• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
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    • 专利摘要:
    An axial compressor (30) for a gas turbine contains one or more end wall devices for controlling a leakage flow in the compressor (30). The one or more end wall devices have a height (112), which is formed and arranged in an inner surface (83) of a compressor housing (82) or a compressor hub, in order to create a flow (53) adjacent to a plurality of blade tips (81) or a plurality of guide blade tips ( 87) to a cylindrical flow passage (56) upstream from a tapping point of the flow (53). The end wall devices each define a front wall (102), a rear wall (104), an outer wall that runs between the front wall and the rear wall, an axial projection (108), an axial overlap (110), an axial angle of inclination (α1) and a tangential Tilt angle (β1). The axial projection (108) extends upstream to project at least over one of the at least one blade set and the at least one guide blade set. The axial overlap (110) is downstream to overlap at least one of the at least one blade set and the at least one guide blade set.
    公开号:CH710476B1
    申请号:CH01709/15
    申请日:2015-11-23
    公开日:2020-05-15
    发明作者:David Stampfli John；Venkata Mallina Ramakrishna；Michelassi Vittorio；Jothiprasad Giridhar；Keshava Rao Ajay；Konrad Selmeier Rudolf；Giacché Davide；Malcevic Ivan；Yoon Sungho
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    BACKGROUND
    The invention described herein relates generally to gas turbines and, more particularly, relates to an axial compressor end wall device for a gas turbine to control the leakage flow therein.
    As is known, an axial compressor for a gas turbine may include a number of stages arranged along an axis of the compressor. Each stage can include a rotor disk and a number of compressor blades arranged around a circumference of the rotor disk, which are also referred to herein as rotor blades. In addition, each stage may further include a number of vanes located adjacent to the blades and arranged around a circumference of the compressor housing.
    [0003] During operation of a gas turbine using a multi-stage axial compressor, a turbine rotor is rotated by a turbine at high speeds so that air is continuously introduced into the compressor. The air is accelerated by the rotating compressor blades and propelled backwards onto the adjacent rows of guide blades. Each blade / vane level increases the pressure of the air. In addition, during operation some of the compressed air may leak downstream around a tip of each of the compressor blades and / or guide blades. Such a step-to-step leakage of compressed air as a leakage flow can influence the stall point of the compressor.
    Compressor flow breaks can reduce the compressor pressure ratio and reduce the air flow delivered to a combustor, thereby compromising the efficiency of the gas turbine. Rotating flow separation in an axial type compressor usually occurs at a desired peak operating point of the compressor. After rotating flow separation, the compressor may transition to a compressor pump condition or a deep stall condition, which may result in a loss of efficiency and, if allowed to last longer, may result in gas turbine failure.
    The operating range of an axial compressor is generally limited due to weak flow in rotor tips, where the specific stall point of the rotor is determined by the operating conditions and the compressor design. Attempts in the prior art to enlarge the area of this operation and to enlarge the stall boundary area include methods based on flow regulation, such as, for example, control by means of plasma actuators and suction / blowing near a blade tip. However, such attempts significantly increase the complexity and weight of the compressor. Other attempts include end wall devices such as circumferential grooves, axial grooves or the like. Initial trials had a significant impact on design point efficiency with very minimal benefits for the stall boundary.
    There is therefore a desire for an improved axial compressor for a gas turbine and a method for controlling the leakage flow around one or more blade tips on these. Specifically, such a compressor can control leakage of compressed air through a carefully designed end wall device near the blades and / or vanes, which results in a desired circulation of the leakage flow. Leakage control of this type can increase the overall operating range and the limit range of the compressor pumping of the compressor and the gas turbine, while minimizing the adverse effect on efficiency at the design point.
    SUMMARY OF THE INVENTION
    Aspects and advantages of the disclosure are set forth in the description below, or may be obvious from the description, or may be learned by practice of the disclosure.
    [0008] In one aspect, a compressor is created. The compressor includes a compressor end wall that defines a substantially cylindrical flow passage. The compressor end wall includes a compressor housing and a compressor hub, which are arranged concentrically about and coaxially along a longitudinal central axis, at least one set of blades, at least one set of guide vanes and one or more end wall devices with a radial height, which in an inner surface of the housing and / or the hub are formed. Each of the at least one rotor blade set contains a plurality of rotor blades which are connected to the compressor hub and run between the compressor hub and the compressor housing and there define a blade passage between all the rotor blades. The compressor housing circumscribes the at least one rotor blade set in order to define an annular gap between the compressor housing and a plurality of rotor blade tips of the plurality of rotor blades. Each of the at least one guide vane set contains a plurality of guide vanes which are connected to the compressor housing and run between the compressor housing and the compressor hub and there define a vane passage between all the guide vanes. The guide vanes are arranged relative to the compressor hub to define an annular gap between the compressor hub and a plurality of guide vane tips of the plurality of guide vanes. The one or more end wall devices are configured to return a flow adjacent the plurality of blade tips or vane tips to the cylindrical flow passage upstream of a tapping point of the flow. The end wall devices each define a front wall that contains a first axial angle of inclination α1 relative to the longitudinal central axis, a rear wall that contains a second axial angle of inclination α2 relative to the longitudinal central axis, an outer wall that extends between the front wall and the rear wall, an axial projection that extends extending upstream, to project at least over one of the at least one blade set or the at least one guide blade set, an axial overlap extending downstream to overlap at least one of the at least one blade set or the at least one guide blade set relative to a first tangential inclination angle β1 a peripheral surface of the compressor end wall and a second tangential inclination angle β2 relative to the peripheral surface of the compressor end wall. Either the axial angle of inclination α1 is not equal to the axial angle of inclination α2 or the tangential angle of inclination β1 is not equal to the tangential angle of inclination β2.
    In the compressor mentioned above, the one or more end wall devices can have a plurality of separate axial recesses which are defined in the circumferential direction to extend around the compressor hub and / or the compressor housing.
    [0010] Specifically, each vane passage can contain 0 to 10 separate axial recesses.
    In the compressor of any type mentioned above, the one or more end wall devices may have a radial height that is in the range of 5 to 50% of a span of the plurality of blades and / or the plurality of vanes.
    Furthermore, the first axial inclination angle α1 and the second axial inclination angle α2 can be in a range from 10 to 170 degrees.
    [0013] Still further, the first tangential inclination angle β1 and the second tangential inclination angle β2 can be in a range from 10 to 170 degrees.
    In one embodiment, the first axial tilt angle, the second axial tilt angle, the first tangential tilt angle and the second tangential tilt angle cannot be the same.
    In a further embodiment, the axial projection can be -10% to 60% of a blade chord length.
    [0016] In one embodiment, the axial projection can be 0% of a blade chord length.
    [0017] In yet another embodiment, the axial overlap can be -10 to 60% of a blade chord length.
    In one embodiment, the axial overlap can be 0% of a blade chord length.
    In any compressor mentioned above, a non-metal area of the recess may be 10% to 90% of an area of the blade passage.
    In a still further aspect, an engine or an engine is created. The engine or engine includes a fan assembly and a core engine downstream of the fan assembly. The core engine contains a compressor, a combustion chamber and a turbine. The compressor, combustor, and turbine are configured in a downstream axial flow relationship. The compressor further includes a compressor end wall defining a generally cylindrical flow passage, at least one blade set, at least one guide blade set, and one or more end wall means. The compressor end wall includes a compressor housing and a compressor hub which are arranged concentrically about and coaxially along a longitudinal central axis. Each of the at least one rotor blade set contains a plurality of rotor blades which are connected to the compressor hub and run between the compressor hub and the compressor housing. The compressor housing circumscribes the at least one rotor blade set in order to define an annular gap between the compressor housing and a plurality of rotor blade tips of the plurality of rotor blades. Each of the at least one guide vane set contains a plurality of guide vanes which are connected to the compressor housing and run between the compressor housing and the compressor hub. The guide vanes are positioned relative to the compressor hub to define an annular gap between the compressor hub and a plurality of guide vane tips of the plurality of guide vanes. The one or more end wall devices have a height formed in an inner surface of the housing and are configured to return a flow adjacent the plurality of blade tips to the cylindrical flow passage upstream of a tapping point of the flow. The one or more end wall devices each define a front wall that has a first axial angle of inclination α1 relative to the longitudinal central axis, a rear wall that has a second axial angle of inclination α2 relative to the longitudinal central axis, an outer wall that extends between the front wall and the rear wall, an axial projection extending upstream to cantilever over at least one of the one or more sets of vanes or one or more sets of vanes, an axial overlap extending downstream to overlap at least one of the at least one set of vanes or at least one set of vanes first tangential angle of inclination β1 relative to a peripheral surface of the compressor end wall and a second tangential angle of inclination β2 relative to the peripheral surface of the compressor end wall, wherein the axial angle of inclination α1 is not equal to the axial angle of inclination α2 and / or the tangential angle of inclination kel β1 is not equal to the tangential inclination angle β2.
    In the aforementioned engine, the first axial inclination angle α1, the second axial inclination angle α2, the first tangential inclination angle β1 and the second tangential inclination angle β1 can be in a range from 10 to 170 °.
    Preferably, the core engine is set up for use in an aircraft engine.
    BRIEF DESCRIPTION OF THE DRAWINGS
    [0023] A detailed and enabling disclosure of the subject matter of the present disclosure, including the best mode thereof, to those skilled in the art is set forth in more detail in the remainder of the description with reference to the accompanying figures, in which:
    Figure 1 is a schematic longitudinal section of part of an aircraft engine having a compressor having end wall devices according to one or more embodiments shown or described herein;
    Figure 2 is a schematic longitudinal section of part of a compressor known in the art;
    3 is a schematic longitudinal section of a portion of the compressor of the aircraft engine of FIG. 1 having an end wall assembly in accordance with one or more embodiments shown or described herein;
    FIG. 4 is a schematic longitudinal section of the compressor of FIG. 3 having an end wall device in accordance with one or more embodiments shown or described herein;
    5 is a schematic isometric view of a portion of the compressor of FIG. 4 having end wall means in accordance with one or more embodiments shown or described herein;
    Figure 6 is a schematic longitudinal section of another embodiment of a compressor having end wall means according to one or more embodiments shown or described herein;
    Figure 7 is a schematic longitudinal section of another embodiment of a compressor having end wall means according to one or more embodiments shown or described herein;
    Figure 8 is a schematic longitudinal section of another embodiment of a compressor having end wall means according to one or more embodiments shown or described herein;
    9 is a schematic axial cross-section of the compressor of FIG. 7 along line 9-9 having end wall means according to one or more embodiments shown or described herein;
    Fig. 10 is a schematic axial cross section of another embodiment of a compressor having an end wall device, according to one or more embodiments shown or described herein, and
    11 is a graph illustrating the utility of a compressor having the one or more end wall devices according to one or more of the embodiments shown or described herein.
    In the different views of the drawings, corresponding reference numerals indicate corresponding parts throughout.
    DETAILED DESCRIPTION
    The present disclosure is described for purposes of illustration only in connection with certain embodiments; however, it is to be understood that other objects and advantages of the present disclosure will become apparent from the following description of the disclosed drawings. While preferred embodiments are disclosed, they are not intended to be limiting. Rather, the general principles set forth herein are merely illustrative of the scope of the present disclosure, and it should also be appreciated that numerous changes can be made without departing from the scope of the present disclosure.
    Preferred embodiments of the present disclosure are illustrated in the figures, wherein like reference numerals are used to refer to the same and corresponding parts of the different drawings. In addition, reference throughout the specification to "a single embodiment", a "further embodiment", "an embodiment" and so on means that a particular element (e.g. feature, structure and / or property) described in connection with the embodiment. is included in at least one embodiment described herein and may or may not be included in other embodiments. It is understood that the inventive features described can be combined in the various embodiments in any suitable manner. It is also understood that terms such as “above”, “below”, “away”, “inward” and the like are useful words and should not be interpreted as restrictive terms. It should be noted that the terms "first", "second" and the like, as used herein, do not mean order, quantity or meaning, but rather are used to distinguish one element from another. The terms “a” and “one” do not denote a quantitative restriction, but rather denote the presence of at least one of the elements mentioned. The modification "about" used in connection with a quantity is to be understood including the given value and has the meaning prescribed by the context (e.g. contains the degree of error associated with measuring the specific quantity).
    [0038] Embodiments disclosed herein relate to a compressor device of an aircraft engine, which has one or more end wall devices for controlling the leakage through the compressor. In contrast to known means for controlling a leakage flow through a compressor, the end wall devices as disclosed herein enable the limit of the operability of the compressor to be increased, the efficiency disadvantage of the compressor to be minimized and the resulting stalling of the rotor to be delayed as a result.
    Referring to the drawings, wherein identical reference numerals denote the same elements throughout the different views, FIGS. 1 and 2 represent, for example, a schematic representation of an exemplary aircraft engine assembly 10. The embodiments described herein are equally based on a stationary one Gas turbine type, such as a gas turbine used for industrial applications. It is pointed out that the part of the engine arrangement 10 which is illustrated in FIG. 3 is indicated in FIG. 1 with a dashed line. The engine assembly 10 has a longitudinal centerline or axis 12 and an outer stationary annular fan housing 14 which is concentric about and coaxial along the longitudinal axis 12. In addition, the engine assembly 10 has a radial axis 13. In the exemplary embodiment, the engine assembly 10 includes a fan assembly 16, an auxiliary compressor 18, a core gas turbine engine 2,0, and a low pressure turbine 22, 22 that may be connected to the fan assembly 16 and the auxiliary compressor 18. The fan assembly 16 includes a plurality of fan blades 24 that extend substantially radially outward from a fan rotor disk 26, as well as a plurality of structural strut members 28 and outlet vanes (“OGVs”) 29 that may be positioned downstream of the fan blades 24. In this example, separate elements are provided for the aerodynamic and structural functions. In other configurations, each of the OGV s29 can be both an aerodynamic element and a structural support for an annular fan housing. The auxiliary compressor includes a plurality of blades 35 that extend substantially radially outward from a compressor rotor disk or hub 37 connected to a first drive shaft 40.
    The core gas turbine engine 20 includes a high pressure compressor 30, a combustor 32 and a high pressure turbine 34. The high pressure compressor 30 includes a plurality of blades 36 which extend radially outwardly from a compressor hub 38. The high-pressure compressor 30 and the high-pressure turbine 34 are connected to one another by a second drive shaft 41. The first and second drive shafts 40 and 41 are rotatably mounted in bearings 43, which in turn are mounted in a fan frame 45 and a rear turbine frame 47. The engine assembly 10 also includes an intake side 44 that defines a fan inlet 49, a core engine exhaust side 46 thereof, and a fan exhaust side 48.
    During operation, the fan assembly 16 compresses air entering the engine assembly 10 through the suction side 44. The air flow emerging from the blower arrangement 16 is divided so that a part 50 of the air flow is passed as a compressed air flow into the additional compressor 18 and a remaining part 52 of the air flow is guided past the additional compressor 18 and the core gas turbine engine 20 and via a bypass duct 51 through the Blower exhaust side 48 exits the engine assembly 10 as bypass air. More specifically, the bypass duct 51 runs between an inner wall 15 of the fan housing 14 and an outer wall 17 of an additional compressor housing 19. This part 52 of the air flow, which is also referred to herein as the bypass or secondary air flow 52, flows on the structural strut elements 28, the outlet guide vanes 29 and a heat exchanger device 54 and interacts with them. The plurality of fan blades 24 compress and deliver the compressed airflow 50 to the core gas turbine engine 20. Furthermore, the airflow 50 is further compressed by the high pressure compressor 30 and delivered to the combustor 32. The compressed air flow 50 from the combustion chamber 32 also drives the rotating high-pressure turbine 34 and the low-pressure turbine 22 and exits the engine arrangement 10 through the core engine exhaust side 46.
    Referring now to Fig. 2, a portion of a compressor 60 is shown schematically which is well known in the art and is referred to as the prior art. Compressor 60 includes a plurality of blade sets 62 which are circumferentially spaced apart and which extend radially outward from a compressor hub 66 toward a compressor housing 64. A plurality of sets of circumferentially spaced vanes 68 (only a single vane is shown) are positioned adjacent each vane set 62 and in combination form one of a plurality of stages 70 (only a single stage is shown). Each of the guide vanes 68 is securely connected to the compressor housing 64 and extends radially inward for coupling to the compressor hub 66. Each of the blades 62 is surrounded by the compressor housing 64, so that an annular gap 72 is defined in the blade set 62 between the compressor housing 64 and a blade tip 63 of each blade. Likewise, the guide blades 68 are arranged relative to the compressor hub 66 in such a way that an annular gap 73 is defined between the compressor hub 66 and a guide blade tip 69 of each of the guide blades 68.
    [0043] During operation, an operating range of the compressor 60 is generally limited near the blade tips 63 due to a leakage flow, as indicated by directional arrows 74. There may also be a leakage flow (not shown) near the vane tips 69. A specific stall point of the rotor is determined by the operating conditions and the compressor design. To increase the range of this operation, previous compressors, in an attempt to increase the operating range by redirecting and / or minimizing leakage flow 74, included end wall devices (not shown), such as circumferential grooves.
    In Fig. 3, to which reference is now made in more detail, part of the new compressor 30 as presented in Fig. 1 is illustrated. As illustrated, the aircraft engine assembly 10, and more particularly the compressor 30, in the exemplary embodiment includes at least one set of blades 76, each set having a plurality of blades 80 that are circumferentially spaced apart from a compressor hub or rotor disc 84 that coexists of the first drive shaft 40, extend radially outward in the direction of a compressor housing. At least one set of vanes 78, each set having a plurality of circumferentially spaced vanes 86, is positioned adjacent to each set of blades 76 and in combination forms one of a plurality of stages 88. The vanes 86 are securely connected to the compressor housing 82 and extend for coupling with the compressor hub 84 radially inwards. Each of the plurality of stages 88 conducts a flow of compressed air through the compressor 30. The blades 80 are surrounded by the compressor housing 82, so that an annular gap 90 is defined between the compressor housing 82 and a blade tip 81 of each of the blades 80. Likewise, the guide blades 86 are arranged relative to the compressor hub 84 such that an annular gap 92 is defined between the compressor hub 84 and a guide blade tip 87 of each of the guide blades 86.
    As is common in the art, each gap 90 and 92 is sized to allow a compressed amount of air 50, which defines the leakage flow 74 (FIG. 2), to flow around the blades 80 and vanes 86, respectively minimize. To allow circulation of this portion of compressed air 50 near the blade tips 81 and / or the guide blade tips 87, the new compressor 30 disclosed herein has one or more end wall devices 94. The term "end wall" as used herein is intended to refer to the compressor housing 82 and / or include the compressor hub 84 and provide a generally cylindrical flow passage 56.
    Referring now to Figs. 4 and 5, Fig. 4 schematically illustrates a longitudinal section through a portion of compressor 30 having one or more end wall devices 94 (only one of which is shown). 5 illustrates, in a schematic isometric view, the one or more end wall devices 94 and positioning relative to a blade 80, with part of the housing 82 removed for illustration. In this particular embodiment, as illustrated, the one or more end wall devices 94 are configured as a plurality of separate recesses 96 which are formed in an inner surface 83 of the compressor 82 and are arranged there in the circumferential direction near the rotor blade tips 81. Each of the recesses of the plurality of recesses 96 is aligned essentially along the main axis and more particularly along the longitudinal central axis 12 (FIG. 1), so that a current circulation 98 takes place in these recesses generally along this main direction. As indicated by the directional arrow 98 of the flow recirculation, the one or more end wall devices 94 are configured to circulate 98 the flow 50 adjacent the plurality of blade tips 81 and, more specifically, to the cylindrical flow passage 56 upstream of a tapping point for the flow 50. Each recess 96 has a cross section in the plane of this main direction, which supports a recirculation 98 of the flow over the blade tip 81. The position of each of the recesses 96, the orientation, the cross-sectional definition and additional geometric parameters can be optimized to provide a specific solution for each application that wishes to enlarge a stable operating area.
    Specifically, in the exemplary embodiment illustrated, the one or more end wall devices 94, and in particular the plurality of separate recesses 96, enable the adverse effect of leakage flows of compressed air between the compressor housing 82 and the blade tip 81 to be reduced. More specifically, the plurality of separate recesses 96 allow the Conversion of the uselessness of leakage flows into useful flows to enlarge the stall boundary. During operation, the portion of air flow 50 flows through fan inlet 49 (FIG. 1) into aircraft engine assembly 10 and toward compressor 30. The guide vanes 86 direct the compressed air towards the rotor blades 80. The compressed air extracts additional work from the blades 89 that rotate about the longitudinal central axis 12 of the compressor 30 while the vanes 86 remain stationary and further compress the air flowing through each of the plurality of stages 88. In this way, the blades 80 cooperate with the adjacent guide blades 86 in order to impart kinetic energy to the inflowing air stream 50, which is then fed to the combustion chamber 32, and to compress it. Other types of compressor configurations can be used.
    The one or more end wall devices 94, and in particular the plurality of separate recesses 96, help delay the stall of the rotor by initially having a weak peak flow through a rear segment 100 of a portion 58 of the stream 50, herein also referred to as leakage flow, which the blade tip 81 is exposed, withdraws or withdraw. The portion 58 of the flow 50 is then circulated and strengthened within each of the recesses 96 and is blown back into the main flow 50 through the front segment as a re-blown flow 59 in front of the blade 80. It is to be understood that the position of the plurality of recesses 96 relative to the blade tips 81, the circumferential distribution around the housing 82 and the repeating pattern of the plurality of recesses 96 are shown for illustration purposes only. In practice, the specific design of the one or more end wall devices 94 is optimized for the application in which they are used.
    Referring again to FIG. 4, the plurality of recesses 96 are configured relative to the plurality of blades 80 and, in particular, the blade tips 81. As illustrated, each of the plurality of recesses 96 is defined by a front wall 102, a rear wall 104, and an outer wall 106 between the front wall 102 and the rear wall 104. Each of the plurality of recesses 96 is furthermore an axial projection 108, an axial overlap 110, a radial height 112, a first axial inclination angle α1 to the longitudinal central axis 12 (FIG. 1), a second axial inclination angle α2 to the longitudinal central axis 12 (FIG. 1), one first tangential inclination angle and a second tangential inclination angle (described above). In one embodiment, the axial projection 108 extends upstream of the blades 80 and specifically extends with a blade leading edge tip 81 of the blades 80 corresponding to the front wall 102. The axial projection 108 can vary between -10% and 60% of the axial chord “y”. It is understood that an axial projection 108 of -10% of the axial chord “y” means that the front wall 102 of the recess 96 is located 10% downstream of the front blade tip corner 81. The axial overlap 110 extends from the blade leading edge tip 81 of the blades 80 in a downstream direction, thereby substantially overlapping a portion of the blades 80. The axial overlap 110 can vary between -10% and 100% of the axial chord “y”. It is understood that an axial overlap 110 of -10% of the axial chord “y” means that the rear wall 104 of the recess 96 is 10% upstream of the front blade edge tip 81. In one embodiment, the radial height 112 of each of the plurality of recesses 86 is approximately 5 to 50% of the span "x" of the blades 80.
    As already indicated and illustrated, the front wall 102 and the rear wall 104 of each of the plurality of recesses 96 are designed independently so that they refer to the longitudinal central axis 12 at one or more angles, which are referred to herein as axial inclination angles α1 and α2 of the housing 82 incline. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 are between 10 and 170 degrees. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 may be the same. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 cannot be the same. In one embodiment, the first axial angle of inclination α1 is aligned with the incoming main flow 50 to minimize the mixing loss between the incoming flow 50 and the re-injected flow 59 from each of the plurality of recesses 96. In contrast, the second axial inclination angle α2 is designed to effectively extract low-pulse fluids from the main flow 50.
    Referring now to Fig. 6, part of another embodiment of a compressor 120 is illustrated which is generally similar to the compressor 30 of Figs. 3-5. As already indicated, the same elements have the same reference numerals throughout the disclosed embodiments. Similar to the previously disclosed embodiment, compressor 120 includes a plurality of blades 80 that are circumferentially spaced apart and that extend radially outward from a compressor hub 84 toward a compressor housing 82 to a blade tip 81. A plurality of circumferentially spaced vanes 86 are positioned adjacent each blade set 80 and, in combination, form one of a plurality of stages 88. The vanes 86 are securely connected to the compressor housing 82 and extend radially inward from the compressor housing 82 toward the compressor hub 84 a vane tip 87. Each of the plurality of stages 88 directs a flow of compressed air through the compressor 30.
    In this particular embodiment, the new compressor 120 includes one or more end wall devices 94 which are configured as a plurality of separate recesses 96 which extend circumferentially around both the housing 82 and the hub 84 in order to circulate this part 58 to provide the compressed air 50 near the blade tips 81 and the guide blade tips 87. In particular, the recesses 96 in this special embodiment are embedded both in an inner surface 89 of the hub 85 in the hub component and in an inner surface 83 of the housing 82 in the housing component. It should be understood that an embodiment is provided that includes a plurality of recesses 96 that are only embedded in the hub component.
    [0053] The plurality of recesses 96 are set up relative to the plurality of blades 80, and more particularly to the blade tips 81 and the guide blades 86, and more particularly to the guide blade tips 87. As in the previous embodiment, each of the plurality of recesses 96 is defined by a front wall 102, a rear wall 104 and an outer wall 106 between the front wall 102 and the rear wall 104. Each of the plurality of recesses 96 is further defined by an axial projection 108, an axial overlap 11,0, a radial height 112, a first axial inclination angle α1, a second axial inclination angle α2, a first tangential inclination angle and a second tangential inclination angle (as in the present case) described).
    With respect to the axial recess provided near the blades 80, the axial projection 108 extends upstream of the blades 80 and, more particularly, extends with a front blade edge tip 81 of the blades 80 coincident with the front wall 102. The axial projection 108 can vary between -10% and 60% of the axial chord "y". It is understood that an axial projection 108 of -10% of the axial chord "y" means that the front wall 102 of the recess 96 is 10% downstream of the front blade edge tip 81. The axial overlap 110 extends from the front blade edge tip 81 of the blades 80 in a downstream direction, thereby substantially overlapping a portion of the blades 80. The axial overlap 110 can vary between -10% and 100% of the axial chord “y”. It is understood that an axial overlap 110 of -10% of the axial chord “y” means that the rear wall 104 of the recess 96 is 10% upstream of the front blade edge tip 81.
    With respect to the axial recess 96, which is provided near the guide vanes 86, the axial projection 108 extends upstream of the guide vanes 86 and, more particularly, extends with a front vane edge tip 87 of the guide vanes 86 coincident with the front wall 102 axial projection 108 can vary between -10% and 60% of the axial chord “y”. It is understood that an axial projection 108 of -10% of the axial chord "y" means that the front wall 102 of the recess 96 is 10% downstream of the rear blade edge tip 87. The axial overlap 110 extends from the front blade edge tip 87 of the guide blades 86 in a downstream direction, whereby it essentially overlaps a part of the guide blades 86. The axial overlap 110 can vary between -10% and 100% of the axial chord “y”. It is understood that an axial overlap 110 of -10% of the axial chord “y” means that the rear wall 104 of the recess 96 is 10% upstream of the front blade edge tip 87. In one embodiment, the radial height 112 of each of the plurality of recesses 96 is about 5 to 50% of the span "x" of the blades 80 and the vanes 86.
    As already indicated and illustrated, the front wall 102 and the rear wall 104 of each of the plurality of recesses 96 are independently designed such that they refer to the longitudinal central axis 12 at one or more angles, referred to herein as axial inclination angles α1 and α2 of the housing 82 incline. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 are between 10 and 170 degrees. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 may be the same. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 may be different. In one embodiment, the first axial angle of inclination α1 is aligned with the incoming main flow 50 to minimize the mixing loss between the incoming flow 50 and the re-blown flow 59 from each of the plurality of recesses 96. In contrast, the second axial inclination angle α2 is designed to effectively extract low-pulse fluids from the main flow 50.
    The embodiment disclosed in Figures 3 to 6 includes, as illustrated, one or more end wall devices 94 in the form of a plurality of axial recesses 96. As illustrated, each of the axial recesses 96 includes a geometric shape that has an overall curvature from the front wall 102 to the back wall 104. The correct choice of curvature can minimize aerodynamic losses in the recesses. Each of the axial recesses 96 can be optimized to provide a specific solution for any application that wishes to increase the stable operating range. Some of the aspects that can be optimized include: (i) the axial angle of inclination α1 of the front wall 102 and the axial angle of inclination α2 of the rear wall 104 of the recess 96; (ii) the tangential angles of inclination (as described herein) of the recess 96; (iii) radial height 112 of recess 96; (iv) a length of axial overhang 108 and the length of axial overlap 110; (v) a tangential distance between the recesses 96 and within each recess 96 (as described herein), (vi) a number of recesses 96 spaced circumferentially around the end wall (as described herein); (viii) an overall geometric cross section of each recess 96 when viewed in a radial-axial plane; and (viii) a variation of the above parameters in the radial, axial and tangential directions.
    FIGS. 7 to 9, to which reference is now made, illustrate sections of other embodiments of a compressor 130 which is substantially similar to the compressor 30 of FIGS. 3 to 5. As already indicated, the same elements have the same reference numerals throughout the disclosed embodiments. Similar to the embodiment disclosed above, the compressor 130 includes a plurality of blades 80 that are circumferentially spaced apart and that extend radially outward from a compressor hub 84 toward a compressor housing 82 to a blade tip 81. A plurality of circumferentially spaced vanes 86 are positioned adjacent to each blade set 80, connected to the compressor housing 82, and extend radially inward from the compressor housing 82 toward the compressor hub 84 to a vane tip 87, and in combination form one of a plurality of stages 88 Guide vanes 86 are securely connected to the compressor housing 82 and extend radially inward from the compressor housing 82 to a guide vane tip 87 in the direction of the compressor hub 84. The compressor 132 is about a longitudinal central axis 12 (FIG. 1) of the engine 10 (FIG. 1) rotatable, as indicated by a directional arrow 133.
    In the embodiments of FIGS. 7 to 9, the new compressor 130 contains one or more end wall devices which are set up as a plurality of recesses 132 which extend circumferentially around the housing 82 in order to circulate this part 58 of the compressed air 50 to worry near the blade tips 81. In the illustrated embodiments, the plurality of recesses 132 are shown embedded in the housing component. It is to be understood that an embodiment is provided that includes multiple recesses that are only embedded in the hub component or multiple recesses that are embedded in both the hub and the housing component.
    The plurality of recesses 132 are set up relative to the plurality of blades 80, and more particularly, the blade tips 81. In another embodiment, the plurality of recesses 132 may be embedded in the hub component or both in the hub component and in the housing component. As with the previous embodiments, each of the plurality of recesses 132 is defined by a front wall 102, a rear wall 104 and an outer wall 106 between the front wall 102 and the rear wall 104. Each of the plurality of recesses 132 is also relative to a circumferential surface of the compressor end wall and a second tangential inclination angle by an axial projection 108, an axial overlap 110, a radial height 112, a first axial inclination angle α1, a second axial inclination angle α1, a first tangential inclination angle β1 Defined β2 relative to a circumferential surface of the compressor wall, as best illustrated in FIG. 9. As in the embodiment disclosed above, the first axial inclination angle α1 and the second axial inclination angle α2 are between 10 and 170 degrees. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 may be the same. In one embodiment, the first axial tilt angle α1 and the second axial tilt angle α2 may be different. In one embodiment, the first axial angle of inclination α1 is aligned with the incoming main flow 50 to minimize the mixing loss between the incoming flow 50 and the re-injected flow 59 from each of the plurality of recesses 96. In contrast, the second axial angle of inclination α2 is designed to effectively withdraw low-pulse fluids from the main flow 50.
    In the illustrated embodiment of FIG. 7, the axial projection 108 extends upstream of the blades 80 and more particularly extends with a front blade edge tip 81 of the blades 80 coincident with the front wall 102. The axial projection 108 can be between -10% and 60 % of the axial chord "y" vary. The axial overlap 110 extends from the front blade edge tip 81 of the blades 80 in a downstream direction, thereby substantially overlapping part of the blades 80. The axial overlap 110 can vary between -10% and 100% of the axial chord “y”. In one embodiment, the radial height 112 of each of the plurality of recesses 86 is about 5 to 50% of the span “x” of the blades 80.
    As best illustrated in FIG. 7, the axial projection 108 extends upstream of the blades 80 and more particularly extends with a front blade tip 81 of the blades 80 coincident with the front wall 102. The axial overlap 110 extends in a downstream direction from the front blade tip 81 of the blades 80, thereby substantially overlapping a portion of the blades 80.
    As best illustrated in FIG. 8, in another embodiment of a recess 132, as illustrated on the left blade 80, the end wall device may be located entirely upstream of the front blade edge tip 81. More particularly, if the recess 132 includes an axial projection 108 that extends upstream of the front blade edge tip and a negative overlap 110 relative to the front blade edge tip 81. In this case, the end wall device, and more particularly the recess 132, has the task of correcting the portion 58 of the flow 50 near the housing 82 before the flow 58 enters the vane passage (as described herein).
    As best illustrated in FIG. 8, and particularly at recess 132 illustrated on right blade 80, end wall assembly may be located entirely downstream of front blade edge tip 81. More particularly, if the recess 132 includes an axial overlap 110 that extends downstream of the front blade edge tip 81 and a negative projection 108 relative to the front blade edge tip 81. In this case, the end wall device, and more particularly the recess 96, has the task of withdrawing weak leakage flows and, more particularly, part 58 of the flow 50 near a trailing blade edge 117 and strengthening the flow near the leading blade edge 116.
    Referring more specifically to Figures 9 and 10, a radial passage cross-sectional view illustrates a vane passage 134 (only one of which is illustrated) that is between adjacent blades 80 and more specifically between a suction side 136 of a first blade 138 and one Pressure side 140 of an adjacent positioned second blade 142 is defined. In one embodiment, the circumferential spacing of the plurality of recesses 132 around the housing 82 is approximately 0 to 10 recesses per blade passage 134, as best illustrated in FIGS. 9 and 10, but may vary for each blade passage 134. It should also be noted that in other embodiments, some vane passages may not contain recesses, while other vane passages may contain recesses.
    [0066] As illustrated in FIGS. 9 and 10, each of the plurality of recesses 132 is further defined by a first side wall 144 and a second side wall 146. Substantially similar to the first axial tilt angle α1 and the second axial tilt angle α2, the first side wall 144 and the second side wall 146 of each of the plurality of recesses 132 are inclined at an angle relative to a first tangential tilt angle β1 and a second tangential tilt angle β2 a peripheral surface of the compressor end wall of the housing 82. It is understood that similar tangential angles of inclination can define the recesses 132 when formed into a hub (as described above). In one embodiment, each of the first tangential tilt angle β1 and the second tangential tilt angle β2 ranges between 10 and 170 degrees relative to the peripheral surface 83 of the housing 82. In one embodiment, the tangential tilt angle 148 of the first side wall 144 and second side wall 146 may be the same . In one embodiment, the first tangential inclination angle β1 and a second tangential inclination angle β2 cannot be the same and can be designed independently of one another. When designing the tangential inclination angle, the tangential inclination angle β1 of the first side wall 144 is determined such that the leakage flows 74 are effectively removed. The tangential angle of inclination β2 of the second side wall 146 is determined in order to minimize the mixing loss in the main flow 50. As best illustrated in FIG. 9, each of the axial recesses 132 includes a geometric shape that has an overall curved shape from the first side 144 to the second side wall 146. Proper selection of the curvature can minimize aerodynamic losses in the recesses 132 and, more particularly, minimize energy loss near the sidewalls that meet at angles that are in the recesses 132. In another embodiment, each of the axial recesses 132 includes a geometric shape that has an overall linear shape from the first side 144 to the second side wall 146, as best illustrated in FIG. 10.
    The embodiments disclosed in FIGS. 7 to 10 contain one or more end wall devices in the form of the plurality of axial recesses 132. In one embodiment, each of the axial recesses 132 contains a geometric shape that extends from the front wall 102 to the rear wall 104 as a whole linear shape (FIG. 7) and from the first side wall 133 to the second side wall 146 has an overall linear shape (FIG. 10). In another embodiment, each of the axial recesses 132 includes a geometric shape that is generally linear in shape from the front wall 102 to the rear wall 104 (FIG. 7) and a generally curvilinear shape from the first side wall 144 to the second side wall 146 (FIG. 9) has. In yet another embodiment, each of the axial recesses 132 has a geometric shape, that from the front wall 102 to the rear wall 104 has an overall curvilinear shape (Figures 4 through 6) and from the first side wall 144 to the second side wall 146 an overall linear shape (Fig. 10). In yet another embodiment, each of the axial recesses 132 includes a geometric shape that is generally curvilinear from the front wall 102 to the rear wall 104 (FIGS. 4-6) and from the first side wall 144 to the second side wall 146 an overall curvilinear shape (FIG . 9) has. Some of the aspects that can be optimized include: (i) the axial angle of inclination α1 of the front wall 102 and the axial angle of inclination α2 of the rear wall 104 of the recesses 132; (ii) the tangential angle of inclination β1 of the first side wall 144 and the tangential angle of inclination β2 of the second side wall 146; (iii) radial height 112 of recesses 132; (iv) a length of axial overhang 108 and the length of axial overlap 110; (v) a tangential distance between the recesses 132 and within each recess 132 (as described herein), (vi) a number of recesses 132 spaced circumferentially from the end wall; (viii) an overall geometric cross-section of each recess when viewed in a radial-axial plane; and (viii) any variation of the above parameters in the radial, axial and tangential directions.
    Referring again to FIGS. 9 and 10, a portion of the recess surface can be defined as the non-metal surface 135 of the recess relative to the blade passage surface 134. In one embodiment, the proportion of the non-metal surface 135 of the recess is between 10% and 90% of the blade passage surface 134 and can vary in the radial direction. That is, the circumferential cover of each recess 132 may vary in the radial direction. By varying the circumferential cover in the radial direction, it is possible to minimize aerodynamic losses within the cutouts 132.
    Referring now to Fig. 11, in an exemplary graphical representation, generally designated 150, the benefit of a compressor incorporating the one or more end wall devices 94 as disclosed herein is illustrated, and more particularly when applied to a modern axial compressor rotor according to an exemplary embodiment. More specifically, the graph 150 illustrates the ratio of total to static pressure (drawn on axis 152) with the inlet corrected flow (drawn on axis 154) of a compressor without end wall devices, and in particular housing devices (drawn on line 156), of a compressor with one first end wall device and in particular a first housing device (drawn on a line 158) according to an embodiment described herein and a compressor with a second end wall device and in particular a second housing device (drawn on a line 160) according to an embodiment described here. As indicated by line 158, rotor 158 may continue to provide a pressure increase at a lower mass flow rate compared to a compressor that does not include end wall devices, as drawn on line 156. This extended stable operating range is only representative and can be optimized so that it is specific for a desired application. Furthermore, these results were obtained using a simulation of unsteady flow with numerical flow dynamics (CFD). A detailed study of the flow simulation results also confirms the primary flow mechanism. As already mentioned, the benefit of expanding the stable operating range and the impact on rotor efficiency depends on how the recess is designed relative to the rotor tip.
    Accordingly, as disclosed herein and illustrated in Figures 1-11, various technological advantages and / or improvements over existing compressor end wall devices, and particularly end wall devices, are provided that provide for an increase in the stall boundary area without the negative efficiency loss in a compressor. The proposed axial recesses, as disclosed herein, arranged circumferentially around an end wall of the compressor have the potential to provide larger stall boundary areas and a larger operability area of the compressor. The parameters of the axial recesses can be optimized and adapted for the application in which they are used.
    The proposed compressor end wall directions can also enable an increase in the performance of the gas turbine on hot days, a lower dependency on adjustable guide vanes during start-up, an increase in the performance of the rotor at the end-of-life gap dimensions and a lower dependence on transient blow-off valves in aircraft compressors during ice formation situations .
    Embodiments of an axial compressor end wall device and a method for controlling a leakage flow thereon are described in detail above. While the end wall assemblies have been described with reference to an axial compressor, the end wall assemblies described above can be used in any axial flow system, including other types of machinery including a compressor, and particularly those where an increase in stall boundary area is desired. Other applications will be apparent to those skilled in the art. Accordingly, the axial compressor end wall device and method of controlling leakage flow as disclosed herein are not limited to use with the specified machine device described herein. Furthermore, the present disclosure is not limited to the embodiments of the axial compressor described in detail above. Rather, other variations of the axial, mixing, and radial compressors that include embodiments of end wall devices can be used within the scope and scope of the claims.
    This written description uses examples to disclose the subject matter of the disclosure, including the best mode for carrying it out, and also to enable any person skilled in the art to practice the disclosure, including the making and using of any device or system, and the implementation of integrated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such further examples are intended to fall within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they contain equivalent structural elements with insignificant differences from the literal language of the claims.
    [0074] While shown and described herein what is currently considered the preferred embodiments of the disclosure, it will be apparent to those skilled in the art that various changes and modifications can be made therein without departing from the scope of the disclosure as defined by the appended claims is defined to deviate.
    An axial compressor for a gas turbine includes one or more end wall devices for controlling a leakage flow in the compressor. The one or more end wall devices have a height that is formed and arranged in an inner surface of a compressor housing or a compressor hub in order to return a flow adjacent to a plurality of blade tips or a plurality of guide blade tips to a cylindrical flow passage upstream from a tapping point of the flow. The end wall devices each define a front wall, a rear wall, an outer wall that runs between the front wall and the rear wall, an axial projection, an axial overlap, an axial angle of inclination and a tangential angle of inclination. The axial projection extends upstream to project at least over one of the at least one blade set and the at least one guide blade set. The axial overlap is downstream to overlap at least one of the at least one blade set and the at least one guide blade set.
    Parts list:
    10 aircraft engine assembly 12 central axis 13 radial axis 14 outer stationary annular blower housing 15 inner wall of the housing 16 blower assembly 17 outer wall of the housing 19 18 auxiliary compressor 19 auxiliary compressor housing 20 core gas turbine engine 21 low pressure turbine 22 airfoil 23 span 24 blower rotor blades 26 blower rotor disk 28 high pressure strut blade 30 from structural airfoil blading element 30 Combustion chamber 34 high pressure turbine 35 additional compressor rotor blades in 36 several rotor blades 37 additional compressor rotor disk 38 compressor rotor disk 40 first drive shaft 41 second drive shaft 42 43 bearing 44 suction side 45 blower frame 46 exhaust side of the core engine 47 rear turbine frame 48 blower exhaust side 49 blower inlet 50 part of the air flow 51 blower channel flow device 52 part 53 58 Part of the flow 50 59 Blown-in flow 60 Compressor 62 Blades 63 Blades elspitze 64 compressor housing 66 compressor rotor disk 68 guide blades 69 guide blade tip 70 several stages 72 gap 73 gap 74 76 rotor blade set 78 guide blade set 80 rotor blades 81 blade tip 82 compressor housing 83 inner surface of 82 84 compressor rotor disc 85 compressor hub 86 guide blades 87 guide blade tip 88 several stages 89 inner surface of 92 94 One or more end wall devices 96 Several axial recesses 98 Flow circulation / recirculation 100 Rear segment 102 Front wall 104 Rear wall 106 Outer wall 108 Axial projection 110 Axial overlap 112 Height 113 114 115 116 Blade leading edge 117 Blade trailing edge 118 Rear edge of the stator 120 compressor 122 124 126 128 130 compressor 132 Axial cutouts 133 Direction arrow 134 Blade passage 135 Non-metal surface 136 Suction side 138 First blade 140 Pressure side 141 142 Second blade 144 First side wall 145 146 Second side wall 1 47 148 Tangential inclination angle
    权利要求:
    Claims (9)
    [1]
    1. Compressor (30), which has:a compressor end wall (82, 85) defining a substantially cylindrical flow passage (56), the compressor end wall having a compressor housing (82) and a compressor hub (85) concentrically and coaxially arranged along a longitudinal central axis (12);at least one blade set (76), each of the at least one blade set (76) having a plurality of blades (80) connected to the compressor hub (85) and extending between the compressor hub (85) and the compressor housing (82) and there define a blade passage (134) between all of the blades (80); the compressor housing (82) surrounding the at least one blade set (76) to define an annular gap (90) between the compressor housing (82) and a plurality of blade tips (81) of the plurality of blades (80);at least one guide vane set (78), each of the at least one guide vane set (78) having a plurality of guide vanes (86) which are connected to the compressor housing (82) and extend between the compressor housing (82) and the compressor hub (85) and there defining a vane passage (134) between all of the guide vanes (86), the plurality of guide vanes (86) being positioned relative to the compressor hub (85) around an annular gap (92) between the compressor hub (85) and a plurality of guide vane tips (87) define a plurality of vanes (86); andone or more end wall devices (94) with a radial height (112), which in an inner surface (83, 89) of the housing (82) and / or the hub (95). are formed, the one or more end wall means (94) being arranged to direct a flow (58) adjacent to the plurality of blade tips (81) or the plurality of guide blade tips (87) to the cylindrical flow passage (56) upstream of a point of withdrawal of the flow (58), wherein the one or more end wall devices (94) each have a front wall (102) which contains a first axial inclination angle α1 relative to the longitudinal central axis (12), a rear wall (104) which has a second axial inclination angle α2 relative to the longitudinal central axis (12) includes an outer wall (106) that extends between the front wall (102) and the rear wall (104), an axial projection (108) that extends upstream to extend over at least one of the at least one blade set (76 ) and the at least one vane set (78) have an axial overlap (110) extending downstream by at least one of the at least one to overlap a rotor blade set (76) and the at least one guide blade set (78) (a first and a second side wall (144, 146) inclined at an angle) a first tangential inclination angle β1 relative to a circumferential surface (83, 89) of the Define the compressor end wall (82, 85) or a second tangential inclination angle β2 relative to the circumferential surface (83, 89) of the compressor end wall (82, 85), wherein either the axial inclination angle α1 is not equal to the axial inclination angle α2 or the tangential inclination angle β1 is not equal to the tangential Tilt angle β2 is.
    [2]
    2. The compressor (30) of claim 1, wherein the one or more end wall devices (94) have a plurality of separate axial recesses (96) defined circumferentially around the compressor hub (85) and / or the compressor housing (82); each vane passage (134) preferably containing 0 to 10 separate axial recesses (96).
    [3]
    3. A compressor (30) according to claim 1 or 2, wherein the one or more end wall means (94) have a radial height ranging from 5 to 50% of a span of the plurality of blades (80) and / or the plurality of vanes (86) .
    [4]
    The compressor (30) according to any one of the preceding claims, wherein the first axial inclination angle α1 and the second axial inclination angle α2 are in a range from 10 to 170 degrees; and / or wherein the first tangential inclination angle β1 and the second tangential inclination angle β2 are in a range from 10 to 170 degrees.
    [5]
    5. Compressor (30) according to one of the preceding claims, wherein the first axial inclination angle α1, the second axial inclination angle α2, the first tangential inclination angle β1 and the second tangential inclination angle β2 are not the same.
    [6]
    6. Compressor (30) according to one of the preceding claims, wherein the axial projection (108) is -10 to 60% of a blade chord length; wherein the axial projection (108) is preferably 0% of a blade chord length.
    [7]
    7. Compressor (30) according to any one of the preceding claims, wherein the axial overlap (110) is -10 to 60% of a blade chord length; wherein the axial overlap (110) is preferably 0% of a blade chord length.
    [8]
    8. Compressor according to one of the preceding claims, wherein a non-metal area of the recess is 10% to 90% of an area of the blade passage (134).
    [9]
    9. Engine (10), which has:a fan assembly (16);a core engine (20) downstream of the fan assembly (16), the core engine (20) including:a compressor (30);a combustion chamber (32) anda turbine (34), the compressor (30), the combustor (32) and the turbine (34) being arranged in a downstream axial flow relationship.wherein the compressor (30) further comprises:a compressor end wall (82, 85) defining a substantially cylindrical flow passage (56), the compressor end wall (82, 85) having a compressor housing (82) and a compressor hub (85) concentric about and coaxial along a longitudinal central axis ( 12) are arranged;at least one blade set (76), each of the at least one blade set (76) having a plurality of blades (80) connected to the compressor hub (85) and extending between the compressor hub (85) and the compressor housing (82), wherein the compressor housing (82) surrounds the at least one blade set (76) to define an annular gap (90) between the compressor housing (82) and a plurality of blade tips (81) of the plurality of blades (80);each of the at least one vane set (78) having a plurality of vane vanes (86) connected to the compressor housing (82) and extending between the compressor housing (82) and the compressor hub (85), the at least one vane set (78) is arranged relative to the compressor hub (85) to define an annular gap (92) between the compressor hub (85) and a plurality of guide vane tips (87) of the plurality of guide vanes; andone or more end wall devices (94) having a radial height (112) formed in an inner surface (83, 89) of the compressor housing (82) and / or the compressor hub (95), the one or more end wall devices (94) being set up are to return a flow (58) adjacent to the plurality of blade tips (81) to the cylindrical flow passage (56) upstream of a point of withdrawal of the flow (58), the one or more end wall devices (94) each having a front wall (102), which has a first axial inclination angle α1 relative to the longitudinal central axis (12), a rear wall (104) which has a second axial inclination angle α2 relative to the longitudinal central axis (12), an outer wall (106) which lies between the front wall (102) and the Extends rear wall (104), an axial projection (108) which extends upstream to at least over one of the at least one blade set (76) and the at least one guide blade set ( 78), an axial overlap (110) extending downstream to overlap at least one of the at least one blade set (76) and the at least one guide blade set (78) (first and second side walls (144, 146 ), inclined at an angle by a) first tangential inclination angle β1 relative to a circumferential surface (83, 89) of the compressor end wall (82, 85) or a second tangential inclination angle β2 relative to the circumferential surface (83, 89) of the compressor end wall (82, 85) to be defined, wherein the axial inclination angle α1 is not equal to the axial inclination angle α2 and / or the tangential inclination angle β1 is not equal to the tangential inclination angle β2.
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    同族专利:
    公开号 | 公开日
    US20160153465A1|2016-06-02|
    JP2016109124A|2016-06-20|
    CH710476A2|2016-06-15|
    CN205349788U|2016-06-29|
    DE102015120127A1|2016-06-02|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2019-05-31| NV| New agent|Representative=s name: FREIGUTPARTNERS IP LAW FIRM DR. ROLF DITTMANN, CH |
    2021-06-30| PL| Patent ceased|
    优先权:
    申请号 | 申请日 | 专利标题
    US14/556,452|US20160153465A1|2014-12-01|2014-12-01|Axial compressor endwall treatment for controlling leakage flow therein|
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