Patent ID: 12221898

DETAILED DESCRIPTION

Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,FIG.1depicts an exemplary gas turbine engine10. While the illustrated example is a high-bypass turbofan engine, the principles of the present invention are also applicable to other types of engines, such as low-bypass turbofans, turbojets, turboprops, etc. The engine10has a longitudinal center line or axis11and a stationary core casing12disposed concentrically about and coaxially along the axis11.

It is noted that, as used herein, the terms “axial” and “longitudinal” both refer to a direction parallel to the centerline axis11, 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” inFIG.1. These directional terms are used merely for convenience in description and do not require a particular orientation of the structures described thereby.

The engine10has a fan14, booster16, compressor18, combustor20, high pressure turbine or “HPT”22, and low pressure turbine or “LPT”24arranged in serial flow relationship. In operation, pressurized air from the compressor18is mixed with fuel in the combustor20and ignited, thereby generating combustion gases. Some work is extracted from these gases by the high pressure turbine22which drives the compressor18via an outer shaft26. The combustion gases then flow into the low pressure turbine24, which drives the fan14and booster16via an inner shaft28. The inner and outer shafts28and26are rotatably mounted in bearings30which are themselves mounted in a fan frame32and a turbine rear frame34.

FIGS.2-5illustrate a portion of an exemplary turbine rotor36suitable for inclusion in the HPT22or the LPT24. While the concepts of the present invention will be described using the HPT22as an example, it will be understood that those concepts are applicable to any of the turbines in a gas turbine engine. As used herein, the term “turbine” refers to turbomachinery elements in which kinetic energy of a fluid flow is converted to rotary motion.

The rotor36includes a disk38including an annular flowpath surface40extending between a forward end42and an aft end44. An array of turbine blades46extend from the flowpath surface40. The turbine blades46constitute “turbine airfoils” for the purposes of this invention. Each turbine blade46extends from a root48at the flowpath surface40to a tip50, and includes a concave pressure side52joined to a convex suction side54at a leading edge56and a trailing edge58. The adjacent turbine blades46define spaces60therebetween.

The turbine blades46are uniformly spaced apart around the periphery of the flowpath surface40. A mean circumferential spacing “s” (seeFIG.2) between adjacent turbine blades46is defined as s=2πr/Z, where “r” is a designated radius of the turbine blades46(for example at the root48) and “Z” is the number of turbine blades46. A nondimensional parameter called “solidity” is defined as c/s, where “c” is equal to the blade chord, described in detail below. In the illustrated example, the turbine blades46may 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.

As best seen inFIG.4, each turbine blade46has a span (or span dimension) “S1” defined as the radial distance from the root48to the tip50. Depending on the specific design of the turbine blade46, its span S1may be different at different axial locations.

For reference purposes a relevant measurement is the span S1at the leading edge56. Each turbine blade46has a chord (or chord dimension) “C1” (FIG.3) defined as the length of an imaginary straight line connecting the leading edge56and the trailing edge58. Depending on the specific design of the turbine blade46, its chord C1may be different at different locations along the span S1. For purposes of the present invention, the relevant measurement is the chord C1at the root48, i.e. adjacent the flowpath surface40.

Each turbine blade46has a thickness “T1” defined as the distance between the pressure side52and the suction side54(seeFIG.3). A “thickness ratio” of the turbine blade46is defined as the maximum value of the thickness T1, divided by the chord length, expressed as a percentage.

An array of splitter blades146(FIG.2) extend from the flowpath surface40. The splitter blades constitute “splitter airfoils” for the purposes of this invention. One or more splitter blades146may be disposed in each of the spaces60between the turbine blades46. In the circumferential direction, the splitter blade or blades146may be spaced uniformly or non-uniformly between two adjacent turbine blades46. Each splitter blade146extends from a root148at the flowpath surface40to a tip150, and includes a concave pressure side152joined to a convex suction side154at a leading edge156and a trailing edge158. In the example shown inFIG.2, the splitter blades146are positioned so that their trailing edges158are at approximately the same axial position as the trailing edges58of the turbine blades46; however the axial position of the splitter blades46may be varied to suit a particular application.

As best seen inFIG.5, each splitter blade146has a span (or span dimension) “S2” defined as the radial distance from the root148to the tip150. Depending on the specific design of the splitter blade146, its span S2may be different at different axial locations. For reference purposes a relevant measurement is the span S2at the leading edge156. Each splitter blade146has a chord (or chord dimension) “C2” defined as the length of an imaginary straight line connecting the leading edge156and the trailing edge158. Depending on the specific design of the splitter blade146, its chord C2may be different at different locations along the span S2. For purposes of the present invention, the relevant measurement is the chord C2at the root148, i.e. adjacent the flowpath surface40. Each splitter blade146has a thickness “T2” (FIG.3) defined as the distance between the pressure side152and the suction side154. A “thickness ratio” of the splitter blade146is defined as the maximum value of the thickness T2, divided by the chord length, expressed as a percentage.

The splitter blades146function to locally increase the hub solidity of the rotor36and thereby control undesired secondary flow around the turbine blades46. A similar effect could be obtained by simply increasing the number of turbine blades46, and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing flow blockage and aerodynamic frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter blades146and their position may be selected to control secondary flow while minimizing their surface area.

The thickness of the splitter blades146should be as small as possible consistent with structural, thermal, and aeroelastic considerations. Generally the splitter blades146should have a thickness ratio less than a thickness ratio of the turbine blades46. As one example, the splitter blades146may have a thickness ratio of less than about 5%. As another example, the splitter blades146may have a thickness ratio of about 2%. For comparison purposes, this is substantially less than the thickness of the turbine blades46. For example, the turbine blades46may be about 30% to 40% thick. Other turbine blades within the engine10, such as in the LPT24, may be about 5% to 10% thick.

The span S2and/or the chord C2of the splitter blades146may be equal to the corresponding span S1and chord C1of the turbine blades46. Alternatively, the span S2and/or the chord C2of the splitter blades146may be some fraction less than unity of the corresponding span S1and chord C1of the turbine blades46. These may be referred to as “part-span” and/or “part-chord” splitter blades. For example, the span S2may be equal to or less than the span S1. Preferably for reducing frictional losses, the span S2is 50% or less of the span S1. As another example, the chord C2may be equal to or less than the chord C1. Preferably for the least frictional losses, the chord C2is 50% or less of the chord C1.

The disk38, turbine blades46, and splitter blades146may be constructed from any material capable of withstanding the anticipated stresses and environmental conditions in operation. Non-limiting examples of known suitable alloys include nickel- and cobalt-based alloys.

The operational environment may exceed the temperature capability of metal alloys. Accordingly the turbine blades46may be actively cooled, in accordance with conventional practice, by providing them with a flow of coolant (such as compressor bleed air). The coolant is routed into internal passages of the turbine blades46and used for various forms of cooling such as conduction cooling, impingement cooling, and/or film cooling. As the turbine blades46generally have a significant thickness ratio, internal volume is available to incorporate active cooling features.

Because it is desirable to make the splitter blades146as thin as possible, there may not be internal volume available for active cooling features. Yet, metal alloys may not have sufficient high-temperature capability without active cooling.

This situation may be addressed by manufacturing all or part of the splitter blades146from nonmetallic high-temperature capable materials, such as ceramics, more particularly ceramic matrix composites (“CMC”). CMC is low density and tolerates high temperatures. Generally, commercially available CMC materials include a ceramic type fiber for example SiC, forms of which are coated with a compliant material such as Boron Nitride (BN). The fibers are carried in a ceramic-type matrix, one form of which is Silicon Carbide (SiC). CMC materials are often capable of operating in high-temperature gas environments without active cooling.

Optionally, all or part of the turbine blades46or disk38could be manufactured from the above-noted high-temperature materials.

InFIGS.2-5, the disk38, turbine blades46, and splitter blades146are depicted as an assembly built up from separate components. The principles of the present invention are equally applicable to a rotor with airfoils configured as an integral, unitary, or monolithic whole. This type of structure may be referred to as a “bladed disk” or “blisk”.

The splitter concepts described above may also be incorporated into turbine stator elements within the engine10. For example,FIGS.6-9illustrate a portion of a turbine nozzle62suitable for inclusion in the HPT22or the LPT24.

The turbine nozzle62includes a row of airflow-shaped turbine vanes64bounded at inboard and outboard ends, respectively by an inner band66and an outer band68. The turbine vanes64constitute “stator airfoils” for the purposes of this invention.

The inner band66defines an annular inner flowpath surface70extending between forward and aft ends72,74. The outer band68defines an annular outer flowpath surface76extending between forward and aft ends78,80. Each turbine vane46extends from a root82at the inner flowpath surface70to a tip84at the outer flowpath surface76, and includes a concave pressure side86joined to a convex suction side88at a leading edge90and a trailing edge92. The adjacent turbine vanes46define spaces92therebetween.

The turbine vanes64are uniformly spaced apart around the periphery of the inner flowpath surface70. The turbine vanes64have a mean circumferential spacing “s” and a solidity defined as described above (seeFIG.6). In the illustrated example, the turbine vanes64may have a spacing which is significantly greater than a spacing that would be expected in the prior art, resulting in a vane solidity significantly less than would be expected in the prior art.

As best seen inFIG.8, each turbine vane64has a span (or span dimension) “S3” defined as the radial distance from the root82to the tip84. Depending on the specific design of the turbine vane64, its span S3may be different at different axial locations. For reference purposes a relevant measurement is the span S3at the leading edge90. Each turbine vane64has a chord (or chord dimension) “C3” defined as the length of an imaginary straight line connecting the leading edge90and the trailing edge92. Depending on the specific design of the turbine vane64, its chord C3may be different at different locations along the span S3. For purposes of the present invention, the relevant measurement would be the chord C3at the root82or tip84, i.e. adjacent flowpath surfaces70or76.

Each turbine vane64has a thickness “T3” defined as the distance between the pressure side86and the suction side88A “thickness ratio” of the turbine vane64is defined as the maximum value of the thickness T3, divided by the chord length, expressed as a percentage.

One or both of the inner and outer flowpath surfaces70,76may be provided with an array of splitter vanes. In the example shown inFIG.6, an array of splitter vanes164extend radially inward from the outer flowpath surface76. The splitter vanes constitute “splitter airfoils” for the purposes of this invention. One or more splitter vanes164are disposed between each pair of turbine vanes64. In the circumferential direction, the splitter vane or vanes164may be spaced uniformly or non-uniformly between two adjacent turbine vanes64. Each splitter vane164extends from a tip184at the outer flowpath surface76to a root182, and includes a concave pressure side186joined to a convex suction side188at a leading edge190and a trailing edge192. In the example shown inFIGS.6and7, the splitter vanes164are positioned so that their trailing edges192are at approximately the same axial position as the trailing edges92of the stator vanes64; however the axial position of the splitter vanes164may be varied to suit a particular application.

As best seen inFIG.9, each splitter vane164has a span (or span dimension) “S4” defined as the radial distance from the root182to the tip184, and a chord (or chord dimension) “C4” defined as the length of an imaginary straight line connecting the leading edge190and the trailing edge192. Depending on the specific design of the splitter vane164, its chord C4may be different at different locations along the span S4. For purposes of the present invention, the relevant measurement is the chord C4at the tip184, i.e. adjacent flowpath surface76. Each splitter vane164has a thickness “T4” defined as the distance between the pressure side186and the suction side188. A “thickness ratio” of the splitter vane164is defined as the maximum value of the thickness T2, divided by the chord length, expressed as a percentage.

The splitter vanes164function to locally increase the solidity of the nozzle and thereby prevent the above-mentioned secondary flows. A similar effect could be obtained by simply increasing the number of turbine vanes64, and therefore reducing the vane-to-vane spacing. This, however, has the undesirable side effect of increasing flow blockage and aerodynamic frictional losses which would manifest as reduced aerodynamic efficiency and increased nozzle weight. Therefore, the dimensions of the splitter vanes164and their position may be selected to prevent secondary flows while minimizing their surface area.

The thickness of the splitter vanes164should be as small as possible consistent with structural, thermal, and aeroelastic considerations. Generally the splitter vanes164should have a thickness ratio less than a thickness ratio of the turbine vane64. As one example, the splitter vanes164may have a thickness ratio of less than about 5%. As another example, the splitter vanes164may have a thickness ratio on the order of about 2%.

The span S4and/or the chord S4of the splitter vanes164may be equal to the corresponding span S3and chord C3of the turbine vanes64. Alternatively, the span S4and/or the chord C4of the splitter vanes164may be some fraction less than unity of the corresponding span S3and chord C3of the turbine vanes64. These may be referred to as “part-span” and/or “part-chord” splitter vanes. For example, the span S4may be equal to or less than the span S3. Preferably for reducing frictional losses, the span S4is 50% or less of the span S3. As another example, the chord C4may be equal to or less than the chord C3. Preferably for the least frictional losses, the chord C4is 50% or less of the chord C3.

All or part of the splitter vanes164may comprise high-temperature capable materials such as the CMC materials discussed above.

FIG.10illustrates an array of splitter vanes264extending radially outward from the inner flowpath surface70. Other than the fact that they extend from the inner flowpath surface70, the splitter vanes264may be identical to the splitter vanes164described above, in terms of their shape, circumferential position relative to the stator vanes64, their thickness, span, and chord dimensions, and their material composition. As noted above, splitter vanes may optionally be incorporated at the inner flowpath surface70, or the outer flowpath surface76, or both.

The turbine apparatus described herein incorporating splitter blades and/or splitter vanes increases the endwall solidity level locally, to locally increase solidity in regions of high secondary flow without incurring the penalty from profile loss due to surface area in regions outside the region of interest.

The foregoing has described a turbine 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 to 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.