through a restriction and can be expressed by the equation Cd = M1/M2. The invention can be characterized based on test data which shows clearly that a higher Cd is achieved for the diffusing slot 62 compared to a standard orif.i.ce 42 each having the same 15 cross-sectional area. More particularly, in a typical 9-stage compressor, the prior orifice 42 when applied to the 5th stage could achieve a discharge pressure of 132 Asia (16/in2) at temperature of 1207°R (Rankine).
In contrast, the present invention, when applied to 20 the 4th stage of the same compressor, can achieve a discharge pressure of 118 Asia at temperature of 1089°R; thus, improving the efficiency of the engine.
Accordingly, the diffuser extraction slot 62 of the present invention allows a portion of the gas 25 flowpath velocity pressure to be recovered as usable manifold static pressure: This higher pressurized flow allows the bleed extraction point to be relocated at least one stage forward in the compressor and represents an overall increase.in efficiency and 30 engine performance which can be reflected in lower specific fuel consumption. In addition, the extraction of air earlier in the compressor provides a lower temperature source for turbine cooling systems.
Although the invention has been shown and 35 described in detail with respect to the preferred Zr ~ U ~ Ll -9_ embodiments thereof, it should be recognized by those skilled in the art that various changes in the form and detail thereof may be made without departing from the true spirit arid scope of the present invention.
Accordingly, the size and location of the diffuser slot can be changed to reflect the pressure drop and flow requirements of the systems) that the bleed slots supplies. Further, the shape of the orifice can be changed such that the pressure gradient across the opening can be minimized to insure a high pressure flow. Accordingly, the bleed diffuser slot construction of the present invention can be adapted to fit a number of gas turbine engines as described herein.
It will be readily understood by those skilled in the art that the present invention is not limited to the specific embodiments described and illustrated herein. Different embodiments and adaptations besides those shown herein and described, as well as many variations, modifications and equivalent arrangements will now be apparent or will be reasonably suggested by the foregoing specification and drawings, without departing from the substance or scope of the invention. While the present invention has been described herein in detail in relation to its preferred embodiments, it is to be understood that this disclosure is only illustrative and exemplary of the present invention and is made merely for purposes of providing a full and enabling disclosure of the invention: Accordingly, it is intended that the invention be limited only by the spirit and scape of the claims appended hereto.
Claims
1. An energy efficient compressor air bleed structure for use in an axial flow compressor including a compressor casing having an outer band, said air bleed structure comprising:
a) a means for recovering and converting a portion of a gas stream dynamic pressure into a manifold static pressure rise;
b) wherein said means for recovering and converting comprises an articulated diffuser slot including an articulated portion of said compressor casing outer band;
and c) wherein said articulated portion i) is articulated at an angle of approximately 10-20 degrees from a compressor casing outer band baseline, and ii) includes an angled recessed surface which acts as a diffuser to decelerate bleed air as it passes through said articulated diffuser slot, and wherein said articulated portion comprises a punched-out portion of said compressor casing outer band.
a) a means for recovering and converting a portion of a gas stream dynamic pressure into a manifold static pressure rise;
b) wherein said means for recovering and converting comprises an articulated diffuser slot including an articulated portion of said compressor casing outer band;
and c) wherein said articulated portion i) is articulated at an angle of approximately 10-20 degrees from a compressor casing outer band baseline, and ii) includes an angled recessed surface which acts as a diffuser to decelerate bleed air as it passes through said articulated diffuser slot, and wherein said articulated portion comprises a punched-out portion of said compressor casing outer band.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US61567690A | 1990-11-19 | 1990-11-19 | |
| US615,676 | 1990-11-19 |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CA2048829A1 CA2048829A1 (en) | 1992-05-20 |
| CA2048829C true CA2048829C (en) | 2001-12-18 |
Family
ID=24466390
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002048829A Expired - Fee Related CA2048829C (en) | 1990-11-19 | 1991-08-08 | High pressure compressor flowpath bleed valve extraction slot |
Country Status (4)
| Country | Link |
|---|---|
| EP (1) | EP0487242B1 (en) |
| JP (1) | JP2513954B2 (en) |
| CA (1) | CA2048829C (en) |
| DE (1) | DE69113209T2 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9528391B2 (en) | 2012-07-17 | 2016-12-27 | United Technologies Corporation | Gas turbine engine outer case with contoured bleed boss |
Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9260974B2 (en) | 2011-12-16 | 2016-02-16 | General Electric Company | System and method for active clearance control |
| US9145772B2 (en) | 2012-01-31 | 2015-09-29 | United Technologies Corporation | Compressor disk bleed air scallops |
| JP6000142B2 (en) * | 2013-01-28 | 2016-09-28 | 三菱重工業株式会社 | Rotating machine and gas turbine provided with the same |
| EP2803822B1 (en) * | 2013-05-13 | 2019-12-04 | Safran Aero Boosters SA | Air-bleeding system of an axial turbomachine |
| JP6134628B2 (en) | 2013-10-17 | 2017-05-24 | 三菱重工業株式会社 | Axial flow compressor and gas turbine |
| GB201518448D0 (en) | 2015-10-19 | 2015-12-02 | Rolls Royce | Compressor |
| CN113847280B (en) * | 2021-10-10 | 2024-08-02 | 中国航发沈阳发动机研究所 | Interstage air entraining structure of compressor rotor |
Family Cites Families (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3108767A (en) * | 1960-03-14 | 1963-10-29 | Rolls Royce | By-pass gas turbine engine with air bleed means |
| GB936635A (en) * | 1961-04-21 | 1963-09-11 | Rolls Royce | Multi-stage axial-flow compressor |
| US3777489A (en) * | 1972-06-01 | 1973-12-11 | Gen Electric | Combustor casing and concentric air bleed structure |
| DE2247400C2 (en) * | 1972-09-27 | 1975-01-16 | Motoren- Und Turbinen-Union Muenchen Gmbh, 8000 Muenchen | Device for blowing off compressed air from a compressor of a gas turbine jet engine |
| JPS5364112A (en) * | 1976-11-19 | 1978-06-08 | Hitachi Ltd | Gas turbine compressor |
| US4344282A (en) * | 1980-12-16 | 1982-08-17 | United Technologies Corporation | Compressor bleed system |
| US4711084A (en) * | 1981-11-05 | 1987-12-08 | Avco Corporation | Ejector assisted compressor bleed |
| US4546605A (en) * | 1983-12-16 | 1985-10-15 | United Technologies Corporation | Heat exchange system |
| JPS6124675U (en) * | 1984-07-17 | 1986-02-14 | 日本電気株式会社 | Withstand voltage tester |
| FR2640685B1 (en) * | 1988-12-15 | 1991-02-08 | Snecma | TURBOREACTOR COMPRESSOR DISCHARGE VALVE |
-
1991
- 1991-08-08 CA CA002048829A patent/CA2048829C/en not_active Expired - Fee Related
- 1991-11-13 EP EP91310455A patent/EP0487242B1/en not_active Expired - Lifetime
- 1991-11-13 DE DE69113209T patent/DE69113209T2/en not_active Expired - Fee Related
- 1991-11-14 JP JP3325119A patent/JP2513954B2/en not_active Expired - Fee Related
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9528391B2 (en) | 2012-07-17 | 2016-12-27 | United Technologies Corporation | Gas turbine engine outer case with contoured bleed boss |
Also Published As
| Publication number | Publication date |
|---|---|
| JP2513954B2 (en) | 1996-07-10 |
| CA2048829A1 (en) | 1992-05-20 |
| EP0487242A1 (en) | 1992-05-27 |
| DE69113209D1 (en) | 1995-10-26 |
| JPH04284136A (en) | 1992-10-08 |
| EP0487242B1 (en) | 1995-09-20 |
| DE69113209T2 (en) | 1996-05-02 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US5209633A (en) | High pressure compressor flowpath bleed valve extraction slot | |
| CA2089278C (en) | Gas turbine engine cooling system | |
| US5269135A (en) | Gas turbine engine fan cooled heat exchanger | |
| CA1052106A (en) | Split fan work gas turbine engine | |
| US6735956B2 (en) | High pressure turbine blade cooling scoop | |
| CA1108872A (en) | Variable area bypass injector for a double bypass variable cycle gas turbofan engine | |
| EP2519723B1 (en) | Gas turbine engine | |
| US4329114A (en) | Active clearance control system for a turbomachine | |
| US3966354A (en) | Thermal actuated valve for clearance control | |
| CA2107429C (en) | Gas turbine engine cooling systems | |
| US5984630A (en) | Reduced windage high pressure turbine forward outer seal | |
| US4893984A (en) | Clearance control system | |
| US3777489A (en) | Combustor casing and concentric air bleed structure | |
| US7200999B2 (en) | Arrangement for bleeding the boundary layer from an aircraft engine | |
| US5531568A (en) | Turbine blade | |
| US5316437A (en) | Gas turbine engine structural frame assembly having a thermally actuated valve for modulating a flow of hot gases through the frame hub | |
| US9982629B2 (en) | Engine driven by SC02 cycle with independent shafts for combustion cycle elements and propulsion elements | |
| EP1173656B1 (en) | High pressure turbine cooling of gas turbine engine | |
| US6513330B1 (en) | Diffuser for a gas turbine engine | |
| US6401447B1 (en) | Combustor apparatus for a gas turbine engine | |
| US20080016878A1 (en) | Aeroengine bleed valve | |
| EP0212834A2 (en) | Variable inlet for a radial turbine | |
| GB1535681A (en) | Turbofan engine nacelle | |
| US4302148A (en) | Gas turbine engine having a cooled turbine | |
| CA2048829C (en) | High pressure compressor flowpath bleed valve extraction slot |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request | ||
| MKLA | Lapsed |