Gas turbine engine including compressor bleed system

Abstract

A gas turbine includes a compressor bleed system including a compressor shroud, a compressor bleed plenum, and a plurality of compressor bleed passages disposed through the compressor shroud, the compressor bleed passages providing fluid communication between a compressor airflow passage and the compressor bleed plenum. The plurality of compressor bleed channels are arranged circumferentially spaced apart from each other, where a first compressor bleed channel and a second compressor bleed channel, which are circumferentially adjacent to each other, are spaced apart from each other by a first angular pitch, and corresponding inlets of at least a portion of a remaining portion of the plurality of compressor bleed channels are circumferentially spaced apart from each other by a second angular pitch, the second angular pitch is smaller than the first angular pitch.

Description

[0001] Government licensing rights

[0002] This invention was made with Government support under Award W58RGZ-19-C-0003 from the Department of Defense. The Government has certain rights in this invention. Technical Field

[0003] The present disclosure relates to a compressor bleed air system for a gas turbine. Background Art

[0004] Gas turbine engines typically include a compressor that compresses an incoming air flow before the compressed air flow is provided to a combustor. In some gas turbines, a compressor bleed air system may be provided. The compressor bleed air system may include a bleed air passage through a compressor shroud. A portion of the compressed air flow passing through the compressor may be discharged to a bleed air plenum through the bleed air passage. The air in the bleed air plenum may then be used for various other purposes, such as providing cooling air flow to other parts of the gas turbine engine, or providing air flow to an aircraft environmental system. BRIEF DESCRIPTION OF THE DRAWINGS

[0005] Features and advantages of the present disclosure will become apparent from the following description of various exemplary embodiments as illustrated in the accompanying drawings, wherein like reference numerals generally indicate identical, functionally similar, and / or structurally similar elements.

[0006] Figure 1 is a partial cross-sectional view of a gas turbine engine according to an aspect of the present disclosure.

[0007] Figure 2 According to one aspect of the present disclosure, Figure 1 Detailed view of the compressor bleed air passage taken at 84 .

[0008] Figure 3 According to one aspect of the present disclosure, Figure 1 A partial cross-sectional view of the circumferentially arranged compressor bleed air passages through the compressor shroud taken at plane 3-3 of FIG.

[0009] Figure 4 According to one aspect of the present disclosure, Figure 3 Detailed view 152 of the compressor is an enlarged view of the compressor bleed air passage.

[0010] Figure 5 Described is a Figure 2 An alternative arrangement of the compressor bleed air passage is shown.

[0011] Figure 6 A compressor bleed air passage replacement according to one aspect of the present disclosure Figure 3An alternative partial cross-sectional view of .

[0012] Figure 7 According to one aspect of the present disclosure, Figure 6 Detailed view of 170 Figure 6 An enlarged view of a portion of.

[0013] Figure 8 is a schematic partial cross-sectional view of an exemplary high bypass turbofan jet engine according to an aspect of the present disclosure.

[0014] Fig. 9 According to one aspect of the present disclosure, Figure 8 A partial cross-sectional view of an exemplary portion of a high pressure compressor taken at detailed view 200 of FIG.

[0015] Fig.10 According to one aspect of the present disclosure, Fig. 9 A partial cross-sectional view of a circumferentially arranged compressor bleed air passage through the compressor shroud taken at plane 10 - 10 of FIG.

[0016] Fig.11 According to another aspect of the present disclosure Fig. 9 A cross-sectional view of an alternative high-pressure compressor to the one shown.

[0017] Fig.12 According to another aspect of the present disclosure Fig. 9 A cross-sectional view of another alternative high-pressure compressor is shown. DETAILED DESCRIPTION

[0018] Features, advantages and embodiments of the present disclosure are set forth or apparent by considering the following detailed description, drawings and claims.In addition, the following detailed description is exemplary and is intended to provide further explanation without limiting the scope of the present disclosure as claimed.

[0019] Various embodiments are discussed in detail below. Although specific embodiments are discussed, this is for illustrative purposes only. Those skilled in the relevant art will recognize that other components and configurations may be used without departing from the spirit and scope of the present disclosure.

[0020] As used herein, the terms “first,” “second,” “third,” “fourth,” “fifth,” “sixth,” etc. may be used to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0021] Gas turbine engines mounted on aircraft typically include a compressor that compresses an incoming airflow before the compressed airflow is provided to a combustor. In some gas turbines, a compressor bleed air system may be provided. The compressor bleed air system may include a bleed air passage and a bleed air chamber passing through a compressor shroud. A portion of the compressed airflow passing through the compressor may be discharged to the bleed air chamber through the bleed air passage. The air in the bleed air chamber may then be used for various other purposes, such as providing cooling airflow to other parts of the gas turbine engine, or providing airflow to an aircraft environmental system on which the gas turbine engine is mounted.

[0022] In a conventional bleed air system, the bleed air passages may be evenly spaced around the compressor shroud circumference. The airflow through the bleed air passages may cause a wake in the airflow, which is felt as a pulse by the wing passing through. The combination of the number of bleed air passages (pulses) and the rotational speed of the compressor produces a stimulus that may be consistent with the natural frequency of the compressor rotor blades. That is, the uniform circumferential spacing of the bleed air passages may result in a repetition frequency around the compressor airflow passages, thereby introducing a stimulus on the rotating hardware (e.g., compressor rotor) upstream of the bleed air passages and / or downstream of the bleed air passages. In the present disclosure, non-uniform spacing of the bleed air passages is provided to facilitate changing the frequency, thereby reducing the vibration response caused in adjacent rotor blades. The frequency of the change reduces the repetition frequency, thereby reducing the vibration of the compressor rotor blades, which may cause metal fatigue of the compressor rotor blades. In addition, reducing the repetition frequency vibration may reduce the impact of vibration on other engine components.

[0023] With reference to the accompanying drawings, Figure 1 is a partial cross-sectional view of a gas turbine engine 10 according to one aspect of the present disclosure. Figure 1 Half of a gas turbine engine 10 is depicted, but the gas turbine engine 10 generally extends circumferentially about a longitudinal centerline axis 12. The gas turbine engine 10 includes, in sequential flow order, a compressor section 14, a combustion section 16, and a turbine section 18. As used herein, the terms "axial" and "longitudinal" both refer to directions (L) parallel to the longitudinal centerline axis 12, while the term "radial" refers to directions (R) perpendicular to or extending outward from the longitudinal centerline axis 12, and "tangential" or "circumferential" refers to directions (C) extending about the longitudinal centerline axis 12. As used herein, the terms "front" or "forward" refer to a relatively upstream position in an airflow through or around a component, and the terms "rear" or "rearward" refer to a relatively downstream position in an airflow through or around a component.

[0024] Generally speaking, in operation of the gas turbine engine 10, inlet air 20, schematically illustrated via arrows, enters the inlet end 22 of the compressor section 14, where the inlet air 20 is compressed by one or more compressor rotors 40 and impellers 42 of the compressor section 14 to produce compressed air 24. The compressed air 24 is forced radially outward through the compressor outlet duct 48 via the impellers 42, such that the compressor section 14 may be referred to as a centrifugal flow compressor. Although Figure 1 4. Although not shown in the drawing, the compressor outlet duct 48 may be in the form of a diffuser to reduce the flow velocity of the compressed air 24 while increasing the pressure of the compressed air 24 entering the combustion section 16. The compressed air 24 is provided to the combustion section 16 from the compressor section 14. A portion of the compressed air 24 flows through the swirler assembly 26 and mixes with the fuel (not shown) provided by the fuel nozzle assembly 28 to form a fuel / air mixture 30. The fuel / air mixture 30 is ignited in the combustion chamber 32 to produce combustion gases 34 that are discharged to the turbine section 18. The combustion gases 34 expand through the turbine nozzle 36 in the turbine section 18 and drive one or more turbine rotors 38, and then exit through the exhaust nozzle 46 (not shown, but only indicated by numbers). The turbine rotor 38 may be connected to a shaft 44, which may also be connected to the compressor rotor 40 and the impeller 42, so that the driving of the turbine rotor 38 by the combustion gases 34 supports the driving of the compressor rotor 40 and the impeller 42. The shaft 44 may also be connected to a propeller (not shown) located upstream of the compressor section 14 or downstream of the turbine section 18. Therefore, this gas turbine engine 10 can be a turboshaft engine. However, the principles described herein are equally applicable to any compressor. For example, they can be applied to compressors of gas turbine engines of different configurations, or to compressors driven by external prime movers.

[0025] The compressor section 14 includes an upstream portion 50 and a downstream portion 52. The upstream portion 50 of the compressor section 14 is configured for axial fluid flow and may also be referred to as an axial flow portion or simply an axial portion. The upstream portion 50 includes a plurality of stages ( Figure 1 ). Each stage includes a compressor rotor 40 including a plurality of rotor blades 54, and a circumferentially spaced row of stationary stator blades 56 connected to a compressor shroud 58. The compressor shroud 58 extends circumferentially about the longitudinal centerline axis 12 (which may therefore also correspond to the centerline axis of the compressor shroud 58), and thus the stator blades 56 are circumferentially spaced about the compressor shroud 58 to form a row of stator blades 56. The compressor shroud 58 also defines a compressor airflow passage 60 therein, in which the inlet air 20 flows and is compressed. Figure 1The compressor section 14 is merely one example of a compressor that may be implemented in a gas turbine engine, and the principles described herein are applicable to any type of compressor, including a purely axial flow compressor as will be described in greater detail below.

[0026] The downstream portion 52 of the compressor section 14 is configured for centrifugal or mixed axial-centrifugal fluid flow and may be referred to as a centrifugal compressor airflow passage 53 or simply a centrifugal portion. The downstream portion 52 includes an impeller 42 mounted to rotate with the shaft 44. An annular array of airfoil impeller blades 62 extends outward from the impeller 42. The impeller blades 62 are configured in terms of their size, cross-sectional shape, orientation, spacing, and other parameters (according to conventional practice) to provide an incremental pressure increase to the compressed air 24 flowing through them as the impeller 42 rotates. The compressor shroud 58 extends radially outward and surrounds the impeller 42. A compressor bleed air system 63, which will be described in more detail below, is disposed at the downstream portion 52 of the compressor section 14. The compressor bleed air system 63 may include a compressor bleed air plenum 64 that may be defined within a tapered wall 66, a compressor casing 68, and the compressor shroud 58.

[0027] The compressor bleed air system 63 also includes a plurality of compressor bleed air passages 70 extending through the compressor shroud 58 and providing fluid communication between the compressor airflow passage 60 and the compressor bleed air chamber 64. A portion of the compressed air 24 flowing through the compressor airflow passage 60 can enter the compressor bleed air chamber 64 as a compressor bleed air flow 72 through the compressor bleed air passage 70. The compressor bleed air passage 70 can be positioned at a distance 80 upstream of the outlet end 82 of the impeller 42. The distance 80 along the compressor airflow passage 60 can vary according to the specific application based on the pressure level requirement of the compressor bleed air chamber 64. These requirements are established based on the understanding of the pressure level requirements of the aircraft system 76 and the losses in the discharge pipe 74. The compressor bleed air passage 70 allows a portion of the air velocity head to be recovered from the inlet air 20 entering the compressor section 14. The distance 80 is selected so that the discharge pressure of the compressor bleed air passage 70 is at least the static pressure level required for the compressor bleed air chamber 64.

[0028] The compressor bleed air system 63 may also include one or more discharge pipes 74, which provide compressed bleed air flow 72 from the compressor bleed air chamber 64 to an engine or aircraft system 76, schematically shown as a box, which can be used for various purposes, such as cooling, driving or environmental control. The flow can be controlled or regulated by appropriate equipment (e.g., valve 78, schematically shown).

[0029] Figure 2 According to one aspect of the present disclosure, Figure 1 Detailed view 84 of FIG. 8 is an enlarged view of compressor bleed air passage 70 . Figure 2One compressor bleed air passage 70 is shown, however, in the compressor section 14, multiple compressor bleed air passages 70 are circumferentially spaced apart about the longitudinal centerline axis 12 by the compressor shroud 58. The circumferential spacing of the multiple compressor bleed air passages 70 will be discussed in more detail below, but each of the multiple compressor bleed air passages 70 may be spaced apart as shown. Figure 2 Configure as shown.

[0030] exist Figure 2 , it can be seen that the compressor bleed air passage 70 extends from the inlet 86 at the compressor air flow passage 60 through the compressor shroud 58 to the outlet 88 at the compressor bleed air chamber 64. The compressor bleed air passage 70 provides fluid communication between the compressor air flow passage 60 and the compressor bleed air chamber 64. In the longitudinal direction L, the compressor bleed air passage 70 may be defined between the front side wall 90 and the rear side wall 92. The cross-sectional shape of the compressor bleed air passage 70 (e.g., intercepted through the plane 106) may be various shapes, such as circular, elliptical or polygonal. Therefore, the front side wall 90 and the rear side wall 92 may include a plurality of spaced-apart surfaces, or they may include a single continuous wall surface extending around a closed perimeter. The centerline 94 extends from the inlet center reference point 99 at the inlet 86 to the outlet center reference point 101 at the outlet 88. For any given axial position along the centerline 94 extending from the inlet 86 to the outlet 88, the centerline 94 passes through the centroid of the flow area of ​​the compressor bleed air passage 70 at the axial position.

[0031] At the inlet 86, the centerline 94 is disposed at a non-parallel, non-perpendicular angle (generally designated as "θ") to the inner surface 96 of the compressor shroud 58. For reference purposes, the angle θ is located in an axial-radial plane (i.e., in a plane defined by the longitudinal direction L and the radial direction R) and is measured between a reference line 98 tangent to the inner surface 96 at the inlet 86 of the compressor shroud 58 and the centerline 94. For reference purposes, the axial and circumferential positions of the inlet 86 may be defined by an inlet center reference point 99 defined at the intersection of the centerline 94 and the reference line 98. The angle θ1 at the inlet 86 may be referred to as the "discharge angle", while the angle θ2 may be referred to as the "exit angle". The value of the angle θ1 at the inlet 86 may be selected in accordance with conventional practice so as to extract or discharge air from the compressor airflow passage 60 at a desired flow rate and temperature conditions while minimizing pressure losses. The value of θ2 may be selected based on a desired flow rate into the compressor bleed air plenum 64.

[0032] The compressor bleed air passage 70 includes a throat 100 at an upstream end 102 of the compressor bleed air passage 70, the throat 100 being disposed proximate to and in communication with the inlet 86. The throat 100 may extend a distance 108 from the intersection of the reference line 98 and the centerline 94 to a plane 106 that is perpendicular to the centerline 94. Within the throat 100, the front sidewall 90 and the rear sidewall 92 are parallel to each other such that the flow area between the front sidewall 90 and the rear sidewall 92 along the length of the throat 100 is constant (i.e., the front sidewall 90 and the rear sidewall 92 do not diverge).

[0033] The compressor bleed air passage 70 includes a diffuser flow passage 104 located downstream of the throat 100, and the diffuser flow passage 104 extends from a plane 106 to an outlet 88 along a centerline 94. In the diffuser flow passage 104, the front sidewall 90 and the rear sidewall 92 are separated from each other as they extend from the plane 106 toward the outlet 88 in a downstream direction. Therefore, the diffuser flow passage 104 has an increased flow area from the plane 106 to the outlet 88 between the front sidewall 90 and the rear sidewall 92. The increase in flow area results in the flow diffusion of the compressor bleed air flow 72 through the compressor bleed air passage 70. As the compressor bleed air flow 72 passes through the diffuser flow passage 104 of the compressor bleed air passage 70 downstream, the diffusion reduces the speed of the compressor bleed air flow 72 through the compressor bleed air passage 70 and increases the static pressure. The diffusion rate is related to the diffusion angle α between the front sidewall 90 and the rear sidewall 92. For any given cross-sectional cutting plane 110 taken perpendicular to the centerline 94 , the divergence angle α is the angle between lines tangent to the front and rear sidewalls 90 , 92 and intersecting the cutting plane 110 .

[0034] In order to provide an appropriate diffusion rate while minimizing pressure losses, the diffusion angle α can vary along the length of the compressor bleed air passage 70 (e.g., from α1 to α2). In other words, the front sidewall 90 and the rear sidewall 92 can have a variable taper along their respective lengths instead of having a single straight taper. In the case of using a variable taper, the diffuser flow passage 104 may include a plurality of discrete longitudinal segments (not shown), each longitudinal segment having a constant diffusion angle α within the segment, or the diffusion angle α can be continuously varied within the segment. The diffusion angle α can be selected for a specific application using appropriate design techniques. For example, conventional computational fluid dynamics ("CFD") software or analysis methods can be used to model the airflow through the compressor bleed air passage 70 in order to determine the appropriate diffusion angle α. In terms of reducing pressure losses, the diffusion angle α can have a relatively large value (α1) upstream of the diffuser flow passage 104 or closer to the inlet 86, and can have a relatively low value (α2) downstream of the diffuser flow passage 104 near the outlet 88. In other words, the divergence angle α may have its maximum value immediately downstream of the throat 100 and may gradually decrease to a lower value toward the outlet 88 .

[0035] The flow path of the compressor bleed air passage 70 defined by the centerline 94 can but need not follow a straight line (e.g., a straight line in the radial direction R). On the contrary, all or part of the centerline 94 (and accordingly the front sidewall 90 and the rear sidewall 92) can be curved. The curvature can be used to reduce or eliminate the flow separation that may occur if the centerline 94 is straight instead of a curve. For example, in a radial-axial plane (i.e., in a plane defined by the longitudinal direction L and the radial direction R), the centerline 94 is disposed adjacent to the inlet 86 at a first angle θ1, and adjacent to the outlet 88 and the reference line 110 at a second angle θ2 that is different from the first angle θ1. The second angle θ2 can be referred to as an "injection angle". The radius of curvature and direction of the centerline 94 can be selected using appropriate design techniques for a particular application. For example, CFD software or analysis techniques can be used to model the airflow through the compressor bleed air passage 70 to determine the appropriate radius of curvature and direction of the centerline 94.

[0036] Figure 3 According to one aspect of the present disclosure, Figure 1 A partial cross-sectional view of the circumferentially arranged compressor bleed air passage 70 through the compressor shroud 58 taken at plane 3-3 of FIG. Figure 1 A cross-sectional view of the upper half of the gas turbine engine 10 is depicted (the upper half of the engine is above a horizontal reference plane 112 passing through the longitudinal centerline axis 12), but Figure 3 The lower half of the compressor shroud 58 (eg, the lower half of the engine below a horizontal reference plane 112 passing through the longitudinal centerline axis 12) and the impeller 42 ( Figure 1 ), and an example of a compressor bleed air passage 70 passing through a lower half of the compressor shroud 58. As described above, the compressor shroud 58 extends circumferentially about the longitudinal centerline axis 12, and Figure 3 A full circumferential depiction of the compressor shroud 58 is included.

[0037] like Figure 3 As shown, the compressor shroud 58 includes a plurality of compressor bleed air passages 70 therethrough, which provide fluid communication between the compressor airflow passages 60 and the compressor bleed air plenum 64 such that the flow of the compressor bleed air flow 72 flows from the outer edges of the impeller blades 62 within the compressor airflow passages 60 to the compressor bleed air plenum 64. Figure 3, sixteen compressor bleed air passages 70 are shown spaced circumferentially about the longitudinal centerline axis 12. However, more than sixteen or fewer than sixteen compressor bleed air passages 70 may be included instead. The sixteen compressor bleed air passages 70 include a first compressor bleed air passage 114, a second compressor bleed air passage 116, a third compressor bleed air passage 118, a fourth compressor bleed air passage 120, a fifth compressor bleed air passage 122, a sixth compressor bleed air passage 124, a seventh compressor bleed air passage 126, an eighth compressor bleed air passage 128, a ninth compressor bleed air passage 130, a tenth compressor bleed air passage 132, an eleventh compressor bleed air passage 134, a twelfth compressor bleed air passage 136, a thirteenth compressor bleed air passage 138, a fourteenth compressor bleed air passage 140, a fifteenth compressor bleed air passage 142, and a sixteenth compressor bleed air passage 144. Each of the first compressor bleed air passage 114 to the sixteenth compressor bleed air passage 144 may have a Figure 2 As shown above Figure 2 As discussed above, the location of the inlet 86 of each compressor bleed air passage 70 is defined by an inlet center reference point 99 located at the intersection of centerline 94 and reference line 98 .

[0038] like Figure 3 As shown, each of the plurality of compressor bleed air passages 70 is circumferentially spaced apart from each other based on the angular spacing between the corresponding inlet center reference points 99 of each of the plurality of compressor bleed air passages 70. Therefore, the first inlet center reference point 115 defines the circumferential position of the first compressor bleed air passage 114. Similarly, the second inlet center reference point 117 defines the circumferential position of the second compressor bleed air passage 116. The circumferential spacing between the first compressor bleed air passage 114 and the second compressor bleed air passage 116 is defined by the first angular spacing 146. Similarly, the third inlet center reference point 119 defines the circumferential position of the third compressor bleed air passage 118, and the third compressor bleed air passage 118 is angularly spaced apart from the second compressor bleed air passage 116 by the first angular spacing 146. Similarly, the angular spacing between the third compressor bleed air passage 118 and the fourth compressor bleed air passage 120 may be the first angular spacing 146, and the angular spacing between the fourth compressor bleed air passage 120 and the fifth compressor bleed air passage 122 may be the first angular spacing 146. The angular spacing between fifth compressor bleed air passage 122 and sixth compressor bleed air passage 124 may be first angular spacing 146 , the angular spacing between sixth compressor bleed air passage 124 and seventh compressor bleed air passage 126 may be first angular spacing 146 , and the angular spacing between seventh compressor bleed air passage 126 and eighth compressor bleed air passage 128 may be first angular spacing 146 .

[0039] The same first angular spacing 146 may also be implemented between the ninth compressor bleed air passage 130 and the tenth compressor bleed air passage 132, between the tenth compressor bleed air passage 132 and the eleventh compressor bleed air passage 134, between the eleventh compressor bleed air passage 134 and the twelfth compressor bleed air passage 136, between the twelfth compressor bleed air passage 136 and the thirteenth compressor bleed air passage 138, and between the thirteenth compressor bleed air passage 138 and the fourteenth compressor bleed air passage 140. Likewise, the same first angular spacing 146 may be implemented between the fourteenth compressor bleed air passage 140 and the fifteenth compressor bleed air passage 142, and between the fifteenth compressor bleed air passage 142 and the sixteenth compressor bleed air passage 144.

[0040] On the other hand, a second angular spacing 148 is implemented between the first compressor bleed air passage 114 and the sixteenth compressor bleed air passage 144. The second angular spacing 148 can be, for example, twice the first angular spacing 146. For example, when the first angular spacing 146 is twenty degrees, the second angular spacing is forty degrees. The same second angular spacing 148 can also be implemented between the eighth compressor bleed air passage 128 and the ninth compressor bleed air passage 130. Of course, the second angular spacing does not need to be twice the first angular spacing 146, and any other ratio between the first angular spacing 146 and the second angular spacing 148 can be implemented as long as the ratio is greater than 1:1. That is, the second angular spacing 148 and the first angular spacing 146 are different from each other, and preferably, the second angular spacing 148 is greater than the first angular spacing 146.

[0041] In addition, although Figure 3 A circumferential clocking arrangement is depicted, wherein the first compressor bleed air passage 114 is arranged along a vertical reference plane 150 and the remaining compressor bleed air passages are arranged circumferentially relative thereto, but the circumferential clocking of the compressor bleed air passages 70 need not be as Figure 3 As shown. Conversely, for example, the first compressor bleed air passage 114 may alternatively be circumferentially offset to the position of the second compressor bleed air passage 116, with each of the second compressor bleed air passage 116 through the sixteenth compressor bleed air passage 144 being circumferentially offset clockwise by the same angular amount. With the above-described circumferential spacing of the compressor bleed air passages 70, by providing a different (larger) second angular spacing 148 between at least two consecutively spaced compressor bleed air passages 70, the uneven spacing of the bleed air passages 70 facilitates changing the frequency of the air flow pulses relative to the compressor rotor blades to reduce the vibration response induced in the compressor rotor blades.

[0042] exist Figure 3In the embodiment of the present invention, the plurality of compressor bleed air passages 70 may also be considered to be grouped into a plurality of groups. For example, the first group 145 compressor bleed air passages 70 may include the first compressor bleed air passage 114 to the eighth compressor bleed air passage 128, and the second group 147 compressor bleed air passages 70 may include the ninth compressor bleed air passage 130 to the sixteenth compressor bleed air passage 144. In the first group 145 compressor bleed air passages 70, each compressor bleed air passage 70 is spaced apart by equal circumferential spacing, such as the first angular spacing 146. Similarly, in the second group 147 compressor bleed air passages 70, each compressor bleed air passage 70 is spaced apart by equal circumferential spacing, such as the first angular spacing 146. Therefore, the first angular spacing 146 may be implemented in the first group 145 and in the second group 147. Alternatively, in the second group 147, the equal circumferential spacing of each compressor bleed air passage 70 may be different from the equal circumferential spacing in the first group 145. For example, each compressor bleed air passage 70 within the second group 147 may be equally spaced apart at a first angular spacing 146 ′ that is different than the first angular spacing 146 within the first group 145 .

[0043] On the other hand, the first group 145 and the second group 147 are spaced apart from each other at an angular spacing that is greater than the first angular spacing 146. For example, the first group 145 and the second group 147 may be spaced apart by a second angular spacing 149, which may be the same as the second angular spacing 148. That is, for example, in the first group 145, the first compressor bleed air passage 114 may be considered to be located at a first end 155 of the first group 145, and the eighth compressor bleed air passage 128 may be considered to be located at a second end 157 of the first group 145. The ninth compressor bleed air passage 130 may be considered to be located at a first end 159 of the second group 147, and the sixteenth compressor bleed air passage 144 may be located at a second end 161 of the second group 147. The second angular spacing 149 between the first group 145 and the second group 147 is, for example, the angular spacing between the first end 155 of the first group 145 and the second end 161 of the second group 147 (i.e., the same as the second angular spacing 148 between the sixteenth compressor bleed air passage 144 and the first compressor bleed air passage 114), and the angular spacing between the second end 157 of the first group 145 and the first end 159 of the second group 147 (i.e., the same as the second angular spacing 148 between the eighth compressor bleed air passage 128 and the ninth compressor bleed air passage 130). In this case, the first group 145 and the second group 147 are spaced apart from each other. Figure 3 Considered as mirror images of each other in the cross-sectional view of , the compressor bleed air passages 70 within each group have equal first angular spacing 146 between them, and the first group 145 and the second group 147 have equal second angular spacing 149 between them.

[0044] Alternatively, instead of the adjacent ends of the first group 145 and the second group 147 being equally spaced apart by the second angular spacing 149, different angular spacings may be implemented between the adjacent ends of the first group 145 and the second group 147. For example, the second angular spacing 151 may be implemented between the first end 155 of the first group 145 and the second end 161 of the second group 147, wherein the second angular spacing 151 may be greater than the second angular spacing 149, and the third angular spacing 153 may be implemented between the second end 157 of the first group 145 and the first end 159 of the second group 147, wherein the third angular spacing 153 is less than the second circumferential spacing 151 and also less than the second angular spacing 149, but greater than the first angular spacing 146. In this case, the second group 147 may be spaced apart from the second angular spacing 151. Figure 3 The positions shown are offset counterclockwise around the longitudinal centerline axis 12 to provide a second angular spacing 151 that is larger than the second angular spacing 149, and to provide a second angular spacing 153 that is smaller than the second angular spacing 149. Utilizing the first angular spacing 146 of the compressor bleed air passages 70 in each group, and the second angular spacing 148, the second angular spacing 149, the second angular spacing 151, or the third angular spacing 153 between each group (where each angular spacing is greater than the first angular spacing 146), the uneven spacing of the bleed air passages 70 is conducive to changing the frequency of the air flow pulses relative to the compressor rotor blades to reduce the vibration response induced in the compressor rotor blades. Of course, the present disclosure is not limited to two groups (i.e., the first group 145 and the second group 147) of compressor bleed air passages 70, but more than two groups of compressor bleed air passages 70 can be implemented, as long as the angular spacing between adjacent groups is different from the angular spacing between the corresponding compressor bleed air passages 70 in the corresponding groups.

[0045] Figure 4 According to one aspect of the present disclosure, Figure 3 Detailed view 152 of FIG. 1 is an enlarged view of the compressor bleed air passage 70. Figure 4 In the present invention, the compressor bleed air passage 70' passing through the compressor shroud 58 is assumed to be a straight passage in the radial direction R (ie, in Figure 2 The radial-longitudinal RL plane extends in a radial direction R perpendicular to the longitudinal direction L), rather than Figure 2 The curved compressor bleed air passage 70 is shown in FIG. Figure 4 A compressor bleed air passage 70' is depicted having a throat 100' extending in a radial direction R and angled in a circumferential direction C relative to the longitudinal centerline axis 12. Figure 4 In FIG. 1 , the centerline 94′ is shown as being at an angle β relative to the radial direction R in the circumferential direction C. Alternatively, a radial direction β may be used that is perpendicular to the centerline 94 ( Figure 2 ) is similarly curved with a curved center line 94". Figure 4In the example shown, the centerline 94' extends from the inlet center reference point 99' to the outlet center reference point 101', and is arranged at a first angle β1 (i.e., the tangential component of the above-mentioned discharge angle) near the inlet 86', and is arranged at a second angle β2 (i.e., the tangential component of the above-mentioned injection angle) different from the first angle β1 at the outlet 88'. Therefore, both the discharge angle and the injection angle can be three-dimensional parameters. These angles and curvatures can be configured to provide a low composite discharge angle at the intersection with the outlet of the compressor airflow passage 60, and then approach a functionally optimal injection angle in three dimensions, which results in a low swirl discharge flow or a zero swirl discharge flow.

[0046] Figure 5 Described is a Figure 2 An alternative arrangement of compressor bleed air passages is shown. Figure 5 In the aspect of the invention, in the side view depicted therein, an annular groove compressor bleed air passage 154 is depicted and is configured to Figure 2 The compressor bleed air passage 70 is substantially the same as the compressor bleed air passage 70 of FIG. 1 , with one difference being that the annular groove compressor bleed air passage 154 is configured as an annular groove extending partially or entirely around the central longitudinal axis 12 . The annular groove compressor bleed air passage 154 is defined between a front sidewall 156 of the compressor shroud 58 and a rear sidewall 158 of the compressor shroud 58 . Similar to FIG. Figure 2 In the compressor bleed air passage 70 , the annular groove compressor bleed air passage 154 extends from an inlet 160 at the compressor air flow passage 60 to an outlet 162 at the compressor bleed air chamber 64 . Figure 5 The other elements and components for the annular groove compressor bleed passage 154 are depicted in Figure 2 The same elements are depicted in the aspects, but include an "a" designation. These same elements include centerline 94a, reference plane 98a, inlet center reference point 99a, outlet center 101a, throat 100a, upstream end 102a, diffuser flow passage 104a, plane 106a, distance 108a, and cutting plane 110a. Figure 5 The same components with an "a" mark are described in Figure 2 The descriptions of the corresponding elements without the “a” mark are the same, and the descriptions of these same elements will not be repeated here.

[0047] Figure 6 According to one aspect of the present disclosure Figure 3 Alternative cross-sectional view depicting Figure 5 An alternative annular groove compressor bleed air passage 154. Figure 7 It is captured at detail view 170 Figure 6 An enlarged view of a portion of. Common Reference Figure 6 and Figure 7, the annular groove compressor bleed air passage 154 may be circumferentially divided into a plurality of compressor bleed air passages 166 by including a plurality of vanes 164 within the annular groove compressor bleed air passage 154. As will be described below, each of the plurality of vanes 164 may have the same shape and may be spaced apart from each other by the same amount, such that each of the plurality of compressor bleed air passages 166 also has the same shape.

[0048] Figure 6 The configuration also includes at least one large vane 168, which extends in the circumferential direction across a larger circumferential span than each vane 164. Each of the vanes 164 and the at least one large vane 168 extends in the axial direction L ( Figure 5 ) extends between the front sidewall 156 and the rear sidewall 158. A plurality of compressor bleed air passages 166 are circumferentially spaced apart from each other by a first angular spacing 172. For example, the first compressor bleed air passage 176 and the second compressor bleed air passage 178 are spaced apart from each other by a first angular spacing 172. The first angular spacing 172 is due to the inclusion of vanes 164 disposed between the first compressor bleed air passage 176 and the second compressor bleed air passage 178. Similarly, the second compressor bleed air passage 178 and the third compressor bleed air passage 180 are circumferentially spaced apart from each other by a first angular spacing 172. Again, the first angular spacing 172 is due to the inclusion of vanes 164 disposed between the second compressor bleed air passage 178 and the third compressor bleed air passage 180. Therefore, in the case where the vanes 164 are arranged to define the compressor bleed air passage 166, the spacing between those adjacent compressor bleed air passages 166 is the first angular spacing 172.

[0049] On the other hand, fourth compressor bleed air passage 182 and first compressor bleed air passage 176 are circumferentially spaced apart from each other by a second angular spacing 174 that is greater than first angular spacing 172. Second angular spacing 174 is due to the inclusion of large vanes 168 disposed between first compressor bleed air passage 176 and fourth compressor bleed air passage 182.

[0050] like Figure 7 As shown, the vanes 164 may have a shape defined by a perimeter having a first side 183, a second side 184, an inner end 186, and an outer end 188. The first side 183 and the second side 184 may have a curved shape. The shape of the compressor bleed air passage 166 is defined by the shapes of adjacent vanes 164. For example, the shape of the second compressor bleed air passage 178 disposed between the first vane 190 and the second vane 192 is defined by the shapes of the second side 184 of the first vane 190 and the first side 183 of the second vane 192. In the case where the shape of each vane 164 is the same and the first angular spacing 172 between the vanes 164 is the same, the shape of the compressor bleed air passage 166 is also the same. Figure 3 In the same way, Figure 7The circumferential spacing of the compressor bleed air passages 166 in is taken at the inlet center reference point 99a and the longitudinal centerline axis 12.

[0051] Figures 1 to 7 The foregoing description relates to Figure 1 A centrifugal flow compressor section 14 is shown, but the present disclosure with respect to the compressor bleed air passages may also be implemented in an axial flow compressor arrangement of a gas turbine engine. Figure 8 2 is a schematic partial cross-sectional view of an exemplary high bypass gas turbine engine 210 (also referred to herein as "engine 210") according to an aspect of the present disclosure. Although further described below with reference to a high bypass turbine engine, the present disclosure is also applicable to turbomachinery in general, including turbojets, turbopropellers, and turboshaft gas turbine engines, including marine turbine engines, industrial turbine engines, and auxiliary power units. Figure 8 As shown, for reference purposes, engine 210 has a longitudinal centerline axis 212 extending through engine 210 from upstream end 198 to downstream end 199. Generally speaking, engine 210 may include a fan assembly 214 and a core engine 216 disposed downstream of fan assembly 214.

[0052] The core engine 216 may generally include an outer casing 218 defining an annular inlet 220. The outer casing 218 surrounds or at least partially forms a compressor section 221 having a low pressure (LP) compressor 222 and a high pressure (HP) compressor 224, a diffuser 225, a combustion section 226, a turbine section 227 including a high pressure (HP) turbine 228 and a low pressure (LP) turbine 230, and an ejection exhaust nozzle section 232 in a series flow relationship. A high pressure (HP) rotor shaft 234 drivingly connects the HP turbine 228 to the HP compressor 224. A low pressure (LP) rotor shaft 236 drivingly connects the LP turbine 230 to the LP compressor 222. The LP rotor shaft 236 may also be connected to a fan shaft 238 of the fan assembly 214. In certain embodiments, as shown in FIG. Figure 8 As shown, the LP rotor shaft 236 may be connected to the fan shaft 238 via a reduction gearbox 240 , such as in an indirect drive configuration or a gear drive configuration.

[0053] like Figure 8As shown, fan assembly 214 includes a plurality of fan blades 242 that are coupled to and extend radially outward from fan shaft 238. An annular fan casing or nacelle 244 circumferentially surrounds fan assembly 214 and / or at least a portion of core engine 216. Nacelle 244 may be supported relative to core engine 216 by a plurality of circumferentially spaced outlet guide vanes or struts 246. In addition, at least a portion of nacelle 244 may extend over an exterior of core engine 216 to define a bypass airflow passage 248 therebetween.

[0054] During operation of the engine 210, a volume of air 273, as schematically indicated by arrows, enters the engine 210 from the upstream end 198 through the nacelle 244 and / or an associated nacelle inlet 276 of the fan assembly 214. As the air 273 passes through the fan blades 242, a portion of the air 273 is directed or directed into the bypass airflow passage 248 as a bypass airflow 278, while another portion of the air 273 is directed or directed into the LP compressor 222 via the annular inlet 220 as compressor inlet air 280. As the compressor inlet air 280 flows from the annular inlet 220 through the LP compressor 222 and the HP compressor 224 toward the combustion section 226, the compressor inlet air 280 is progressively compressed to produce compressed air 282. The compressed air 282 flows through the diffuser 225 and into the combustion section 226, where the compressed air 282 is mixed with fuel to produce an air / fuel mixture that is ignited and combusted to produce combustion gases 286. Combustion gases 286 flow into HP turbine 228, thereby causing HP rotor shaft 234 to rotate, thereby supporting operation of HP compressor 224. Combustion gases 286 are then directed through LP turbine 230, thereby causing LP rotor shaft 236 to rotate, thereby supporting operation of LP compressor 222 and / or rotation of fan shaft 238. Combustion gases 286 are then exhausted through jet exhaust nozzle section 232 of core engine 216 to provide propulsion at downstream end 199.

[0055] Fig. 9 According to one aspect of the present disclosure, Figure 8 A partial cross-sectional view of an exemplary portion of a high pressure compressor 224 is shown taken at the detailed view 200 of FIG. Fig. 9In the aspect of , the high-pressure compressor 224 is an axial flow compressor in which the compressed air 282 flows in a generally axial direction (in a generally longitudinal direction L relative to the longitudinal centerline axis 212). The high-pressure compressor 224 includes a compressor shroud 250 and a compressor rotor shaft assembly 252 that extend circumferentially around the longitudinal centerline axis 212 (which may also therefore correspond to the centerline axis of the compressor shroud 250). An axial compressor airflow passage 251 is defined between the compressor shroud 250 and the compressor rotor shaft assembly 252. The compressor rotor shaft assembly 252 may include, for example, a plurality of compressor rotors connected to the HP shaft 234, including a first compressor rotor 254, a second compressor rotor 256, a third compressor rotor 258, and a fourth compressor rotor 260. Of course, the compressor rotor shaft assembly 252 is not limited to four compressor rotors, but may include less than four compressor rotors or more than four compressor rotors. Each compressor rotor includes a plurality of circumferentially spaced rotor blades. For example, the first compressor rotor 254 includes a plurality of circumferentially spaced rotor blades 262 ( Fig. 9 ), the second compressor rotor 256 includes a plurality of circumferentially spaced rotor blades 264, and the third compressor rotor 258 includes a plurality of circumferentially spaced rotor blades 266 ( Fig. 9 ), and the fourth compressor rotor 260 includes a plurality of circumferentially spaced rotor blades 268 ( Fig. 9 ). The first compressor rotor 254 may be referred to as a first rotor stage 270, the second compressor rotor 256 may be referred to as a second rotor stage 272, the third compressor rotor 258 may be referred to as a third rotor stage 274, and the fourth compressor rotor 260 may be referred to as a fourth rotor stage 284. As the compressed air 282 flows axially downstream through the compressor air flow passage 251, each of the first rotor stage 270, the second rotor stage 272, the third rotor stage 274, and the fourth rotor stage 284 gradually compress the compressed air 282.

[0056] The high pressure compressor 224 also includes a plurality of stages of stator blades connected to the compressor shroud 250 and extending into the compressor air flow passage 251. The plurality of stages of stator blades include a first stator blade stage 288 including a plurality of stator blades 290 ( Fig. 9 ), a second stator blade stage 292 including a plurality of stator blades 294 ( Fig. 9 ), and a third stator blade stage 296 including a plurality of stator blades 298 ( Fig. 9). The stator vanes of each of first stator vane stage 288, second stator vane stage 292, and third stator vane stage 296 are arranged to redirect the flow of compressed air 282 exiting the compressor rotor stage immediately upstream of the respective stator vane stage.

[0057] Although Fig. 9 Four compressor rotor stages and three stator vane stages are depicted, but the present disclosure is not limited to four compressor rotor stages and three stator vane stages. Rather, any number of additional compressor rotor stages and corresponding stator vane stages may be arranged upstream of the first rotor stage 270 and / or downstream of the fourth rotor stage 284. For example, a plurality of compressor rotor stages and corresponding stator vane stages may be arranged upstream of the first rotor stage 270 as shown in FIG. Fig. 9 Three additional stator blade stages and three additional rotor blade stages are arranged alternately in a similar manner as shown, or six additional stator blade stages and six additional rotor blade stages may be provided upstream of the first rotor stage 270 .

[0058] Compressor section 221 includes a compressor bleed air system 299 at high pressure compressor 224. Compressor bleed air system 229 includes a compressor shroud 250, which may include an inner wall 253 and an outer wall 255, which define a compressor bleed air chamber 300 therebetween. Inner wall 253, outer wall 255, and compressor bleed air chamber 300 may extend circumferentially about longitudinal centerline axis 212. Compressor bleed air system 229 also includes at least one compressor bleed air passage 302 ( Fig. 9 ), which extends through the compressor shroud 250 and provides fluid communication between the compressor airflow passage 251 and the compressor bleed air chamber 300. The compressor bleed air passage 302 may have a Figure 2 The compressor bleed air passage 70 shown in FIG. 302 may be of a similar form (ie, the compressor bleed air passage 302 may be such as Figure 2 However, the compressor bleed air passage 302 is not limited to Figure 3 The form shown, but can be provided in any other form. Fig. 9 In the aspect of FIG. 3 , the compressor bleed air passage 302 is shown as being arranged through the compressor shroud on the downstream side 305 of the first stator blade stage 288. Therefore, when the compressor bleed air passage 302 has Figure 2 In the example of the compressor bleed air passage 70, the inlet 86 ( Figure 2) is arranged downstream of the downstream side 305 of the first stator blade stage 288. In operation, a portion of the compressed air 282 flowing through the compressor air flow passage 251 is directed through at least one compressor bleed air passage 302 to enter the compressor bleed air chamber 300 as a compressor bleed air flow 283. The compressor bleed air discharge line 303 can provide the compressor bleed air flow 283 from the compressor bleed air chamber 300 to other parts of the engine 210 or to other aircraft systems (not shown).

[0059] Fig.10 According to one aspect of the present disclosure, Fig. 9 FIG. 1 is a partial cross-sectional view of the circumferential arrangement of compressor bleed air passages 302 through compressor shroud 250 taken at plane 10-10 of FIG. Fig. 9 A cross-sectional view of the upper half of the high pressure compressor 224 is depicted (ie, the upper half above a horizontal reference plane 212 ( a ) passing through the longitudinal centerline axis 212 ), but Fig.10 An example of a compressor bleed air passage 302 including a lower portion of the compressor shroud 250 (i.e., a lower portion of the high pressure compressor 224 below a horizontal reference plane 212(a) passing through the longitudinal centerline axis 212) and passing through the lower portion of the compressor shroud 250. As described above, the compressor shroud 250 extends circumferentially around the longitudinal centerline axis 212, and Fig.10 A full circumferential depiction of the compressor shroud 250 is included. Fig.10 As shown, the compressor shroud 250 includes a plurality of compressor bleed air passages 302 therethrough, which provide fluid communication between the compressor airflow passage 251 and the compressor bleed air chamber 300, so that the flow of the compressor bleed air flow 283 flows from the compressor airflow passage 251 to the compressor bleed air chamber 300. Fig.10 In, similar to Figure 3 In the aspect of the present invention, sixteen compressor bleed air passages 302 are shown, which are circumferentially spaced about the longitudinal centerline axis 212. However, more than sixteen compressor bleed air passages 302 or less than sixteen compressor bleed air passages 302 may alternatively be included. The sixteen compressor bleed air passages 302 include a first compressor bleed air passage 304, a second compressor bleed air passage 306, a third compressor bleed air passage 308, a fourth compressor bleed air passage 310, a fifth compressor bleed air passage 312, a sixth compressor bleed air passage 314, a seventh compressor bleed air passage 316, an eighth compressor bleed air passage 318, a ninth compressor bleed air passage 320, a tenth compressor bleed air passage 322, an eleventh compressor bleed air passage 324, a twelfth compressor bleed air passage 326, a thirteenth compressor bleed air passage 328, a fourteenth compressor bleed air passage 330, a fifteenth compressor bleed air passage 332, and a sixteenth compressor bleed air passage 334.

[0060] Each of the first compressor bleed air passage 304 to the sixteenth compressor bleed air passage 334 has a Figure 2 With the configuration shown, the location of the inlet 86 of each compressor bleed air passage 302 is defined by an inlet center reference point 99 located at the intersection of centerline 94 and reference line 98 .

[0061] like Fig.10 As shown, each of the plurality of compressor bleed air passages 302 is circumferentially spaced apart from one another based on the angular spacing between the respective inlet center reference points 99 of each of the plurality of compressor bleed air passages 302. Thus, the first inlet center reference point 340 of the first compressor bleed air passage 304 (corresponding to Figure 2 The inlet center reference point 99 shown in FIG. 1 defines the circumferential position of the first compressor bleed air passage 304. Similarly, the second inlet center reference point 342 defines the circumferential position of the second compressor bleed air passage 306. The circumferential spacing between the first compressor bleed air passage 304 and the second compressor bleed air passage 306 is defined by the first angular spacing 336. Similarly, the third inlet center reference point 344 defines the circumferential position of the third compressor bleed air passage 308, and the third compressor bleed air passage 308 is angularly spaced from the second compressor bleed air passage 306 by the first angular spacing 336. Similarly, the angular spacing between the third compressor bleed air passage 308 and the fourth compressor bleed air passage 310 may be the first angular spacing 336, and the angular spacing between the fourth compressor bleed air passage 310 and the fifth compressor bleed air passage 312 may be the first angular spacing 336. The angular spacing between the fifth compressor bleed air passage 312 and the sixth compressor bleed air passage 314 may be the first angular spacing 336, the angular spacing between the sixth compressor bleed air passage 314 and the seventh compressor bleed air passage 316 may be the first angular spacing 336, and the angular spacing between the seventh compressor bleed air passage 316 and the eighth compressor bleed air passage 318 may be the first angular spacing 336.

[0062] The same first angular spacing 336 may also be implemented between the ninth compressor induction air passage 320 and the tenth compressor induction air passage 322, between the tenth compressor induction air passage 322 and the eleventh compressor induction air passage 324, between the eleventh compressor induction air passage 324 and the twelfth compressor induction air passage 326, between the twelfth compressor induction air passage 326 and the thirteenth compressor induction air passage 328, and between the thirteenth compressor induction air passage 328 and the fourteenth compressor induction air passage 330. Likewise, the same first angular spacing 336 may be implemented between the fourteenth compressor induction air passage 330 and the fifteenth compressor induction air passage 332, and between the fifteenth compressor induction air passage 332 and the sixteenth compressor induction air passage 334.

[0063] On the other hand, the second angular spacing 338 is implemented between the first compressor bleed air passage 304 and the sixteenth compressor bleed air passage 334. The second angular spacing 338 can be, for example, twice the first angular spacing 336. For example, when the first angular spacing 336 is twenty degrees, the second angular spacing 338 is forty degrees. The same second angular spacing 338 can also be implemented between the eighth compressor bleed air passage 318 and the ninth compressor bleed air passage 320. Of course, the second angular spacing 338 does not need to be twice the first angular spacing 336, and any other ratio between the first angular spacing 336 and the second angular spacing 338 can be implemented as long as the ratio is greater than 1:1. That is, the second angular spacing 338 and the first angular spacing 336 are different from each other, and preferably, the second angular spacing 338 is greater than the first angular spacing 336.

[0064] In addition, although Fig.10 A circumferential clocking arrangement is depicted, wherein the first compressor bleed air passage 304 is arranged along a vertical reference plane 346 and the remaining compressor bleed air passages are arranged circumferentially relative thereto, but the circumferential clocking of the compressor bleed air passages 302 need not be as Fig.10 As shown. Instead, for example, the first compressor bleed air passage 304 may alternatively be circumferentially offset to the position of the second compressor bleed air passage 306, and each of the second compressor bleed air passage 308 to the sixteenth compressor bleed air passage 334 is offset clockwise by the same angular amount. With the aforementioned circumferential spacing of the compressor bleed air passages 302, by providing a different (larger) second angular spacing 338 between at least two consecutively spaced compressor bleed air passages 302, the uneven spacing of the bleed air passages 302 facilitates changing the frequency of the air flow pulses relative to the compressor rotor blades to reduce the vibration response induced in the compressor rotor blades.

[0065] exist Fig.10 In terms of Figure 3In the aspect of the invention, the plurality of compressor bleed air passages 302 may also be considered to be grouped into a plurality of groups. For example, the first group 335 compressor bleed air passages 302 may include the first compressor bleed air passage 304 to the eighth compressor bleed air passage 318, and the second group 337 compressor bleed air passages 302 may include the ninth compressor bleed air passage 320 to the sixteenth compressor bleed air passage 334. Within the first group 335 compressor bleed air passages 302, each compressor bleed air passage 302 is spaced apart by equal circumferential spacing, such as the first angular spacing 336. Similarly, within the second group 337 compressor bleed air passages 302, each compressor bleed air passage 302 is spaced apart by equal circumferential spacing, such as the first angular spacing 336. Therefore, the first angular spacing 336 may be implemented in the first group 335 and the second group 337. Alternatively, within the second group 337, the equal circumferential spacing of each compressor bleed air passage 302 may be different from the equal circumferential spacing in the first group 335. For example, each compressor bleed air passage 302 within second group 337 may be equally spaced apart at a first angular spacing 336 ′ that is different than first angular spacing 336 within first group 335 .

[0066] On the other hand, the first group 335 and the second group 337 are spaced apart from each other at an angular spacing that is greater than the first angular spacing 336. For example, the first group 335 and the second group 337 may be spaced apart by a second angular spacing 339, which may be the same as the second angular spacing 338. That is, for example, in the first group 335, the first compressor bleed air passage 304 may be considered to be located at the first end 341 of the first group 335, and the eighth compressor bleed air passage 318 may be considered to be located at the second end 343 of the first group 335. The ninth compressor bleed air passage 320 may be considered to be located at the first end 345 of the second group 337, and the sixteenth compressor bleed air passage 334 may be located at the second end 347 of the second group 337. The second angular spacing 339 between the first group 335 and the second group 337 is, for example, the angular spacing between the first end 341 of the first group 335 and the second end 347 of the second group 337 (i.e., the same as the second angular spacing 338 between the sixteenth compressor bleed air passage 334 and the first compressor bleed air passage 304), and the angular spacing between the second end 343 of the first group 335 and the first end 345 of the second group 337 (i.e., the same as the second angular spacing 338 between the eighth compressor bleed air passage 318 and the ninth compressor bleed air passage 320). In this case, the first group 335 and the second group 337 are spaced apart from each other. Fig.10 Considered as mirror images of one another in the cross-sectional view, the compressor bleed air passages 302 within each group have equal first angular spacing 336 between them, and equal second angular spacing 339 between the first group 335 and the second group 337 .

[0067] Alternatively, instead of the adjacent ends of the first group 335 and the second group 337 being equally spaced apart by the second angular spacing 339, different angular spacings may be implemented between the adjacent ends of the first group 335 and the second group 337. For example, the second angular spacing 349 may be implemented between the first end 341 of the first group 335 and the second end 347 of the second group 337, wherein the second angular spacing 349 may be greater than the second angular spacing 339, and the third angular spacing 351 may be implemented between the second end 343 of the first group 335 and the first end 345 of the second group 337, wherein the third angular spacing 351 is less than the second angular spacing 349 and also less than the second angular spacing 338, but greater than the first angular spacing 336. In this case, the second group 337 may be spaced apart from the adjacent ends of the first group 335 and the second group 337. Fig.10 The positions shown are offset counterclockwise around the longitudinal centerline axis 12 to provide a second angular spacing 349 that is larger than the second angular spacing 339, and to provide a third angular spacing 351 that is smaller than the second angular spacing 339. Utilizing the first angular spacing 336 of the compressor bleed air passages 302 in each group, and the second angular spacing 338, the second angular spacing 339, the second angular spacing 349, or the third angular spacing 351 between each group (where each angular spacing is greater than the first angular spacing 336), the uneven spacing of the bleed air passages 302 is conducive to changing the frequency of the air flow pulses relative to the compressor rotor blades to reduce the vibration response induced in the compressor rotor blades. Of course, the present disclosure is not limited to two groups (i.e., the first group 335 and the second group 337) of compressor bleed air passages 302, but more than two groups of compressor bleed air passages 302 can be implemented, as long as the angular spacing between adjacent groups is different from the angular spacing between the corresponding compressor bleed air passages 302 in the corresponding group.

[0068] Fig.11 According to another aspect of the present disclosure Fig. 9 A cross-sectional view of an alternative high pressure compressor is shown in FIG. Fig.11 In, with Fig. 9 Elements identical to those shown in FIG. 1 are labeled with the same reference numerals, and the description of these elements is the same as above for Fig. 9 Same as described. However, Fig. 9 Aspects and Fig.11 One difference between the aspects is that in Fig.11 In terms of the compressor bleed air passages 302', Fig. 9 The axial position of the compressor inlet air passage 302 is axially offset rearward. Fig. 9 In one aspect, the plurality of compressor bleed air passages 302 are axially located on the downstream side 305 of the first stator blade stage 288. On the other hand, in Fig.11 In the aspect of FIG. 2 , the plurality of compressor bleed air passages 302 ′ are arranged axially (ie, in the longitudinal direction L) with the second rotor stage 272 (eg, above the rotor blades of the compressor rotor). Fig.11 In the aspect of the invention, the circumferential clock and angular spacing between the plurality of compressor bleed air passages 302' is Fig.10 Targeted Fig. 9 Therefore, Fig.10 The description also applies to Fig.11 The circumferential spacing of the plurality of compressor bleed air passages 302' is shown in FIG.

[0069] Fig.12 According to another aspect of the present disclosure Fig. 9 A cross-sectional view of an alternative high pressure compressor is shown in FIG. Fig.12 In, with Fig. 9 Elements identical to those shown in FIG. 1 are labeled with the same reference numerals, and the description of these elements is the same as above for Fig. 9 Same as described. However, Fig. 9 Aspects and Fig.12 One difference between the aspects is that Fig.12 In terms of the compressor bleed air passages 302″, Fig. 9 The axial position of the compressor bleed air passage 302 is axially offset forward. Fig. 9 In one aspect, the plurality of compressor bleed air passages 302 are axially located on the downstream side 305 of the first stator blade stage 288. On the other hand, in Fig.12 In the aspect of FIG. 3 , a plurality of compressor bleed air passages 302 ″ are arranged through the base of the stator blades 290 or circumferentially between each stator blade 290 to provide a compressor bleed air flow 283 therethrough into the compressor bleed air chamber 300. Fig.12 In the aspect of the invention, the circumferential clock and angular spacing between the plurality of compressor bleed air passages 302" may be the same as Fig.10 Targeted Fig. 9 Therefore, Fig.10 The description also applies to Fig.12 The circumferential spacing of the plurality of compressor bleed air passages 302″ in the aspect of FIG.

[0070] The compressor bleed air system described herein implements non-uniform compressor bleed air passage spacing to reduce irritation on the compressor rotor upstream and downstream of the compressor bleed air passages. The use of non-uniform compressor bleed air passage spacing changes the forcing effect around the compressor rotor. This can be achieved through a variety of asymmetric patterns, including but not limited to split case spacing (right half / left half) differences, large gaps at specific clockwise locations, multiple sections of non-uniform spacing, or any other combination (especially for additively manufactured parts). This solution does not require redesigning the rotating part by adding weight or other performance-impairing features to make it more robust to irritation to meet life requirements.

[0071] Although the foregoing description is directed to some exemplary embodiments of the present disclosure, other variations and modifications will be apparent to those skilled in the art and may be made without departing from the spirit or scope of the present disclosure. In addition, the features described in conjunction with one embodiment of the present disclosure may be used in conjunction with other embodiments, even if not explicitly stated above.

[0072] Further aspects of the disclosure are provided by the subject matter of the following clauses.

[0073] A gas turbine engine comprises: a compressor; and a compressor bleed air system, the compressor bleed air system comprising: a compressor shroud, a compressor airflow passage defined in the compressor shroud; a compressor bleed air chamber; and a plurality of compressor bleed air passages, the plurality of compressor bleed air passages being arranged to pass through the compressor shroud and provide fluid communication between the compressor airflow passage and the compressor bleed air chamber, each of the plurality of compressor bleed air passages comprising an inlet at the compressor airflow passage and an outlet at the compressor bleed air chamber, wherein respective inlets of respective compressor bleed air passages in the plurality of compressor bleed air passages are arranged circumferentially spaced apart from each other around a centerline axis of the compressor shroud, a first compressor bleed air passage in the plurality of compressor bleed air passages and a second compressor bleed air passage in the plurality of compressor bleed air passages are arranged circumferentially adjacent to each other, and the inlet of the first compressor bleed air passage and the inlet of the second compressor bleed air passage are circumferentially spaced apart from each other by a first angular spacing, and respective inlets of at least a portion of the remaining portions of the plurality of compressor bleed air passages arranged circumferentially sequentially adjacent to each other are circumferentially spaced apart from each other by a second angular spacing, the second angular spacing being smaller than the first angular spacing.

[0074] A gas turbine engine according to the preceding clause, wherein at least one of said plurality of compressor bleed air passages defines a constant width throat at said inlet and defines a diverging portion from a downstream end of said throat to said outlet.

[0075] A gas turbine engine according to any of the preceding clauses, wherein the compressor air flow passage defines a centrifugal compressor air flow passage, and the plurality of compressor bleed air passages are arranged to receive bleed air flow from outer edges of impeller blades of an impeller rotating within the centrifugal compressor air flow passage.

[0076] A gas turbine engine according to any one of the preceding clauses, wherein the compressor air flow passage defines an axial compressor air flow passage, the compressor shroud includes a plurality of stator blades arranged in a stage circumferentially around the compressor shroud, and the plurality of compressor bleed air passages are arranged on a downstream side of the plurality of stator blades.

[0077] A gas turbine engine as described in any preceding clause, wherein respective inlets of said portion of said remainder of said plurality of compressor bleed air passages are circumferentially spaced apart from one another at equal angular intervals.

[0078] A gas turbine engine as described in any one of the preceding clauses, wherein respective inlets of said portion of said remaining portion of said plurality of compressor bleed air passages are arranged circumferentially between said first compressor bleed air passage and said second compressor bleed air passage.

[0079] A gas turbine engine according to any of the preceding clauses, wherein the inlet of a first portion of the portion of the remainder of the plurality of compressor bleed air passages is arranged adjacent to the first compressor bleed air passage and is arranged at the second angular spacing from the inlet of the first compressor bleed air passage, and the inlet of a second portion of the portion of the remainder of the plurality of compressor bleed air passages is arranged adjacent to the second compressor bleed air passage and is arranged at the second angular spacing from the inlet of the second compressor bleed air passage.

[0080] A gas turbine engine as in any one of the preceding clauses, wherein at least one of the plurality of compressor bleed air passages defines a diffuser flow passage, wherein the diffuser flow passage increases in size from the inlet to the outlet.

[0081] A gas turbine engine according to any preceding clause, wherein the inlet and the outlet are circular openings.

[0082] A gas turbine engine as described in any preceding clause, wherein the diffuser flow passage defines a centerline extending from a center of the inlet to a center of the outlet.

[0083] A gas turbine engine as described in any one of the preceding clauses, wherein the centerline is a curved line extending from the center of the inlet to the center of the outlet.

[0084] A gas turbine engine according to any of the preceding clauses, wherein the compressor shroud includes an annular slot compressor bleed air passage therethrough, and the plurality of blades are arranged within the annular slot compressor bleed air passage, the plurality of compressor bleed air passages being defined between respective ones of the plurality of blades.

[0085] A gas turbine engine as in any one of the preceding clauses, wherein each of said plurality of buckets extends axially through said annular slot compressor bleed air passage from a forward sidewall of said annular slot compressor bleed air passage to a rearward sidewall of said annular slot compressor bleed air passage.

[0086] A gas turbine engine comprises: a compressor; and a compressor bleed air system, the compressor bleed air system comprising: a compressor shroud, a compressor air flow passage defined in the compressor shroud; a compressor bleed air chamber; and a plurality of compressor bleed air passages, the plurality of compressor bleed air passages being arranged to pass through the compressor shroud and provide fluid communication between the compressor air flow passage and the compressor bleed air chamber, each of the plurality of compressor bleed air passages comprising an inlet at the compressor air flow passage and an outlet at the compressor bleed air chamber, wherein (a) the plurality of compressor bleed air passages are arranged in a first group of the plurality of compressor bleed air passages and a second group of the plurality of compressor bleed air passages, (b) the respective inlets of each of the plurality of compressor bleed air passages in the first group and the respective inlets of each of the plurality of compressor bleed air passages in the second group are equidistantly spaced from each other by a first angular spacing around a centerline axis of the compressor shroud, and (c) a first end of the first group and a second end of the second group are arranged to be spaced from each other by a second angular spacing greater than the first angular spacing.

[0087] A gas turbine engine according to the preceding clause, wherein the compressor air flow passage defines a centrifugal compressor air flow passage, and the plurality of compressor bleed air passages are arranged to receive bleed air flow from outer edges of impeller blades of an impeller rotating within the centrifugal compressor air flow passage.

[0088] A gas turbine engine according to any of the preceding clauses, wherein the compressor air flow passage defines an axial compressor air flow passage, the compressor shroud includes a plurality of compressor stator blades arranged in a stage circumferentially around the compressor shroud, and the plurality of compressor bleed air passages are arranged on a downstream side of the plurality of compressor stator blades.

[0089] A gas turbine engine as in any preceding clause, wherein at least one of said plurality of compressor bleed air passages defines a diffuser flow passage, wherein said diffuser flow passage increases in size as extending from said inlet to said outlet.

[0090] A gas turbine engine as in any of the preceding clauses, wherein the second end of the first group and the first end of the second group are arranged to be spaced apart from each other by a third angular spacing greater than the first angular spacing.

[0091] A gas turbine engine as described in any preceding clause, wherein said second angular spacing is the same as said third angular spacing.

[0092] A gas turbine engine as described in any preceding clause, wherein said second angular spacing is different from said third angular spacing.

Claims

1. A gas turbine engine, characterized in that: include: compressor; and A compressor bleed air system, the compressor bleed air system comprising: a compressor shroud, wherein a compressor airflow passage is defined within the compressor shroud; Compressor bleed air plenum; and a plurality of compressor bleed air passages arranged through the compressor shroud and providing fluid communication between the compressor airflow passage and the compressor bleed air plenum, each of the plurality of compressor bleed air passages comprising an inlet at the compressor airflow passage and an outlet at the compressor bleed air plenum, wherein corresponding inlets of corresponding compressor bleed air passages among the plurality of compressor bleed air passages are arranged circumferentially spaced apart from each other around a centerline axis of the compressor shroud, a first compressor bleed air passage among the plurality of compressor bleed air passages and a second compressor bleed air passage among the plurality of compressor bleed air passages are arranged circumferentially adjacent to each other, and an inlet of the first compressor bleed air passage and an inlet of the second compressor bleed air passage are circumferentially spaced apart from each other by a first angular spacing, and corresponding inlets of at least a portion of the remaining portions of the plurality of compressor bleed air passages that are arranged circumferentially sequentially adjacent to each other are circumferentially spaced apart from each other by a second angular spacing, and the second angular spacing is smaller than the first angular spacing.

2. The gas turbine engine according to claim 1, characterized in that Wherein at least one of the plurality of compressor bleed air passages defines a throat of constant width at the inlet and defines a diverging portion from a downstream end of the throat to the outlet.

3. The gas turbine engine according to claim 1, characterized in that Wherein the compressor air flow passage defines a centrifugal compressor air flow passage, and the plurality of compressor bleed air passages are arranged to receive bleed air flow from outer edges of impeller blades of an impeller rotating within the centrifugal compressor air flow passage.

4. The gas turbine engine according to claim 1, characterized in that The compressor air flow passage defines an axial compressor air flow passage, the compressor shroud includes a plurality of stator blades circumferentially arranged in one stage around the compressor shroud, and the plurality of compressor bleed air passages are arranged at a downstream side of the plurality of stator blades.

5. The gas turbine engine according to claim 1, characterized in that wherein the respective inlets of the portion of the remaining portion of the plurality of compressor bleed air passages are circumferentially spaced apart from one another by equal angular intervals.

6. The gas turbine engine according to claim 5, characterized in that wherein the respective inlets of the portion of the remaining portion of the plurality of compressor bleed air passages are circumferentially arranged between the first compressor bleed air passage and the second compressor bleed air passage.

7. The gas turbine engine according to claim 6, characterized in that wherein the inlet of a first portion of the portion of the remaining portions of the plurality of compressor bleed air passages is arranged adjacent to the first compressor bleed air passage and is arranged at the second angular spacing from the inlet of the first compressor bleed air passage, and the inlet of a second portion of the portion of the remaining portions of the plurality of compressor bleed air passages is arranged adjacent to the second compressor bleed air passage and is arranged at the second angular spacing from the inlet of the second compressor bleed air passage.

8. The gas turbine engine according to claim 1, characterized in that Wherein at least one of the plurality of compressor bleed air passages defines a diffuser flow passage, wherein the diffuser flow passage increases in size from the inlet to the outlet.

9. The gas turbine engine according to claim 8, characterized in that Wherein the inlet and the outlet are circular openings.

10. The gas turbine engine according to claim 9, characterized in that Wherein the diffuser flow passage defines a centerline extending from a center of the inlet to a center of the outlet.

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