Patent Application
20170114796
This invention relates generally to turbomachinery compressors and more particularly relates to rotor and stator airfoils of such compressors.
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 stages each including a rotating disk with a row of axial-flow airfoils, referred to as rotor blades. Typically, this type of compressor also includes stationary airfoils alternating with the rotor airfoils, referred to as stator vanes. The stator vanes are typically bounded at their inner and outer ends by arcuate endwall structures (e.g. a hub or a case).
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.
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 or stator row. However, as airfoil count is reduced the accompanying reduced endwall solidity tends to cause the airflow in the endwall region of the airfoil to undesirably separate from the airfoil surface.
Accordingly, there remains a need for a compressor that is operable with sufficient stall range and an acceptable balance of aerodynamic and structural performance.
This need is addressed by an axial compressor having a stator vane row including stator vane airfoils and splitter airfoils.
The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:
Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,
It will be understood that the fan 12 and booster 14 are driven by a low-pressure turbine (“LPT”) which is not illustrated in
It will be further understood that the HPC 16 is driven by a high-pressure turbine (“HPT”) which is not illustrated in
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.
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
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 mounted to a rotating disk, and row of stationary airfoils or vanes. The vanes serve to turn the airflow exiting an upstream row of blades before it enters the downstream row of blades.
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.
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
As seen in
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 C1 from the leading edge 62.
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
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
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
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.
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
The compressor blades 252 are uniformly spaced apart around the periphery of the flowpath surface 250. A mean circumferential spacing “s” (see
As seen in
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 C3 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.
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
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
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.
Several portions of the engine 10 shown in
The inner band 444 defines an annular inner flowpath surface 450 extending between forward and aft ends 446, 448. The casing 470 defines an annular outer flowpath surface 472 extending between forward and aft ends 474, 476.
The stator vanes 452 extend between the inner and outer flowpath surfaces 450, 472. Each stator vane 452 extends from a root 454 at the inner flowpath surface 450 to a tip 456 at the outer flowpath surface 472, and includes a concave pressure side 458 joined to a convex suction side 460 at a leading edge 462 and a trailing edge 464. As best seen in
The stator vanes 452 are uniformly spaced apart around the periphery of the inner flowpath surface 450. The stator vanes 452 have a mean circumferential spacing “s”, defined as described above (see
As seen in
The reduced vane solidity will have the effect of reducing weight, improving stator performance, and simplify manufacturing by minimizing the total number of airfoils used in a given stator stage. An aerodynamically adverse side effect of reduced stator solidity is to increase the rotor passage flow area between adjacent stator vanes 452. This increase in stator passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side 460 of the stator vane 452, at the inboard portion near the root 454, and at an aft location, for example approximately 75% of the chord distance C5 from the leading edge 462, also referred to as “hub flow separation”. It also tends to cause undesirable flow separation on the suction side 460 of the stator vane 452, at the outboard portion near the tip 456, and at an aft location, for example approximately 75% of the chord distance C5 from the leading edge 462, also referred to as “case flow separation”. Generally, both of these conditions may be referred to as “endwall separation”. For any given stator design, the stator vane spacing may be intentionally selected to produce a solidity low enough to result in endwall separation under expected operating conditions.
To counter this adverse side effect, one or both of the inner and outer flowpath surfaces 450, 472 may be provided with an array of splitter vanes. In the example shown in
The splitter vanes 552 function to locally increase the hub solidity of the stator and thereby prevent the above-mentioned flow separation from the stator vanes 452. A similar effect could be obtained by simply increasing the number of stator vanes 452, and therefore reducing the vane-to-vane spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased stator weight. Therefore, the dimensions of the splitter vanes 552 and their position may be selected to prevent flow separation while minimizing their surface area. The splitter vanes 552 are positioned so that their trailing edges 564 are at approximately the same axial position as the trailing edges 464 of the stator vanes 452, relative to the outer flowpath surface 472. This can be seen in
The compressor apparatus described herein with splitter blades and/or splitter vanes increases the endwall solidity level locally, reduces the endwall aerodynamic loading level locally, and suppresses the tendency of the airfoil portion adjacent the endwall to want to separate in the presence of the non-axisymmetric contoured endwall flowpath surface, or with a reduced airfoil count on an axisymmetric flowpath. The use of a partial-span and/or partial-chord splitter blade or vane is effective to keep the solidity levels of the middle and upper sections of the airfoil unchanged from a nominal value, and therefore to maintain middle and upper airfoil section performance.
The foregoing has described a compressor 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.
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.
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.
