Patent Abstract:
A compressor apparatus includes: a rotor having: a disk mounted for rotation about a centerline axis, an outer periphery of the disk defining a flowpath surface; an array of airfoil-shaped axial-flow compressor blades extending radially outward from the flowpath surface, wherein the compressor blades each have a root, a tip, a leading edge, and a trailing edge, wherein the compressor blades have a chord dimension and are spaced apart by a circumferential spacing, the ratio of the chord to the circumferential spacing defining a blade solidity parameter; and an array of airfoil-shaped splitter blades alternating with the compressor blades, wherein the splitter blades each have a root, a tip, a leading edge, and a trailing edge; wherein at least one of a chord dimension of the splitter blades at the roots thereof and a span dimension of the splitter blades is less than the corresponding dimension of the compressor blades.

Full Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to turbomachinery compressors and more particularly relates to rotor blade stages of such compressors. 
         [0002]    A gas turbine engine includes, in serial flow communication, a compressor, a combustor, and turbine. The turbine is mechanically coupled to the compressor and the three components define a turbomachinery core. The core is operable in a known manner to generate a flow of hot, pressurized combustion gases to operate the engine as well as perform useful work such as providing propulsive thrust or mechanical work. One common type of compressor is an axial-flow compressor with multiple rotor stages each including a disk with a row of axial-flow airfoils, referred to as compressor blades. 
         [0003]    For reasons of thermodynamic cycle efficiency, it is generally desirable to incorporate a compressor having the highest possible pressure ratio (that is, the ratio of inlet pressure to outlet pressure). It is also desirable to include the fewest number of compressor stages. However, there are well-known inter-related aerodynamic limits to the maximum pressure ratio and mass flow possible through a given compressor stage. 
         [0004]    It is known to reduce weight, improve rotor performance, and simplify manufacturing by minimizing the total number of compressor airfoils used in a given rotor blade row . However, as airfoil blade count is reduced the accompanying reduced hub solidity tends to cause the airflow in the hub region of the rotor airfoil to undesirably separate from the airfoil surface. 
         [0005]    It is also known to configure the disk with a non-axisymmetric “scalloped” surface profile to reduce mechanical stresses in the disk. An aerodynamically adverse side effect of this feature is to increase the rotor blade row through flow area and aerodynamic loading level promoting airflow separation. 
         [0006]    Accordingly, there remains a need for a compressor rotor that is operable with sufficient stall range and an acceptable balance of aerodynamic and structural performance. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0007]    This need is addressed by the present invention, which provides an axial compressor having a rotor blade row including compressor blades and splitter blade airfoils. 
         [0008]    According to one aspect of the invention, a compressor apparatus includes: a rotor including: a disk mounted for rotation about a centerline axis, an outer periphery of the disk defining a flowpath surface; an array of airfoil-shaped axial-flow compressor blades extending radially outward from the flowpath surface, wherein the compressor blades each have a root, a tip, a leading edge, and a trailing edge, wherein the compressor blades have a chord dimension and are spaced apart by a circumferential spacing, the ratio of the chord dimension to the circumferential spacing defining a blade solidity parameter; and an array of airfoil-shaped splitter blades alternating with the compressor blades, wherein the splitter blades each have a root, a tip, a leading edge, and a trailing edge; wherein at least one of a chord dimension of the splitter blades at the roots thereof and a span dimension of the splitter blades is less than the corresponding dimension of the compressor blades. 
         [0009]    According to another aspect of the invention, the solidity parameter is selected to as to result in hub flow separation under normal operating conditions. 
         [0010]    According to another aspect of the invention, the flowpath surface is not a body of revolution. 
         [0011]    According to another aspect of the invention, the flowpath surface includes a concave scallop between adjacent compressor blades. 
         [0012]    According to another aspect of the invention, the scallop has a minimum radial depth adjacent the roots of the compressor blades, and has a maximum radial depth at a position approximately midway between adjacent compressor blades. 
         [0013]    According to another aspect of the invention, each splitter blade is located approximately midway between two adjacent compressor blades. 
         [0014]    According to another aspect of the invention, the splitter blades are positioned such that their trailing edges are at approximately the same axial position as the trailing edges of the compressor blades, relative to the disk. 
         [0015]    According to another aspect of the invention, the span dimension of the splitter blades is 50% or less of the span dimension of the compressor blades. 
         [0016]    According to another aspect of the invention, the span dimension of the splitter blades is 30% or less of the span dimension of the compressor blades. 
         [0017]    According to another aspect of the invention, the chord dimension of the splitter blades at the roots thereof is 50% or less of the chord dimension of the compressor blades at the roots thereof. 
         [0018]    According to another aspect of the invention, the chord dimension of the splitter blades at the roots thereof is 50% or less of the chord dimension of the compressor blades at the roots thereof. 
         [0019]    According to another aspect of the invention, a compressor includes a plurality of axial-flow stages, at least a selected one of the stages includes: a disk mounted for rotation about a centerline axis, an outer periphery of the disk defining a flowpath surface; an array of airfoil-shaped axial-flow compressor blades extending radially outward from the flowpath surface, wherein the compressor blades each have a root, a tip, a leading edge, and a trailing edge, wherein the compressor blades have a chord dimension and are spaced apart by a circumferential spacing, the ratio of the chord dimension to the circumferential spacing defining a blade solidity parameter; and an array of airfoil-shaped splitter blades alternating with the compressor blades, wherein the splitter blades each have a root, a tip, a leading edge, and a trailing edge; wherein at least one of a chord dimension of the splitter blades at the roots thereof and a span dimension of the splitter blades is less than the corresponding dimension of the compressor blades. 
         [0020]    According to another aspect of the invention, the solidity parameter is selected to as to result in hub flow separation under normal operating conditions. 
         [0021]    According to another aspect of the invention, the flowpath surface is not a body of revolution. 
         [0022]    According to another aspect of the invention, the flowpath surface includes a concave scallop between adjacent compressor blades. 
         [0023]    According to another aspect of the invention, the span dimension of the splitter blades is 50% or less of the span dimension of the compressor blades. 
         [0024]    According to another aspect of the invention, the span dimension of the splitter blades is 30% or less of the span dimension of the compressor blades. 
         [0025]    According to another aspect of the invention, the chord dimension of the splitter blades at the roots thereof is 50% or less of the chord dimension of the compressor blades at the roots thereof. 
         [0026]    According to another aspect of the invention, the chord dimension of the splitter blades at the roots thereof is 50% or less of the chord dimension of the compressor blades at the roots thereof 
         [0027]    According to another aspect of the invention, the selected stage is the aft-most rotor of the compressor. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0028]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0029]      FIG. 1  is a cross-sectional, schematic view of a gas turbine engine that incorporates a compressor rotor apparatus constructed in accordance with an aspect of the present invention; 
           [0030]      FIG. 2  is a perspective view of a portion of a rotor of a compressor apparatus; 
           [0031]      FIG. 3  is a top plan view of a portion of a rotor of a compressor apparatus; 
           [0032]      FIG. 4  is an aft elevation view of a portion of a rotor of a compressor apparatus; 
           [0033]      FIG. 5  is a side view taken along lines  5 - 5  of  FIG. 4 ; 
           [0034]      FIG. 6  is a side view taken along lines  6 - 6  of  FIG. 4 ; 
           [0035]      FIG. 7  is a perspective view of a portion of a rotor of an alternative compressor apparatus; 
           [0036]      FIG. 8  is a top plan view of a portion of a rotor of an alternative compressor apparatus; 
           [0037]      FIG. 9  is an aft elevation view of a portion of a rotor of an alternative compressor apparatus; 
           [0038]      FIG. 10  is a side view taken along lines  10 - 10  of  FIG. 9 ; and 
           [0039]      FIG. 11  is a side view taken along lines  11 - 11  of  FIG. 9 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0040]    Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates a gas turbine engine, generally designated  10 . The engine  10  has a longitudinal centerline axis  11  and includes, in axial flow sequence, a fan  12 , a low-pressure compressor or “booster”  14 , a high-pressure compressor (“HPC”)  16 , a combustor  18 , a high-pressure turbine (“HPT”)  20 , and a low-pressure turbine (“LPT”)  22 . Collectively, the HPC  16 , combustor  18 , and HPT  20  define a core  24  of the engine  10 . The HPT  20  and the HPC  16  are interconnected by an outer shaft  26 . Collectively, the fan  12 , booster  14 , and LPT  22  define a low-pressure system of the engine  10 . The fan  12 , booster  14 , and LPT  22  are interconnected by an inner shaft  28 . 
         [0041]    In operation, pressurized air from the HPC  16  is mixed with fuel in the combustor  18  and burned, generating combustion gases. Some work is extracted from these gases by the HPT  20  which drives the compressor  16  via the outer shaft  26 . The remainder of the combustion gases are discharged from the core  24  into the LPT  22 . The LPT  22  extracts work from the combustion gases and drives the fan  12  and booster  14  through the inner shaft  28 . The fan  12  operates to generate a pressurized fan flow of air. A first portion of the fan flow (“core flow”) enters the booster  14  and core  24 , and a second portion of the fan flow (“bypass flow”) is discharged through a bypass duct  30  surrounding the core  24 . While the illustrated example is a high-bypass turbofan engine, the principles of the present invention are equally applicable to other types of engines such as low-bypass turbofans, turbojets, and turboshafts. 
         [0042]    It is noted that, as used herein, the terms “axial” and “longitudinal” both refer to a direction parallel to the centerline axis  11 , while “radial” refers to a direction perpendicular to the axial direction, and “tangential” or “circumferential” refers to a direction mutually perpendicular to the axial and tangential directions. As used herein, the terms “forward” or “front” refer to a location relatively upstream in an air flow passing through or around a component, and the terms “aft” or “rear” refer to a location relatively downstream in an air flow passing through or around a component. The direction of this flow is shown by the arrow “F” in  FIG. 1 . These directional terms are used merely for convenience in description and do not require a particular orientation of the structures described thereby. 
         [0043]    The HPC  16  is configured for axial fluid flow, that is, fluid flow generally parallel to the centerline axis  11 . This is in contrast to a centrifugal compressor or mixed-flow compressor. The HPC  16  includes a number of stages, each of which includes a rotor comprising a row of airfoils or blades  32  (generically) mounted to a rotating disk  34 , and row of stationary airfoils or vanes  36 . The vanes  36  serve to turn the airflow exiting an upstream row of blades  32  before it enters the downstream row of blades  32 . 
         [0044]      FIGS. 2-6  illustrate a portion of a rotor  38  constructed according to a first exemplary embodiment of the present invention and suitable for inclusion in the HPC  16 . As an example, the rotor  38  may be incorporated into one or more of the stages in the aft half of the HPC  16 , particularly the last or aft-most stage. 
         [0045]    The rotor  38  includes a disk  40  with a web  42  and a rim  44 . It will be understood that the complete disk  40  is an annular structure mounted for rotation about the centerline axis  11 . The rim  44  has a forward end  46  and an aft end  48 . An annular flowpath surface  50  extends between the forward and aft ends  46 ,  48 . 
         [0046]    An array of compressor blades  52  extend from the flowpath surface  50 . Each compressor blade extends from a root  54  at the flowpath surface  50  to a tip  56 , and includes a concave pressure side  58  joined to a convex suction side  60  at a leading edge  62  and a trailing edge  64 . As best seen in  FIG. 5 , each compressor blade  52  has a span (or span dimension) “S 1 ” defined as the radial distance from the root  54  to the tip  56 , and a chord (or chord dimension) “C 1 ” defined as the length of an imaginary straight line connecting the leading edge  62  and the trailing edge  64 . Depending on the specific design of the compressor blade  52 , its chord C 1  may be different at different locations along the span S 1 . For purposes of the present invention, the relevant measurement is the chord C 1  at the root  54 . 
         [0047]    As seen in  FIG. 4 , the flowpath surface  50  is not a body of revolution. Rather, the flowpath surface  50  has a non-axisymmetric surface profile. As an example of a non-axisymmetric surface profile, it may be contoured with a concave curve or “scallop”  66  between each adjacent pair of compressor blades  52 . For comparison purposes, the dashed lines in  FIG. 4  illustrate a hypothetical cylindrical surface with a radius passing through the roots  54  of the compressor blades  52 . It can be seen that the flowpath surface curvature has its maximum radius (or minimum radial depth of the scallop  66 ) at the compressor blade roots  54 , and has its minimum radius (or maximum radial depth “d” of the scallop  66 ) at a position approximately midway between adjacent compressor blades  52 . 
         [0048]    In steady state or transient operation, this scalloped configuration is effective to reduce the magnitude of mechanical and thermal hoop stress concentration at the airfoil hub intersections on the rim  44  along the flowpath surface  50 . This contributes to the goal of achieving acceptably-long component life of the disk  40 . An aerodynamically adverse side effect of scalloping the flowpath  50  is to increase the rotor passage flow area between adjacent compressor blades  52 . This increase in rotor passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side  60  of the compressor blade  52 , at the inboard portion near the root  54 , and at an aft location, for example approximately 75% of the chord distance C 1  from the leading edge  62 . 
         [0049]    An array of splitter blades  152  extend from the flowpath surface  50 . One splitter blade  152  is disposed between each pair of compressor blades  52 . In the circumferential direction, the splitter blades  152  may be located halfway or circumferentially biased between two adjacent compressor blades  52 , or circumferentially aligned with the deepest portion d of the scallop  66 . Stated another way, the compressor blades  52  and splitter blades  152  alternate around the periphery of the flowpath surface  50 . Each splitter blade  152  extends from a root  154  at the flowpath surface  50  to a tip  156 , and includes a concave pressure side  158  joined to a convex suction side  160  at a leading edge  162  and a trailing edge  164 . As best seen in  FIG. 6 , each splitter blade  152  has a span (or span dimension) “S 2 ” defined as the radial distance from the root  154  to the tip  156 , and a chord (or chord dimension) “C 2 ” defined as the length of an imaginary straight line connecting the leading edge  162  and the trailing edge  164 . Depending on the specific design of the splitter blade  152 , its chord C 2  may be different at different locations along the span S 2 . For purposes of the present invention, the relevant measurement is the chord C 2  at the root  154 . 
         [0050]    The splitter blades  152  function to locally increase the hub solidity of the rotor  38  and thereby prevent the above-mentioned flow separation from the compressor blades  52 . A similar effect could be obtained by simply increasing the number of compressor blades  152 , and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter blades  152  and their position may be selected to prevent flow separation while minimizing their surface area. The splitter blades  152  are positioned so that their trailing edges  164  are at approximately the same axial position as the trailing edges of the compressor blades  52 , relative to the rim  44 . This can be seen in  FIG. 3 . The span S 2  and/or the chord C 2  of the splitter blades  152  may be some fraction less than unity of the corresponding span S 1  and chord C 1  of the compressor blades  52 . These may be referred to as “part-span” and/or “part-chord” splitter blades. For example, the span S 2  may be equal to or less than the span S 1 . Preferably for reducing frictional losses, the span S 2  is 50% or less of the span S 1 . More preferably for the least frictional losses, the span S 2  is 30% or less of the span S 1 . As another example, the chord C 2  may be equal to or less than the chord C 1 . Preferably for the least frictional losses, the chord C 2  is 50% or less of the chord C 1 . 
         [0051]    The disk  40 , compressor blades  52 , and splitter blades  152  may be constructed from any material capable of withstanding the anticipated stresses and environmental conditions in operation. Non-limiting examples of known suitable alloys include iron, nickel, and titanium alloys. In  FIGS. 2-6  the disk  40 , compressor blades  52 , and splitter blades  152  are depicted as an integral, unitary, or monolithic whole. This type of structure may be referred to as a “bladed disk” or “blisk”. The principles of the present invention are equally applicable to a rotor built up from separate components (not shown). 
         [0052]      FIGS. 7-11  illustrate a portion of a rotor  238  constructed according to a second exemplary embodiment of the present invention and suitable for inclusion in the HPC  16 . As an example, the rotor  238  may be incorporated into one or more of the stages in the aft half of the HPC  16 , particularly the last or aft-most stage. 
         [0053]    The rotor  238  includes a disk  240  with a web  242  and a rim  244 . It will be understood that the complete disk  240  is an annular structure mounted for rotation about the centerline axis  11 . The rim  244  has a forward end  246  and an aft end  248 . An annular flowpath surface  250  extends between the forward and aft ends  246 ,  248 . 
         [0054]    An array of compressor blades  252  extend from the flowpath surface  250 . Each compressor blade  252  extends from a root  254  at the flowpath surface  250  to a tip  256 , and includes a concave pressure side  258  joined to a convex suction side  260  at a leading edge  262  and a trailing edge  264 . As best seen in  FIG. 10 , each compressor blade  252  has a span (or span dimension) “S 3 ” defined as the radial distance from the root  254  to the tip  256 , and a chord (or chord dimension) “C 3 ” defined as the length of an imaginary straight line connecting the leading edge  262  and the trailing edge  264 . Depending on the specific design of the compressor blade  252 , its chord C 3  may be different at different locations along the span S 3 . For purposes of the present invention, the relevant measurement is the chord C 3  at the root  254 . 
         [0055]    The compressor blades  252  are uniformly spaced apart around the periphery of the flowpath surface  250 . A mean circumferential spacing “s” (see  FIG. 9 ) between adjacent compressor blades  252  is defined as s=2πr/Z, where “r” is a designated radius of the compressor blades  252  (for example at the root  254 ) and “Z” is the number of compressor blades  252 . A nondimensional parameter called “blade solidity” is defined as c/s, where “c” is equal to the blade chord as described above. In the illustrated example, the compressor blades  252  may have a spacing which is significantly greater than a spacing that would be expected in the prior art, resulting in a blade solidity significantly less than would be expected in the prior art. 
         [0056]    As seen in  FIG. 9 , the flowpath surface  250  is depicted as a body of revolution (i.e. axisymmetric). Optionally, the flowpath surface  250  may have a non-axisymmetric surface profile as described above for the flowpath surface  250 . 
         [0057]    The reduced blade solidity will have the effect of reducing weight, improving rotor performance, and simplify manufacturing by minimizing the total number of compressor airfoils used in a given rotor stage. An aerodynamically adverse side effect of reduced blade solidity is to increase the rotor passage flow area between adjacent compressor blades  252 . This increase in rotor passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side  260  of the compressor blade  252 , at the inboard portion near the root  254 , and at an aft location, for example approximately 75% of the chord distance C 3  from the leading edge  262 , also referred to as “hub flow separation”. For any given rotor design, the compressor blade spacing may be intentionally selected to produce a solidity low enough to result in hub flow separation under expected operating conditions. 
         [0058]    An array of splitter blades  352  extend from the flowpath surface  250 . One splitter blade  352  is disposed between each pair of compressor blades  252 . In the circumferential direction, the splitter blades  352  may be located halfway or circumferentially biased between two adjacent compressor blades  252 . Stated another way, the compressor blades  252  and splitter blades  352  alternate around the periphery of the flowpath surface  250 . Each splitter blade  352  extends from a root  354  at the flowpath surface  250  to a tip  356 , and includes a concave pressure side  358  joined to a convex suction side  360  at a leading edge  362  and a trailing edge  364 . As best seen in  FIG. 11 , each splitter blade  352  has a span (or span dimension) “S 4 ” defined as the radial distance from the root  354  to the tip  356 , and a chord (or chord dimension) “C 4 ” defined as the length of an imaginary straight line connecting the leading edge  362  and the trailing edge  364 . Depending on the specific design of the splitter blade  352 , its chord C 4  may be different at different locations along the span S 4 . For purposes of the present invention, the relevant measurement is the chord C 4  at the root  354 . 
         [0059]    The splitter blades  352  function to locally increase the hub solidity of the rotor  238  and thereby prevent the above-mentioned flow separation from the compressor blades  252 . A similar effect could be obtained by simply increasing the number of compressor blades  252 , and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter blades  352  and their position may be selected to prevent flow separation while minimizing their surface area. The splitter blades  352  are positioned so that their trailing edges  364  are at approximately the same axial position as the trailing edges  264  of the compressor blades  252 , relative to the rim  244 . This can be seen in  FIG. 8 . The span S 4  and/or the chord C 4  of the splitter blades  352  may be some fraction less than unity of the corresponding span S 3  and chord C 3  of the compressor blades  252 . These may be referred to as “part-span” and/or “part-chord” splitter blades. For example, the span S 4  may be equal to or less than the span S 3 . Preferably for reducing frictional losses, the span S 4  is 50% or less of the span S 3 . More preferably for the least frictional losses, the span S 4  is 30% or less of the span S 3 . As another example, the chord C 4  may be equal to or less than the chord C 3 . Preferably for the least frictional losses, the chord C 4  is 50% or less of the chord C 3 . 
         [0060]    The disk  240 , compressor blades  252 , and splitter blades  352  using the same materials and structural configuration (e.g. monolithic or separable) as the disk  40 , compressor blades  52 , and splitter blades  152  described above. 
         [0061]    The rotor apparatus described herein with splitter blades increases the rotor hub solidity level locally, reduces the hub aerodynamic loading level locally, and suppresses the tendency of the rotor airfoil hub to want to separate in the presence of the non-axisymmetric contoured hub flowpath surface, or with a reduced airfoil count rotor on an axisymmetric flowpath. The use of a partial-span and/or partial-chord splitter blade is effective to keep the solidity levels of the middle and upper sections of the rotor unchanged from a nominal value, and therefore to maintain middle and upper airfoil section performance. 
         [0062]    The foregoing has described a compressor rotor apparatus. All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive. 
         [0063]    Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features. 
         [0064]    The invention is not restricted to the details of the foregoing embodiment(s). The invention extends any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.

Technology Classification (CPC): 5