Patent Publication Number: US-2023151825-A1

Title: Compressor shroud with swept grooves

Description:
TECHNICAL FIELD 
     The disclosure relates generally to aircraft engines and, more particularly, to compressors for such engines. 
     BACKGROUND 
     Compressor stall margin is one of many aspects that may affect the overall performance of aircraft engines. While compressor shrouds or casings may have various configurations in order to enhance rotor stall margin, such as surface treatment and/or structural modifications provided on the surface of the shroud, minimizing performance loss in this regard remains desirable. 
     SUMMARY 
     There is accordingly provided a compressor for an aircraft engine, comprising: a rotor having a plurality of blades mounted for rotation about a central axis, the plurality of blades having blade tips extending between leading and trailing edges; and a shroud surrounding the rotor and having an inner surface surrounding the blade tips, a plurality of grooves defined in said inner surface of the shroud adjacent said blade tips, the plurality of grooves extending circumferentially about the shroud and extending radially from groove inlet openings defined in the inner surface to closed end surfaces of the plurality of grooves, the plurality of grooves having sidewalls extending circumferentially about the central axis, the plurality of grooves being axially spaced-apart from each other, the groove inlet opening of the most upstream one of the plurality of grooves having an upstream end disposed upstream of the leading edges of the plurality of blades, the plurality of grooves having a swept angle from the inner surface such that a center of the groove inlet openings is axially offset of a center of a closed-end surface of each of the plurality of grooves, the plurality of grooves spanning an overall axial distance corresponding to 30% or more of a chord length of the plurality of blades, wherein the plurality of grooves have circumferential interruptions defined by a plurality of baffles such that the plurality of grooves extend non-continuously around a shroud circumference. 
     There is also provided a compressor for an aircraft engine, comprising: a rotor having a plurality of blades mounted for rotation about a central axis, the plurality of blades having blade tips extending between leading and trailing edges; and a shroud surrounding the rotor and having an inner surface surrounding the blade tips, a plurality of grooves defined in said inner surface of the shroud adjacent said blade tips, the plurality of grooves extending circumferentially about the shroud and extending radially from groove inlet openings defined in the inner surface to closed end surfaces of the plurality of grooves, the plurality of grooves having sidewalls extending circumferentially about the central axis, the plurality of grooves being axially spaced-apart from each other, the leading edge of the plurality of blades axially disposed between an upstream end of the groove inlet opening of the most upstream one of the plurality of grooves and a downstream end of the groove inlet opening of the most upstream one of the plurality of grooves, the plurality of grooves having a swept angle from the inner surface such that a center of the groove inlet openings is axially offset of a center of a closed-end surface of each of the plurality of grooves, wherein the plurality of grooves have circumferential interruptions defined by a plurality of baffles such that the plurality of grooves extend non-continuously around a shroud circumference. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG.  1    is a schematic cross sectional view of a gas turbine engine; 
         FIG.  2    is a schematic cross-sectional view of an exemplary part of the compressor rotor casing of the engine shown in  FIG.  1   ; 
         FIG.  3    is an enlarged perspective view of an exemplary part of the compressor rotor casing of  FIGS.  1 - 2   , defining a cross-section A-A and a cross-section B-B; 
         FIG.  3 A  is a schematic cross-sectional view taken through A-A in  FIG.  3   ; and 
         FIG.  3 B  is a schematic cross-sectional view taken through A-A of an alternate compressor rotor casing; 
         FIG.  4    is another perspective view of the exemplary part of  FIG.  3   , showing the cross-section B-B in a different angle; 
         FIG.  5    is a schematic cross-sectional view of another exemplary part of a compressor rotor casing of the engine shown in  FIG.  1   ; 
         FIG.  6    is a side view of an exemplary part of the compressor rotor casing of  FIG.  5   ; 
         FIGS.  7 A- 7 C  are graphical representations of various groove taper angles in a compressor rotor casing; and 
         FIGS.  8 A- 8 B  are schematic cross-sectional views taken through A-A in  FIG.  3    of various groove and baffle configuration options. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    illustrates a turbofan gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a transonic fan  12  through which ambient air is propelled, a multistage compressor  14  for pressurizing the air, a combustor  16  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section  18  for extracting energy from the combustion gases. 
     The fan  12 , also referred to as a low compressor, comprises a rotor  13  mounted for rotation about the engine central axis  11 . The rotor  13  is provided with a plurality of radially extending blades  15 . Each blade  15  has a leading edge  17  and a trailing edge  19  extending radially outwardly from the rotor hub to a tip  21 . The rotor  13  is surrounded by a casing  20  including a stationary annular shroud disposed adjacent the tips  21  of the blades  15  and defining an outer boundary for the main flow path. As shown in  FIG.  2   , the casing inner surface is lined with a layer of non-abradable material  22 . The layer of non-abradable material  22  may thus be considered as part of the casing inner surface, forming part of the hard shroud wall. In other cases, an abradable material that may detach or break from the casing  20  without causing damages, may be used. The radial distance or gap between the tip  21  of the blades  15  and the adjacent inner surface of the casing  20  is defined as the rotor tip clearance. Each rotor is designed with a nominal rotor tip clearance to prevent or limit interference between the tip  21  of the blades  15  and the casing  20 , which may occur due to rotor imbalance. 
     Referring to  FIG.  2   , it can be seen that a surface treatment is applied to the low pressure compressor or fan casing  20 , though such surface treatment may be applied to a high pressure compressor. As will be seen hereinafter, the surface treatment allows stall margin to be increased and/or tip clearance vortex flow to be weakened and may help to direct the vortex flow in the main flow stream direction. The rotor casing treatment comprises a series of regularly axially spaced-apart circumferential grooves  24  defined in the non-abradable region of the casing inner surface (region of the casing  20  having the layer of non-abradable material  22 ) axially aligned with the tips  21  of the blades  15 . Having regularly axially spaced-apart grooves  24 , as opposed to irregularly spaced-apart grooves may facilitate manufacturing and/or parametric design of the engine  10  and/or the surface treatment. In other cases, the grooves  24  may be irregularly or non-uniformly spaced apart in an axial direction along the casing inner surface, as will be discussed in further detail below. 
     As shown in  FIG.  3   , the grooves  24  do not extend continuously around 360 degrees. Stated differently, each groove  24  is intersected or interrupted over the circumference of the casing  20 . In other words, the grooves  24  have circumferential interruptions such that the grooves  24  extend non-continuously around a shroud circumference. In the depicted embodiment, the circumferential interruptions are defined by a plurality of baffles  30 . In other words, each groove  24  comprises a plurality of segments  24 A extending circumferentially and separated from an adjacent one of the segments  24 A by one of the baffles  30 . Although not “continuous” along the full circumference of the casing inner surface, each interrupted groove will be referred to as one groove  24  that comprises a plurality of groove segments  24 A, for simplicity. 
     In the illustrated example, six shallow circumferentially extending grooves  24  are embedded in the non-abradable layer  22  of the rotor shroud around the blades  15 . However, it is understood that the series of grooves  24  could be composed of more or less than six grooves  24 . For instance, the rotor casing treatment could comprise from 2 to 15 grooves depending on the rotor configuration. The grooves  24  may also be irregularly or non-uniformly axially spaced-apart in other embodiments. 
     Returning to  FIG.  2   , in the depicted embodiment, each groove  24  is defined by a pair of axially opposed sidewalls  26 , in this embodiment substantially flat, extending forwardly (i.e. towards the front of the engine) from a groove opening (or groove inlet)  25  defined in the shroud surface  27  to a closed-end surface  28 . The closed-end surface  28  may be flat, rounded or semi-circular in various embodiments, as will be discussed in further detail below. In the depicted embodiment, opposed sidewalls  26  of adjacent grooves  24  intersect at the opening (or “inlet”)  25  with the shroud surface  27 , corresponding to a portion of the casing inner surface between adjacent grooves  24 , forming a sharp edge. Such edge may be rounded up in other embodiments. Illustratively, each opening  25  includes an upstream end  25 A and a downstream end  25 B relative to the main flow through the compressor rotor. 
     As shown in  FIG.  2   , each groove  24  has a depth D and a width W. The grooves  24  are spaced apart from one another by a spacing X taken axially along the shroud inner surface  27  (distance between the opening of adjacent grooves  24 ). Such spacing may be equal between each pair of axially adjacent grooves  24 . In other cases, the spacing X between a first pair of axially adjacent grooves  24  may be different, i.e. greater or lesser in magnitude, than the spacing X between another pair of axially adjacent grooves  24 . Each groove  24  has a depth projection Y normal to the casing inner surface. 
     As depicted in  FIG.  2   , the groove inlet opening  25  of the first or upstream groove  24  is axially located upstream of the leading edge  17  of the blades  15 . More particularly, the upstream end  25 A of the groove inlet opening  25  of the first or upstream groove  24  is axially located upstream of the leading edge  17  of the blades  15  relative to the main flow through the compressor rotor. The upstream end  25 A is axially spaced from the leading edge  17  by a distance L corresponding to, for instance, 0% to 10% of the chord length of the blades  15 . Other distances may be contemplated as well. In the shown embodiment, although not necessarily the case in all embodiments, the leading edge  17  of the blades  15  is axially disposed between the upstream end  25 A and the downstream end  25 B of the groove inlet opening  25  of the first or upstream groove  24 . Other arrangements may be contemplated as well, for instance both the upstream end  25 A and the downstream end  25 B of the groove inlet opening  25  of the first or upstream groove  24  being axially disposed upstream of the leading edge  17 . In the depicted embodiment, the last or downstream groove  24  is positioned upstream of the blade trailing edges  19 . The grooves  24  may occupy an axial distance AD spanning from the first or upstream groove to the last or downstream groove corresponding to 30% or more of the chord length of the blades  15 . Illustratively, such axial distance AD may be taken from the upstream-most portion of the closed-end surface  28  of the first or upstream groove  24  to the downstream end  25 B of the last or downstream groove  24 . Other reference points for axial distance AD may be contemplated as well. Having the distance L and axial distance AD within these ranges may optimize their effect on the flow vortex. 
     In the shown case, the grooves  24  are forwardly swept (i.e. swept towards a front of the engine, which may also be upstream relative to the main gas flow through the compressor rotor) at an angle θ. In other words, when viewed axially along the tip  21  of a blade  15  from its leading edge  17  to its trailing edge  19 , such as in  FIGS.  2  and  4   , the closed-end surface  28  of each of the grooves  24  is located upstream of the opening  25  of the corresponding groove  24 . Alternately defined, the grooves  24  are inclined such that a center of their inlet openings  25  is located axially rearward of a center of their closed-end surfaces  28  with respect to the orientation of the grooves  24  of the casing  20  in the engine  10 . The angle θ is taken between an axis P normal to the casing inner surface  27  and a central axis GA extending longitudinally through a center of the grooves  24 . Angle θ may be referred to as the groove swept angle, or groove sweep angle, and is more than 0° and less than 75°. In an embodiment, the angle θ is at least 10° but no more than 75°. Due to the groove swept angle within this range, the swept angled grooves  24  may contribute to minimizing total pressure loss by having the flow exiting from the grooves  24  with a sufficient main flow stream direction component, and/or may allow maximizing an internal volume of the grooves  24  although the layer of non-abradable material  22  of the rotor casing may be thin, for maximizing compactness of the rotor casing  20  (to reduce weight and/or size of the rotor casing  20 ). In other embodiments, the grooves  24  may be rearwardly swept (i.e. swept towards a rear of the engine, which may also be downstream relative to the main gas flow through the compressor rotor) at an angle θ. In such cases, the groove swept angle, or groove sweep angle, may be less than 0° and more than -75° (i.e. a maximum angle of 75° in a rearward direction). In the depicted embodiment, the grooves are all angled identically, but one or more of the grooves  24  may have a different angle θ than other ones or more of the grooves  24  in other embodiments. 
     In one embodiment, the width W of the grooves  24  is between about 1% to about 15% of the chord length of the blades  15 . The spacing X may have any suitable value, for instance respecting an aspect ratio X/W is from about 0.1 to about 5. Other spacing X between grooves  24  may be contemplated, for instance irregular or uneven distributions. In one particular embodiment, the ratio Y/W ranges from about 0.5 to 10. In most cases, larger ratios may be better to trap the tip vortex, though manufacturing may limit the possibilities to have a greater ratio (e.g. a ratio greater or much greater than 10). 
     While in some embodiments the grooves  24  may all have a same geometry, one or more of the grooves may have a respective geometry that may differ in one or more dimensions, in some cases. 
     As shown in  FIGS.  2  and  4   , the respective depths D of the grooves  24  may vary from the first (most upstream groove  24 ) to the last, more particularly, in this case the respective depths D of the grooves  24  increase from the first to the last groove  24 , although they may all have an equal depth D in other embodiments. Depending on the embodiments, the respective depths D of the grooves  24  may substantially correspond to the thickness of the layer of non-abradable material  22  at the local areas where they are defined. Stated differently, the depth projection Y of the grooves  24  may substantially correspond to the thickness of the non-abradable material  22 . In other cases, the depths of the grooves  24  may increase or decrease at various rates, or remain constant, from the first to the last groove  24 , as will be discussed in further detail below. 
     Now referring to  FIG.  3   , the arrays of baffles  30  in the grooves  24  may be angularly aligned with respect to each other. However, the baffles  30  could as well be angularly staggered in the different grooves  24 . In addition, the number of baffles in the grooves  24  does not have to be the same. In an embodiment, the number of baffles  30  in each groove  24  is greater than the number of rotor blades  15  but less than 5 times of the latter. In a particular embodiment, the number of baffles  30  in each groove  24  is between 2 and 5 times the number of rotor blades  15 . In another particular embodiment, there are two times more baffles  30  per groove  24  than rotor blades  15 . Other ratios of baffles  30  per groove  24  may be contemplated as well. Having a greater number of baffles  30  per groove  24  may impede the effects of the casing treatment. 
     As shown in  FIG.  3 A , the baffles  30  are provided in the form of projections from the closed-end surface  28  of the grooves  24  to the inlet opening  25  thereof. That is, the baffles  30  protrude from the closed-end surface  28  over a distance corresponding to the full depth D of the groove  24  in which the baffles  30  are located. The baffles  30  do not necessarily have to be the same shape. The baffles  30  may be integrally machined, moulded or otherwise formed on the closed-end surface  28  of the grooves  24 . For instance, cutting tools, such as conventional wood ruff cutters, could be used for machining the grooves  24  and the baffles  30  in the non-abradable layer  22 . In this way, the baffles  30  can be formed in the grooves  24  in a cost effective manner. The reparability of the casing  20  may be good since the grooves  24  and the baffles  30  are machined in non-abradable material. 
     The depicted baffles  30  extend the full width W of the grooves  24  between the groove sidewalls  26  (see  FIG.  3   ). As shown in  FIG.  3   , each baffle  30  has a substantially flat surface  32  extending in the same plane as the shroud inner surface  27 . In other words, the flat surface  32  of the baffles  30  form a continuous surface with adjacent portions of the shroud inner surface. Forming such continuous surface with adjacent portions of the shroud inner surface may contribute to optimizing the effects of the casing treatment herein described. The flat surface  32  may have other shapes, such as concave or other non-flat shape in other embodiments. 
     As shown in  FIG.  3 A , the baffles  30  extends along the full depth D of the grooves  24 . This may maximize the break of the swirl component (circumferential component) of the main flow stream at the tip of the blades  15  (or simply tip vortex). In the depicted embodiment, the baffles  30  have two opposed walls  33  spaced apart circumferentially from each other and defining respective ends of the baffles  30  (i.e. ends that are spaced apart in the circumferential direction of the grooves  24 ). In the depicted embodiment, the two opposed walls  33  merge with the flat surface  32  to form a sharp edge at their junction, though rounded edges may be contemplated in other embodiments. The grooves closed-end surface  28  and the baffles  30  form an intersected radially inwardly facing surface at the closed end of each groove  24 , such that the radially inwardly facing surface is discontinuous along the length (defined along the circumference of the casing inner surface) of each groove  24 . Although such circumferentially intersected grooves  24  may generate flow turbulence due to the baffles  30  opposing the circumferential component of the tip flow vortex entering and exiting the grooves  24 , such turbulence resulting from the presence of the baffles  30  may be more beneficial to the performance of the engine  10  than if the baffles  30  were omitted entirely, where the circumferential component of the main flow stream (or tip vortex), would not be suitably controlled. The presence of groove interruptions, such as the baffles  30  herein described, may enhance the momentum exchanges between main flow and tip clearance flow, hence enhance the effect of the casing treatment. 
     Referring to  FIG.  3 B , another baffle configuration is shown. In the depicted embodiment, the baffles  30  another embodiment the baffles lean with an angle ϕ relative to the axis P normal to the casing inner surface  27 . In some embodiments, the angle ϕ may vary from -75° to +75°, i.e. into or away from a rotational direction of the blades  15 . The shape of the baffles  30  may vary. For instance, the edges of the baffles may be sharp or rounded. A width B of the baffles  30  may be constant along both radial and axial directions, for instance a tenth of the groove width W. In other cases, the baffle width B may vary in one or both of the radial and axial directions. The circumferential distribution of baffles may be uniform or uneven, or may assume other irregular patterns as well. 
     Referring to  FIG.  5   , another exemplary fan casing  20  is shown, with like reference numerals referring to like elements. The various features discussed in relation to the fan casing depicted in  FIG.  2    may be understood to be applicable to the fan casing depicted in  FIG.  5    as well, for instance the upstream end  25 A of the groove inlet opening  25  of the first or upstream groove  24  being axially located upstream of the leading edge  17  of the blades  15  relative to the main flow through the compressor rotor. Of note, in the fan casing  20  shown in  FIG.  5   , the closed-end surfaces  28  of the grooves  24  are rounded or semi-circular. Other shapes for the closed-end surfaces  28  may be contemplated as well. In addition, in the embodiment shown in  FIG.  5   , the depths D of each is the grooves  24  is constant from the most upstream groove  24  to the most downstream groove  24 . Other depths D, for instance increasing or decreasing depths along the downstream direction, may be contemplated as well. In the depicted embodiment, the grooves  24  each have a forward swept angle θ of 45° relative to axis P normal to the casing inner surface  27 . Other angles, including rearward swept angles, may be contemplated as well. 
     Referring to  FIG.  6   , the depicted casing  20  includes unevenly-spaced grooved  24 . In other words, spacing X 1  between a first pair of grooves  24  is different than spacing X 2 , X 3 , X 4 , etc. In the depicted case, the ratio between spacing X (X 1 , X 2 , X 3 , X 4 ) and the groove width W (X/W) may vary between 0.5 and 5. In other embodiments, the ratio (X/W) may vary between 3 and 3.6. Other ratios may be contemplated as well. As discuss above, and in the depicted case, the groove depth D may be consistent for each groove  24 . In the depicted case, each groove  24  includes a rounded or semi-circular closed-end surface  28 . 
     Referring to  FIGS.  7 A- 7 C , in various embodiments, the taper angle of the grooves  24 , i.e. the variation in radius from one groove  24  to the next, can either remain constant (ex:  FIG.  7 A ), decrease (Ex:  FIG.  7 B ) or increase (EX:  FIG.  7 C ) from an upstream end to a downstream end of the casing  20 . In  FIG.  7 A , the taper angle is shown to remain constant, i.e. a taper angle of 0° between grooves  24 . In  FIG.  7 B , an exemplary inward or decreasing taper angle of 10°, is shown. In  FIG.  7 C , an exemplary outward or increasing taper angle of 10° is shown. Other inward or outward taper angles may be contemplated. For instance, in various cases the taper angle may vary from 20° inward to 20° outward. 
     Referring to  FIGS.  8 A- 8 B , the grooves  24  may take on various shapes or patterns when viewed from cross-section A-A. For instance, the grooves  24  depicted in  FIG.  7 A  are shown to have a linearly-circumferential shape, while the grooves  24  depicted in  FIG.  7 B  are shown to have non-linear or curved shape. Other groove patterns or shapes, or instance for instance helically-threaded grooves with baffles, may be contemplated as well. 
     In the present disclosure, when a specific numerical value is provided (e.g. as a maximum, minimum or range of values), it is to be understood that this value or these ranges of values may be varied, for example due to applicable manufacturing tolerances, material selection, etc. As such, any maximum value, minimum value and/or ranges of values provided herein (such as, for example only, the plurality of grooves spanning an overall axial distance corresponding to 30% or more of a chord length of the plurality of blades), include(s) all values falling within the applicable manufacturing tolerances. Accordingly, in certain instances, these values may be varied by ± 5%. In other implementations, these values may vary by as much as ± 10%. A person of ordinary skill in the art will understand that such variances in the values provided herein may be possible without departing from the intended scope of the present disclosure, and will appreciate for example that the values may be influenced by the particular manufacturing methods and materials used to implement the claimed technology. 
     The embodiments described in this document provide non-limiting examples of possible implementations of the present technology. Upon review of the present disclosure, a person of ordinary skill in the art will recognize that changes may be made to the embodiments described herein without departing from the scope of the present technology. Yet further modifications could be implemented by a person of ordinary skill in the art in view of the present disclosure, which modifications would be within the scope of the present technology.