JP2017530299A - Boundary layer control of the diffuser passage of a centrifugal compressor - Google Patents

Abstract

The diffuser 42 of the centrifugal compressor extends through the annular diffuser housing 20, and has a plurality of diffuser flow passages having the diffuser vane 23 as a circumferential boundary and the front and rear walls 101, 100 as axial boundaries. 22 is included. The diffuser boundary layer bleed means 96 faces the downstream side inclined at an acute angle with respect to the downstream direction of the diffuser air flow 103 in the passage and the boundary layer bleed opening 106 or slot 130 disposed through the front wall. Includes a wall 142. The diffuser bleed air flow 112 is extracted from the diffuser boundary layer. The boundary layer bleed opening is located downstream of the throat 28 near the pressure side of the vane. The centrifugal compressor 18 mixes the impeller bleed air 102 from the diffuser surrounding the annular centrifugal compressor impeller 32 and the radial gap between the impeller tip 36 and the annular inlet 27 of the diffuser with the diffuser bleed air or And a device for cooling the turbine 16 separately. [Selection] Figure 1

Description

  The present invention relates to bleed air from a centrifugal compressor of a gas turbine engine.

  Some gas turbine engines include a centrifugal compressor that has a rotatable impeller and accelerates the air flowing therethrough, thereby increasing kinetic energy. The diffuser is generally placed immediately downstream of the impeller and surrounding it. The diffuser reduces the velocity of the air flow exiting the impeller and converts its energy to increase the static pressure and thus pressurize the air.

  Conventional gas turbine engines typically include a compressor, a combustor, and a turbine, with rotating turbine components such as blades, disks, and retainers, and stationary turbine components such as vanes, shrouds, and frames. Both are heated by high-temperature combustion gas, so cooling is necessary on a daily basis. Cooling the turbine, especially rotating components, is important for proper engine functioning and safe operation. It is known that cooling air is extracted from a centrifugal compressor to help cool the turbine.

  For example, if cooling air with insufficient supply pressure, volumetric flow, or temperature margins cannot provide sufficient cooling of the turbine disk and its cascade, it can adversely affect turbine life and mechanical integrity. is there. Depending on the nature and extent of the inadequate cooling, the impact on engine operation may result in unscheduled engine shutdowns from relatively harmless blade tip damage that results in reduced engine power and reduced blade service life. Can result in turbine disk breakage.

  A higher level of engine operating efficiency is desirable, which leads to lower fuel consumption and lower operating costs, balanced with the need to sufficiently cool the turbine. Since turbine cooling air is typically taken from one or more stages of the compressor and sent to the desired components by various means such as pipes, ducts and internal passages, such air is It is mixed with fuel, ignited in the combustor, and cannot be used to extract work in the main gas flow path of the turbine.

  The total cooling flow extracted from the compressor is a loss in the engine operating cycle and it is desirable to minimize such loss.

  Some conventional engines utilize a clean air bleed system to cool the turbine components of a gas turbine using an axial flow-centrifugal compressor, as is done with the General Electric CFE738 engine. Turbine cooling feed air exits the centrifugal diffuser through a narrow gap between the diffuser outlet and the dewar inner shroud. Other turbine cooling air methods include taking cooling air from the impeller or from the gap between the impeller and the diffuser outlet.

  In U.S. Pat. No. 5,555,721, entitled “A Gas Turbine Engine Cooling Supply Circuit”, granted to Bourneuuf et al. On September 17, 1996, in a turbine cooling supply circuit of a gas turbine, The use of bleed from the impeller stage is disclosed. U.S. Pat. No. 5,555,721 discloses an impeller tip forward bleed air flow and an impeller tip bleed air flow to cool turbine components. U.S. Pat. No. 5,555,721 is issued by General Electric Co., the same as this patent. And is hereby incorporated by reference.

  U.S. Pat. No. 8,087,249, entitled “A Turbine Cooling Air From A Centrifugal Compressor”, granted to Ottavianof et al. On Jan. 3, 2012, and a diffuser immediately below the impeller and impeller. A turbine cooling system for a gas turbine engine including bleed means for bleed clean cooling air from downstream is disclosed. U.S. Pat. No. 8,087,249 is issued by General Electric Co., the same as this patent. And is hereby incorporated by reference.

  Thus, there is a continuing need to develop diffuser designs and geometries that improve aerodynamic performance and reduce the overall radial profile of the engine.

U.S. Pat. No. 4,414,815

  The diffuser of the centrifugal compressor includes an annular diffuser housing, a diffuser vane extending axially between the front and rear walls of the diffuser housing, and extending through the housing and spaced around the periphery of the housing And a plurality of diffuser flow passages arranged in this manner. The diffuser flow passage is bounded by the diffuser vane and the front and rear walls. The diffuser boundary layer extraction means is provided for extracting a diffuser extraction air flow from each diffuser boundary layer of the diffuser flow passage.

  The diffuser boundary layer bleed means can be configured to bleed the diffuser bleed air flow from the diffuser boundary layer located in each weak region of the diffuser flow passage.

  The diffuser boundary layer bleed means may include a boundary layer bleed opening disposed through the front wall. Each of the boundary layer bleed openings has a slot including a downstream facing wall that is angled or inclined at an acute angle with respect to the downstream direction of the respective diffuser air flow in the diffuser flow passage. can do.

  The boundary layer bleed opening may be located or located near the pressure side of the diffuser vane and downstream of the throat portion of the diffuser flow passage.

  The centrifugal compressor includes an annular centrifugal compressor impeller, a diffuser that surrounds the impeller in an annular shape, and a plurality of diffuser flow passages that extend through the housing of the diffuser and are spaced apart around the periphery of the housing. Including. Each of the passages includes a throat portion and a speed reduction portion downstream of the throat portion. The diffuser flow passage uses a diffuser vane extending in the axial direction between the front wall and the rear wall of the diffuser as a circumferential boundary, and the diffuser boundary layer bleed means extracts a diffuser extraction from each diffuser boundary layer of the diffuser flow passage. It is provided to extract airflow.

  The centrifugal compressor also mixes the radial clearance between the impeller tip of the impeller and the annular inlet of the diffuser, and the impeller bleed flow from the radial clearance with the diffuser bleed flow from the boundary layer bleed opening to provide turbine cooling air. And means for flowing turbine cooling air to the turbine, or means for separately flowing the impeller bleed air and the diffuser bleed air to the turbine.

1 is a cross-sectional view of a gas turbine engine having a centrifugal compressor that mixes impeller tip bleed air and diffuser bleed air in a compressor section and then uses that flow to cool turbine components. FIG. 2 is an enlarged cross-sectional view of the centrifugal compressor and a diffuser having diffuser bleed holes shown in FIG. 1. It is the perspective view which looked at the front from the back of the diffuser which passes along 3-3 of FIG. 2, and a diffuser bleed hole. FIG. 4 is an enlarged perspective view of an extraction hole shown in FIG. 3. FIG. 3 is a perspective view of a part of the diffuser shown in FIG. 2 and a diffuser bleed hole. It is an expanded sectional view of the centrifugal compressor tip and diffuser bleed hole shown in FIG. FIG. 6 is a cross-sectional view of a gas turbine engine centrifugal compressor having an alternative arrangement for separately flowing impeller chip bleed to cool turbine components. FIG. 8 is a cross-sectional view of the gas turbine engine shown in FIG. 7 having an arrangement for separately flowing a diffuser bleed air to cool turbine components. FIG. 9 is an enlarged perspective view of one of the impeller extraction airflow ports shown in FIG. 7 and cut at 9-9 in FIG. 10. It is the perspective view which looked at the back from the front of the rear casing surrounding the centrifugal compressor and containing the impeller and the diffuser bleed air port shown in FIG. 7 and FIG. 8, respectively. FIG. 10 is a cut perspective view of an impeller bleed passage for one of the impeller bleed air ports shown in FIGS. 7 and 9. FIG. 13 is an enlarged perspective view of one of the diffuser bleed air ports shown in FIG. 8 and cut at 12-12 in FIG. FIG. 13 is a cut perspective view of the diffuser bleed passage shown in FIG. 8 and passing through one of the diffuser bleed ports cut at 12-12 in FIG.

  FIG. 1 shows a gas turbine engine high-pressure centrifugal compressor 18 of a high-pressure gas generator 10 of a gas turbine engine 8. The high pressure centrifugal compressor 18 is the final compressor stage of the high pressure compressor 14. The high pressure gas generator 10 has a high pressure rotor 12 that includes a high pressure compressor 14, a combustor 52, and a high pressure turbine 16 in a serial or flow relationship in the downstream direction. The rotor 12 is rotatably supported around the engine shaft 25 by engine frame bearings not shown herein.

  The exemplary embodiment of the high pressure compressor 14 shown herein includes a five-stage axial compressor 30 in front of the centrifugal compressor 18 having an annular centrifugal compressor impeller 32. An outlet guide vane 40 is disposed between the five-stage axial compressor 30 and the single-stage centrifugal compressor 18. Compressor discharge pressure (CDP) air 76 exits the impeller 32, passes through a diffuser 42 that annularly surrounds the impeller 32, and then passes through a deswirl cascade 44 to a combustion chamber in the combustor 52. Enter 45. The combustion chamber 45 is surrounded by annular radially outer and inner combustor casings 46, 47. Air 76 is typically mixed with fuel supplied by a plurality of fuel nozzles 48, ignited, and burned in an annular combustion region 50 bounded by annular radially outer and inner combustion liners 72, 73.

  As is known in the art, combustion results in hot combustion gas 54 that flows through high pressure turbine 16, rotates high pressure rotor 12, and then flows downstream for further work in low pressure turbine 78. Is removed and finally discharged. In the exemplary embodiment shown herein, the high pressure turbine 16 includes first and second high pressure turbine stages 55, 56 having first and second stage disks 60, 62 in a serial flow relationship in the downstream direction. . A high-pressure shaft 64 of the high-pressure rotor 12 connects the high-pressure turbine 16 to the impeller 32 in a rotational drive engagement state. The first stage nozzle 66 is immediately upstream of the first high pressure turbine stage 55 and the second stage nozzle 68 is immediately upstream of the second high pressure turbine stage.

  Referring to FIG. 1, compressor discharge pressure (CDP) air 76 is discharged from the impeller 32 of the centrifugal compressor 18 to burn fuel in the combustor 52 and to the first stage nozzle 66, And used to cool components of the turbine 16 that are exposed to the hot combustion gases 54, such as first and second stage shrouds 71, 69 surrounding the first and second high pressure turbine stages 55, 56, respectively. The high pressure compressor 14 includes a compressor rear casing 110 and a diffuser front casing 114, as more fully shown in FIGS. The compressor rear casing 110 generally surrounds the axial compressor 30 and the diffuser front casing 114 generally surrounds the centrifugal compressor 18 and supports the diffuser 42 immediately downstream of the centrifugal compressor 18. The compressor discharge pressure (CDP) air 76 is discharged from the impeller 32 of the centrifugal compressor 18 and directly enters the diffuser 42.

  With reference to FIGS. 2 and 3, the impeller 32 includes a plurality of centrifugal compressor blades 84 extending radially from the rotor disk portion 82. Opposite the compressor blade 84 and axially forward is an annular blade tip shroud 90. The shroud 90 is adjacent to the blade tips 86 of the compressor blade 84 and defines a blade tip clearance 80 therebetween. Diffuser 42 includes an annular diffuser housing 20 that has a plurality of tangentially disposed diffuser flow passages 22 extending radially therethrough and spaced around a perimeter 26 of housing 20. The diffuser air flow 103 flows downstream through the diffuser flow passage 22. The diffuser vane 23 extends in the axial direction between the front wall 101 and the rear wall 100 of the diffuser 42.

  Referring to FIGS. 2 and 3, the diffuser vane 23 extends between adjacent passages of the diffuser flow passage 22 in the circumferential direction. The diffuser flow passage 22 is partially defined by diffuser vanes 23 that are spaced apart in the circumferential direction, and these are the circumferential boundaries. Adjacent passages 22 meet each other at the radially inner inlet 24 of the passage 22 and define a pseudo vane-free annular inlet 27 for the diffuser 42. Each passage 22 further includes a throat portion 28 that is integral with it downstream of the inner inlet portion 24. Each passage 22 further includes a speed reducing portion 99 immediately downstream of the throat portion 28.

  With reference to FIGS. 2 and 6, the first cooling air source 92 of the centrifugal compressor for the turbine cooling air 88 is a narrow predetermined position located between the impeller tip 36 of the rotating impeller 32 and the annular inlet 27 of the stationary diffuser 42. The radial clearance (C). Impeller bleed air flow 102 from the radial gap (C) is collected in the radially inner manifold 104. The predetermined radial gap (C) is designed to receive the thermal and mechanical stretch of the impeller 32 and opens into the radial inner manifold 104, ie, is in fluid communication therewith.

  Referring to FIGS. 3-6, the diffuser air flow 103 on one side of a passage (eg, passage 22) of a multiple passage diffuser (eg, diffuser 42) downstream of the centrifugal impeller (eg, impeller 32) is as follows: It was often found to be weak and easy to peel. Separation in the passage may cause large losses and worsen the specific fuel consumption (SFC). This range, that is, the region 127 where the flow is weak, is also considered to be related to a surge that limits the flow range of the compressor.

  The second cooling air source 94 of the centrifugal compressor stage for the turbine cooling air 88 is a diffuser boundary layer bleed 96 for extracting the diffuser bleed air 112 from each diffuser boundary layer 113 of the diffuser flow passage 22 of the diffuser 42. including. Herein, a plurality of boundary layer bleed openings 106 are shown. Diffuser boundary layer bleed 96, also referred to as fluid bleed, helps reduce weak flow and limit or prevent undesired flow separation. The diffuser boundary layer extraction means 96 extracts the diffuser extraction airflow 112 from the diffuser boundary layer 113 into the radially outer manifold 116.

  The radially inner manifold 104 and the radially outer manifold 116 are in fluid communication so that impeller bleed 102 from the radially inner manifold 104 flows into the radially outer manifold 116. The impeller and diffuser bleed air 102, 112 are mixed within the radially outer manifold 116 to provide turbine cooling air 88. The turbine cooling air 88 is then conveyed or flowed from the radially outer manifold 116 to the high pressure turbine 16 through a plurality of circumferentially distributed manifold ports 117. Turbine cooling air 88 is sent or flowed from there by external piping (not shown) to cool the first and second stage shrouds 71, 69 (shown in FIG. 1).

  A substantially axially extending beam or strut 122 separates the radially inner manifold 104 and the radially outer manifold 116, and the impeller bleed air flow 102 from the radially inner manifold 104 into the radially outer manifold 116. Passes between struts 122 when flowing. The fluid bleed shown herein as diffuser boundary layer bleed 96 represents a small flow rate of less than 1% of the engine core flow. Fluid bleeds are extracted strategically near where weak flow begins to improve the overall performance of the diffuser.

  3-5, the boundary layer bleed opening 106 may be a hole or slot 130 that penetrates the front wall 101 of the diffuser 42, as shown herein. The boundary layer bleed opening 106 or slot 130 leads into and is in flow communication with the radially outer manifold 116. The slot 130 is located or located near the pressure side 126 of the diffuser vane 23 and downstream of the throat portion 28. In that position, without diffuser boundary layer bleed 96, the flow will begin to weaken or become unstable within the diffuser. This position is located in an area called a weak flow area 127. The width W of the slot can be determined by manufacturing constraints such as the minimum tool size. The slot length L can be chosen so that up to 3% of the engine core flow rate can be used.

  The slot 130 is ideally positioned so that the diffuser bleed air flow 112 is perpendicular to the front face 105 of the front wall 101 of the diffuser 42 in a radial plane 132 that passes through the engine centerline or axis 25 as shown in FIG. Should be at an angle that leaves However, this angle can be varied due to constraints such as the slot passing through the bend 134 of the front wall 101 of the diffuser 42 or extending very close thereto. As shown in FIG. 6, the slot 130 has radially outer and inner walls 136, 138, and upstream and downstream extending through the front wall 101, respectively, as shown in FIGS. Side walls 140, 142 are provided. The downstream facing wall 142 is designed to scoop boundary layer air 144 only within the diffuser boundary layer 113. Thus, the downstream facing wall 142 is angled at an acute inclination angle B of less than 90 ° with respect to the diffuser air flow 103 (in the diffuser flow passage 22 of the diffuser 42 downstream and parallel to the boundary layer air 144). Or be inclined. The acute inclination angle B seems to be 45 degrees. However, the acute tilt angle B is limited by the geometry and manufacturing constraints outside the diffuser, so that the acute tilt angle is more practical, for example, about 22.5 degrees.

  FIGS. 7-13 show a gas turbine engine having a centrifugal compressor similar to the gas turbine engine shown in FIGS. 1-3, but impeller chip bleed and diffuser extraction to cool the turbine components. It is an alternative arrangement or design that collects airflow separately. The impeller bleed air flow 102 from the radial gap (C) shown in FIG. 9 flows into the radial inner annular manifold 154 shown in FIGS. 7 and 9 and is collected. Inter-manifold opening 160 is disposed between inner manifold 154 and a plurality of radially outer annular manifolds 156 shown in FIGS. The inter-manifold opening 160 allows the impeller bleed air flow 102 to flow from the inner annular manifold 154 into the outer annular manifold 156. Then, the impeller bleed air flow 102 from the outer annular manifold 156 is conveyed or flowed to the high pressure turbine 16 for turbine cooling through the plurality of circumferentially distributed impeller bleed air manifold ports 157 shown in FIG. .

  Referring to FIGS. 8, 10 and 11-13, the diffuser boundary layer bleed means 96 bleeds diffuser bleed air 112 from the diffuser boundary layer 113 into the annular diffuser bleed manifold 158, and then diffuser bleed air 112 is from there Through a plurality of circumferentially distributed diffuser bleed manifold ports 159 to be carried or flowed to high pressure turbine 16 for turbine cooling. FIG. 10 shows the circumferential and axial relative positions of the impeller bleed manifold port 157 and diffuser bleed manifold port 159 on and through the diffuser front casing 114.

  While what has been considered to be preferred and exemplary embodiments of the invention has been described herein, other modifications of the invention should become apparent to those skilled in the art from the teachings herein. Accordingly, it is desired that all such modifications be protected within the scope of the appended claims as falling within the true spirit and scope of the invention. Accordingly, what is desired to be protected by Letters Patent of the United States is an invention as defined and differentiated in the following claims.

8 Gas Turbine Engine 10 High Pressure Gas Generator 12 High Pressure Rotor 14 High Pressure Compressor 16 High Pressure Turbine 18 Centrifugal Compressor 20 Diffuser Housing 22 Diffuser Flow Path 23 Diffuser Vane 24 Radial Inner Inlet Portion 25 Engine Shaft 26 Peripheral 27 Annular Inlet 28 Throat Portion 30 Axial compressor 32 Impeller 36 Impeller tip 40 Outlet guide vane 42 Diffuser 44 Deswar blade cascade 45 Combustion chamber 46 Radial outer combustor casing 47 Radial inner combustor casing 48 Fuel nozzle 50 Combustion region 52 Combustor 54 Hot combustion gas 55 First High Pressure Turbine Stage 56 Second High Pressure Turbine Stage 60 First Stage Disc 62 Second Stage Disk 64 High Pressure Shaft 66 First Stage Nozzle 68 Second Stage Nozzle 69 Second Stage Shroud 71 First Stage Shu Wood 72 Radial outer combustion liner 73 Radial inner combustion liner 76 Air 78 Low pressure turbine 80 Blade tip clearance 82 Rotor disk section 84 Centrifugal compressor blade 86 Blade tip 88 Turbine cooling air 90 Blade tip shroud 92 First cooling air source 94 Second cooling air source 96 Diffuser boundary layer bleed (means)
99 Deceleration section 100 Rear wall 101 Front wall 102 Impeller bleed air flow 103 Diffuser air flow 104 Radial inner manifold 105 Front face 106 Boundary layer bleed opening 110 Compressor rear casing 112 Diffuser bleed air flow 113 Diffuser boundary layer 114 Diffuser front casing 116 Radial direction Outer manifold 117 Manifold port 122 Strut 126 Pressure side 127 Flow weak region 130 Slot 132 Radial plane 134 Bend 136 Radial outer wall 138 Radial inner wall 140 Wall facing upstream 142 Wall facing downstream 144 Boundary Layer air 154 Radial inner annular manifold 156 Radial outer annular manifold 157 Impeller bleed air manifold port 158 Diffuser bleed manifold 1 59 Diffuser bleed manifold port 160 Manifold opening B Sharp angle C Radial gap L Slot length W Slot width

Claims (24)

A diffuser (42) of a centrifugal compressor (18) of a gas turbine engine (8), comprising:
An annular diffuser housing (20);
A diffuser vane (23) extending axially between a front wall (101) and a rear wall (100) of the diffuser housing (20);
A plurality of diffuser flow passages (22) extending through the housing (20) and spaced apart around a periphery (26) of the housing (20),
The diffuser vane (23) and a diffuser flow passage (22) bounded by the front wall (101) and the rear wall (100);
A diffuser (42) comprising a diffuser boundary layer extraction means (96) for extracting a diffuser extraction air flow (112) from each diffuser boundary layer (113) of the diffuser flow passage (22). The diffuser boundary layer bleed means configured to bleed the diffuser bleed air flow (112) from the diffuser boundary layer (113) located in each weak flow region (127) of the diffuser flow passage (22). The diffuser (42) of claim 1, further comprising 96). The diffuser (42) of claim 1, further comprising the diffuser boundary layer bleed means (96) including a boundary layer bleed opening (106) disposed through the front wall (101). A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The diffuser (42) of claim 3, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising a). The boundary layer bleed opening (106) disposed or located near the pressure side (126) of the diffuser vane (23) and downstream of the throat (28) of the diffuser flow passage (22). Item 4. A diffuser (42) according to item 3. A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). 6. The diffuser (42) of claim 5, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising: A centrifugal compressor (18) of a gas turbine engine (8), comprising:
An annular centrifugal compressor impeller (32);
A diffuser (42) annularly surrounding the impeller (32);
A plurality of diffuser flow passages (22) extending through the housing (20) of the diffuser (42) and spaced apart around the periphery (26) of the housing (20),
Each of the passages (22) includes a throat portion (28) and a speed reduction portion (99) downstream of the throat portion (28),
The diffuser flow passage (22) with a diffuser vane (23) extending in the axial direction between the front wall (101) and the rear wall (100) of the diffuser (42) as a circumferential boundary;
A centrifugal compressor (18), comprising a diffuser boundary layer extraction means (96) for extracting a diffuser extraction airflow (112) from each diffuser boundary layer (113) of the diffuser flow passage (22). The diffuser boundary layer bleed means configured to bleed the diffuser bleed air flow (112) from the diffuser boundary layer (113) located in each weak flow region (127) of the diffuser flow passage (22). The centrifugal compressor (18) of claim 7, further comprising 96). The centrifugal compressor (18) of claim 7, further comprising the diffuser boundary layer bleed means (96) including a boundary layer bleed opening (106) disposed through the front wall (101). A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The centrifugal compressor (18) of claim 9, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising: The boundary layer bleed opening (106) disposed or located near the pressure side (126) of the diffuser vane (23) and downstream of the throat (28) of the diffuser flow passage (22). Item 15. The centrifugal compressor (18) according to Item 10. A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The centrifugal compressor (18) of claim 11, further comprising each of said boundary layer extraction openings (106) being slots (130) comprising: A radial clearance (C) between the impeller tip (36) of the impeller (32) and the annular inlet (27) of the diffuser (42);
The impeller bleed air flow (102) from the radial gap (C) is mixed with the diffuser bleed air flow (112) from the boundary layer bleed air opening (106) to supply turbine cooling air (88), and the turbine Means for flowing cooling air (88) to the turbine (16), or
The centrifugal compressor (18) of claim 9, further comprising means for flowing the impeller bleed air (102) and the diffuser bleed air (112) separately to the turbine (16). A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The centrifugal compressor (18) of claim 13, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising The boundary layer bleed opening (106) disposed or located near the pressure side (126) of the diffuser vane (23) and downstream of the throat (28) of the diffuser flow passage (22). Item 14. The centrifugal compressor (18) according to item 13. A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The centrifugal compressor (18) of claim 15, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising A radial clearance (C) between an impeller tip (36) of the impeller (32) and an annular inlet (27) of the diffuser (42),
A radial clearance (C) in fluid communication with the radially inner manifold (104);
The boundary layer bleed opening (106) in fluid communication with the radially outer manifold (116);
The radially outer manifold (102) flows into the radially outer manifold (116) and mixes with the diffuser bleed air (112) to form turbine cooling air (88). Said radially inner manifold in fluid communication with (114);
The centrifugal compressor (18) of claim 9, further comprising means for flowing turbine cooling air (88) from the radially outer manifold (116). A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The centrifugal compressor (18) of claim 17, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising The boundary layer bleed opening (106) disposed or located near the pressure side (126) of the diffuser vane (23) and downstream of the throat (28) of the diffuser flow passage (22). Item 19. A centrifugal compressor (18) according to item 18. A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). 20. The centrifugal compressor (18) of claim 19, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising). A radial clearance (C) between an impeller tip (36) of the impeller (32) and an annular inlet (27) of the diffuser (42),
A radial clearance (C) in fluid communication with the radially inner annular manifold (154);
An inter-manifold opening (160) disposed between the inner annular manifold (154) and a plurality of radially outer annular manifolds (156);
The impeller bleed air flow (102) from the radial gap (C) was distributed in the circumferential direction in the diffuser forward casing (114) surrounding the centrifugal compressor (18) and through the diffuser forward casing (114). Means for conveying or flowing through the plurality of impeller bleed air manifold ports (157) to the high pressure turbine (16) for turbine cooling and an annular diffuser bleed manifold (158) in an annular diffuser bleed manifold The diffuser boundary layer bleed means (96) operable to bleed the diffuser bleed air flow (112) in (158);
The diffuser bleed air flow (112) is passed through a plurality of diffuser bleed manifold ports (159) distributed in a circumferential direction in the diffuser front casing (114) and in the diffuser front casing (114) to cool the turbine. The centrifugal compressor (18) of claim 19, further comprising means for conveying or flowing to the high pressure turbine (16) for storage. A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). The centrifugal compressor (18) of claim 21, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising: The boundary layer bleed opening (106) disposed or located near the pressure side (126) of the diffuser vane (23) and downstream of the throat (28) of the diffuser flow passage (22). Item 23. The centrifugal compressor (18) according to item 22. A wall (142) facing the downstream side that is angled or inclined at an acute inclination angle (B) with respect to the downstream direction of each diffuser air flow (103) in the diffuser flow passage (22). 24. The centrifugal compressor (18) of claim 23, further comprising each of said boundary layer bleed openings (106) being slots (130) comprising: