Gas turbine engine turbine diaphragm with angled holes

A gas turbine engine turbine diaphragm (460) includes an inner cylindrical portion (461), a mounting portion (463), and a disk portion (462). The mounting portion (463) is located radially outward from the inner cylindrical portion (461). The disk portion (462) extends radially between the inner cylindrical portion (461) and the mounting portion (463). The disk portion (462) includes a plurality of angled holes (464). Each angled hole (464) follows a vector which is angled in at least one plane. A component of the vector is located on a plane perpendicular to a radial extending from an axis of the diaphragm (460). The component of the vector is angled relative to an axial direction of the diaphragm (460).

TECHNICAL FIELD

The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a turbine diaphragm with angled holes configured for cooling downstream components.

BACKGROUND

Gas turbine engines include compressor, combustor, and turbine sections. Portions of a gas turbine engine are subject to high temperatures. In particular, the turbine section is subject to such high temperatures that the first stages are cooled by air directed through internal cooling passages from the compressor. The use of air from the compressor for cooling may reduce the efficiency of the gas turbine engine. Loss or uncontrolled cooling air leakage may also lead to a loss of efficiency and may lead to improper cooling.

U.S. patent Application Publication No. 2011-0274536 to A. Inomata discloses a steam turbine where a diaphragm-side cooling path is formed through the internal diaphragm in the axial direction of the rotor and a cooling medium flowing through the rotor-side cooling path diverts into the diaphragm-side cooling path and a labyrinth flow path provided between the internal diaphragm and the rotor.

The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.

SUMMARY OF THE DISCLOSURE

A gas turbine engine turbine diaphragm includes an inner cylindrical portion, a mounting portion, and a disk portion. The mounting portion is located radially outward from the inner cylindrical portion. The disk portion extends radially between the inner cylindrical portion and the mounting portion. The disk portion includes a plurality of angled holes. Each angled hole follows a vector which is angled in at least one plane. A component of the vector is located on a plane perpendicular to a radial extending from an axis of the diaphragm. The component of the vector is angled relative to an axial direction of the diaphragm.

DETAILED DESCRIPTION

The systems and methods disclosed herein include a gas turbine engine diaphragm with angled holes. In embodiments, the diaphragm may be configured to provide a predictable amount of cooling air to the adjacent turbine disk located aft of the diaphragm. The angled holes can be configured to swirl cooling air such that the angular velocity of the cooling air matches the angular velocity of the adjacent turbine disk. Matching the angular velocity of the cooling air with the angular velocity of the adjacent turbine disk can reduce the temperature of the adjacent turbine disk, which can result in a longer service life of the adjacent turbine disk.

FIG. 1is a schematic illustration of an exemplary gas turbine engine. Some of the surfaces have been left out or exaggerated (here and in other figures) for clarity and ease of explanation. Also, the disclosure may reference a forward and an aft direction. Generally, all references to “forward” and “aft” are associated with the flow direction of primary air (i.e., air used in the combustion process), unless specified otherwise. For example, forward is “upstream” relative to primary air flow, and aft is “downstream” relative to primary air flow.

In addition, the disclosure may generally reference a center axis95of rotation of the gas turbine engine, which may be generally defined by the longitudinal axis of its shaft120(supported by a plurality of bearing assemblies150). The center axis95may be common to or shared with various other engine concentric components. All references to radial, axial, and circumferential directions and measures refer to center axis95, unless specified otherwise, and terms such as “inner” and “outer” generally indicate a lesser or greater radial distance from, wherein a radial96may be in any direction perpendicular and radiating outward from center axis95.

A gas turbine engine100includes an inlet110, a shaft120, a gas producer or “compressor”200, a combustor300, a turbine400, an exhaust500, and a power output coupling600. The gas turbine engine100may have a single shaft or a dual shaft configuration.

The compressor200includes a compressor rotor assembly210and compressor stationary vanes (“stators”)250. The compressor rotor assembly210mechanically couples to shaft120. As illustrated, the compressor rotor assembly210is an axial flow rotor assembly. The compressor rotor assembly210includes one or more compressor disk assemblies220. Each compressor disk assembly220includes a compressor rotor disk that is circumferentially populated with compressor rotor blades. Stators250axially precede each of the compressor disk assemblies220. Each compressor disk assembly220paired with the adjacent stators250that precede the compressor disk assembly220is considered a compressor stage. Compressor200includes multiple compressor stages.

The combustor300includes one or more injectors350and includes one or more combustion chambers390.

The turbine400includes a turbine rotor assembly410, turbine nozzles450, and one or more turbine diaphragms460. The turbine rotor assembly410mechanically couples to the shaft120. As illustrated, the turbine rotor assembly410is an axial flow rotor assembly. The turbine rotor assembly410includes one or more turbine disk assemblies420. Each turbine disk assembly420includes a turbine disk430(shown inFIG. 2) that is circumferentially populated with turbine blades440(shown inFIG. 2). Turbine nozzles450axially precede each of the turbine disk assemblies420. The turbine diaphragm460may support turbine nozzles450and may be located radially inward from turbine nozzles450. Each turbine disk assembly420paired with the adjacent turbine nozzles450that precede the turbine disk assembly420is considered a turbine stage. Turbine400includes multiple turbine stages.

The exhaust500includes art exhaust diffuser520and an exhaust collector550.

FIG. 2is a cross-sectional view of a portion of the turbine400ofFIG. 1.FIG. 3is a perspective view of a turbine diaphragm460. The diaphragm ofFIG. 3may be used in the gas turbine engine100ofFIG. 1. As previously mentioned and illustrated inFIG. 2, the turbine400includes turbine diaphragm460, turbine nozzles450, and turbine rotor assembly410(shown inFIG. 1). All references to radial, axial, and circumferential directions and measures for elements of turbine diaphragm460refer to the axis of turbine diaphragm460, which is concentric to center axis95. The axial direction99(illustrated inFIG. 4) of the turbine diaphragm460is the direction traveling from the forward or upstream side of the turbine diaphragm460to the aft or downstream side of the turbine diaphragm460along a path concentric or parallel to the axis of the turbine diaphragm460.

Referring toFIGS. 2 and 3, turbine diaphragm460may include inner cylindrical portion461, disk portion462, and mounting portion463. Inner cylindrical portion461may be in the form of a hollow circular cylinder with a variable thickness, defining a bore there within. Mounting portion463may be a circular piece and may be located radially outward from inner cylindrical portion461. As shown inFIG. 2, mounting portion463may be located radially inward from turbine nozzles450and may be configured to couple with turbine nozzles450. Mounting portion463may include mounting holes467as illustrated inFIG. 3.

Disk portion462may extend radially between inner cylindrical portion461and mounting portion463. Disk portion462may also extend axially forward and axially aft while spanning radially between inner cylindrical portion461and mounting portion463. Disk portion462may also have a variable thickness. In the embodiment shown inFIG. 3, inner cylindrical portion461, disk portion462, and mounting portion463circumferentially extends completely around the axis of the diaphragm. Inner cylindrical portion461, mounting portion463, and disk portion462may be configured to form a first cavity465located axially forward of disk portion462as shown inFIG. 2.

Turbine diaphragm460is configured to include angled holes464, and may also include first stress relief region468(illustrated inFIGS. 3 and 4) and second stress relief region469(shown inFIG. 4). In the embodiment shown inFIGS. 2 and 3, angled holes464are formed in disk portion462. In one embodiment, turbine diaphragm460includes from five to twenty angled holes464. In the embodiment shown inFIG. 3, turbine diaphragm460is configured to include eleven angled holes464.

The diameter of angled holes464may be sized based on the cooling flow needed. In one embodiment, the diameter of each angled hole464taken at a cross-section normal to the angled hole464is ⅛″ to 3/16″. In another embodiment, the diameter of each angled hole464taken at a cross-section normal to the angled hole464is ¼″.

Referring now toFIG. 2, each turbine nozzle450includes an outer wall454, an inner wall455, and a nozzle blade451. Each outer wall454has an arcuate shape and connects to the turbine housing (not shown). An inner wall455is located radially inward from outer wall454. Each inner wall455has an arcuate shape and may connect to turbine diaphragm460at mounting portion463. One or more nozzle blades451span between outer wall454and inner wall455.

Turbine rotor assembly410includes multiple turbine disk assemblies420joined together. A turbine disk assembly420is axially forward of turbine diaphragm460and includes first turbine disk430with multiple turbine blades440. Another turbine disk assembly420is axially aft of turbine diaphragm460and includes second turbine disk435with multiple turbine blades440. First turbine disk430and second turbine disk435may be configured with a bore (not shown) for coupling to shaft120(shown inFIG. 1). First turbine disk430includes disk boles432. The first cavity465may be bound by an aft facing surface of first turbine disk430. Second turbine disk435may include a disk-post and dampers437(only one shown inFIG. 2). Dampers437are located near the radial outer edge of second disk435. A portion of dampers437may be on an axially forward facing surface of second turbine disk435. The axially forward facing surface of second turbine disk435and turbine diaphragm460define a second cavity466.

First turbine disk430may also include first labyrinth threads431extending axially all and radially outward. Second turbine disk435may include second labyrinth threads436extending axially forward and radially outward. The second labyrinth threads436may be located axially aft of the first labyrinth threads431. Both first labyrinth threads431and second labyrinth threads436may be located radially inward of turbine diaphragm460. Bore running surface439may be located radially inward of and radially adjacent to turbine diaphragm460and may be within the bore of turbine diaphragm460. In the embodiment shown inFIG. 2, first labyrinth threads431, second labyrinth threads436, and bore running surface439form a labyrinth seal within the bore of turbine diaphragm460.

Turbine blades440may be installed axially or circumferentially onto first turbine disk430and second turbine disk435. Turbine400also includes shrouds445located radially outward and spaced apart from turbine blades440. Shrouds445may attach to the turbine housing (not shown).

The turbine400may also include a forward diaphragm470, a preswirler (not shown), a forward labyrinth seal480, and an aft labyrinth seal490. The forward diaphragm470is located axially forward of first turbine disk430. Forward diaphragm470may also be configured to couple with turbine nozzles450. The preswirler may be located within a third cavity473formed in forward diaphragm470between radial outer portion471of forward diaphragm470and radial inner portion472of forward diaphragm470. The axially aft end of die third cavity473may be bound by the axially forward facing surface of first turbine disk430.

Forward labyrinth seal480may be located within third cavity473between forward diaphragm470and first turbine disk430. Forward labyrinth seal480may be coupled to first turbine disk430at the forward axial face of first turbine disk430. Forward labyrinth seal480includes forward outer labyrinth threads481, forward inner labyrinth threads482, forward labyrinth hole483, forward outer running surface488, and forward inner running surface489. Forward outer running surface488may be adjacent outer portion471and forward outer labyrinth threads481. Forward outer running surface488may be radially inward from outer portion471and radially outward from forward outer labyrinth threads481. Forward inner running surface489may be adjacent inner portion472and forward inner labyrinth threads482. Forward inner running surface489may be located radially outward from inner portion472and radially inward from forward inner labyrinth threads482.

Aft labyrinth seal490may be located within first cavity465between turbine diaphragm460and first turbine disk430. Aft labyrinth seal490may be coupled to first turbine disk430at the aft axial face of first turbine disk430. Aft labyrinth seal490includes aft outer labyrinth threads491, aft inner labyrinth threads492, aft labyrinth hole493, aft outer running surface498, and aft inner running surface499. Aft outer running surface498may be adjacent mounting portion463and aft outer labyrinth threads491. Aft outer running surface498may be radially inward from mounting portion463and radially outward from aft outer labyrinth threads491. Aft inner running surface499may be adjacent inner cylindrical portion461and aft inner labyrinth threads492. Aft inner running surface499may be located radially outward from cylindrical portion461and radially inward from aft inner labyrinth threads492.

In the embodiment depicted inFIG. 2, forward diaphragm470is a first stage diaphragm, first turbine disk430is a first stage turbine disk, turbine diaphragm460is a second stage diaphragm, and second turbine disk435is a second stage turbine disk.

FIG. 4is a cross-sectional view the turbine diaphragm460shown inFIG. 3. The turbine diaphragm460may be used in the gas turbine engine100ofFIG. 1. As previously mentioned, turbine diaphragm460may be configured to include angled holes464. Angled holes464may follow a vector which is angled in at least one plane. A component of this vector, line97, may be located on a plane perpendicular to a radial extending from the axis of the turbine diaphragm460and may be angled relative to the axial direction99of the turbine diaphragm460illustrated by angle98, the angle between lines94and97inFIG. 4. The radial may be radial96(illustrated inFIG. 1). Line94is a reference line located on the plane perpendicular to the radial extending from the axis of the diaphragm460and is oriented in axial direction99. Line97may also be angled towards the second turbine disk435rotational direction so that angled holes464may direct the gas or air in the same rotational direction as the second turbine disk435.

In one embodiment, angle98is from twenty to eighty-five degrees. In another embodiment, angle98is from fifty to seventy degrees. In another embodiment, angle98is sixty degrees.

A first stress relief region468may be formed contiguous each angled hole464. Each first stress relief region468is in flow communication with the angled hole464. Each first stress relief region468may be an elongated recess and may recede into turbine diaphragm460thereby widening the opening of the angled hole464. Each first stress relief region468may have an angle similar to the angle of the contiguous angled hole464. Each first stress relief region468may have a curved profile and may include multiple curves, arcs, or radii. Each first stress relief region468may be an elongated scoop. The scoop may be wider than the diameter of angled holes464. The elongated length of the scoop may be biased away from the contiguous angled hole464along line97, as illustrated inFIG. 4. In one embodiment, each first stress relief region468is located upstream of an angled hole464and recedes axially aft into turbine diaphragm460. In another embodiment, each first stress relief region468is located downstream of an angled hole464and recedes axially forward into turbine diaphragm460.

A second stress relief region469may be formed contiguous each angled hole464. Each second stress relief region469is in flow communication with the angled hole464. Each second stress relief region469may be an elongated recess and may recede into turbine diaphragm460thereby widening the opening of the angled hole464. Each second stress relief region469may have an angle similar to the angle of the contiguous angled hole464. Each second stress relief region469may have a curved profile and may include multiple curves, arcs, or radii. Each second stress relief region469may be an elongated scoop. The scoop may be wider than the diameter of angled holes464. The elongated length of the scoop may be biased away from the contiguous angled hole464along line97, as illustrated inFIG. 4. In one embodiment, each second stress relief region469is located downstream of an angled hole464opposite an upstream first stress relief region468. The second stress relief region469recedes axially forward into turbine diaphragm460. In another embodiment, each second stress relief region469is located upstream of an angled hole464opposite a downstream first stress relief region468. The second stress relief region469recedes axially aft into turbine diaphragm460.

Each first stress relief region468and second stress relief region469may be formed by manufacturing processes such as ball milling, electrical discharge machining, or drilling.

One or more of the above components (or their subcomponents) may be made from stainless steel and/or durable, high temperature materials known as “superalloys”. A superalloy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance. Superalloys may include materials such as HASTELLOY, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, INCOLOY, MP98T, TMS alloys, and CMSX single crystal alloys.

Industrial Applicability

Gas turbine engines may be suited for any number of industrial applications such as various aspects of the oil and gas industry (including transmission, gathering, storage, withdrawal, and lifting of oil and natural gas), the power generation industry, cogeneration, aerospace, and other transportation industries.

Referring toFIG. 1, a gas (typically air10) enters the inlet110as a “working fluid”, and is compressed by the compressor200. In the compressor200, the working fluid is compressed in an annular flow path115by the series of compressor disk assemblies220. In particular, the air10is compressed in numbered “stages”, the stages being associated with each compressor disk assembly220. For example, “4th stage air” may be associated with the 4th compressor disk assembly220in the downstream or “aft” direction going from the inlet110towards the exhaust500). Likewise, each turbine disk assembly420may be associated with a numbered stage.

Once compressed air10leaves the compressor200, it enters the combustor300, where it is diffused and fuel20is added. Air10and fuel20are injected into the combustion chamber390via injector350and ignited. After the combustion reaction, energy is then extracted from the combusted fuel/air mixture via the turbine400by each stage of the series of turbine disk assemblies420. Exhaust gas90may then be diffused in exhaust diffuser520and collected, redirected, and exit the system via an exhaust collector550. Exhaust gas90may also be further processed (e.g., to reduce harmful emissions, and/or to recover heat from the exhaust gas90).

Operating efficiency of a gas turbine engine generally increases with a higher combustion temperature. Thus, there is a trend in gas turbine engines to increase the temperatures. Gas reaching forward stages of a turbine from a combustion chamber may be 1000 degrees Fahrenheit or more. To operate at such high temperatures a portion of compressed air of a compressor of a gas turbine engine may be diverted through internal passages or chambers to cool various components of a turbine such as disk-posts, dampers, and turbine disks.

Gas reaching forward stages of a turbine may also be under high pressure. Cooling air diverted from a compressor may need to be at compressor discharge pressure to effectively cool turbine components located in forward stages of a turbine. Gas turbine engine100components such as second turbine disk435, damper437, and disk-posts (not shown) may be subject to elevated levels of stress.

Cooling air with a substantially axial flow is diverted from the compressor discharge. deferring toFIG. 2, the cooling air from the compressor discharge may pass through forward diaphragm470and a preswirler (not shown) to the path for cooling air54. Compressor discharge air may exit the preswirler with a tangential component that may match the angular velocity of first turbine disk430. Cooling air may travel along path for coo ling air54from third cavity473, through forward labyrinth hole483of forward labyrinth seal480, and into first turbine disk430and to path for cooling air55.

Path for cooling air55may pass axially through first turbine disk430along disk holes432. A portion of the cooling air may be diverted radially outward to cool turbine blades440that circumferentially surround first turbine disk430. The remainder of the cooling air may continue along path for cooling air55and exits disk holes432on the aft side of first turbine disk430to path for cooling air56. Path for cooling air56may pass through aft labyrinth hole493and into first cavity465. While a particular path along paths for cooling air54,55, and56has been described, alternate paths from the compressor discharge to first cavity465may be used.

It was determined through research and testing that cooling air from the compressor discharge may be directed to the second cavity466to cool the second turbine disk435, dampers437, and the disk-posts. Cooling air from the compressor discharge entering first cavity465may exit first cavity465and travel to second cavity466along path for cooling air59. A portion of the cooling air may also travel along path for cooling air57radially outward towards a gap between a radial outer edge of first turbine disk430and inner wall455.

Cooling air following path for cooling air59may pass through aft labyrinth seal490between aft inner labyrinth threads492and aft inner running surface499, as well as a labyrinth seal formed by first labyrinth threads431, second labyrinth threads436, and bore running surface439. As turbine400heats up or cools down the distance between aft inner labyrinth threads492and aft inner running surface499, as well as the distance between first labyrinth threads431and bore running surface439, and the distance between second labyrinth threads436and bore running surface439may increase or decrease due to thermal expansion. These variable distances may provide uncontrolled amounts of compressor discharge cooling air to second cavity466. An uncontrolled amount of cooling air may lead to improper or insufficient cooling of second turbine disk435, dampers437, and the disk-posts.

It was further determined that angled holes464may provide a controlled flow of cooling air to second cavity466to cool second turbine disk435, dampers437, and disk-posts. Cooling air traveling along path for cooling air59may be minimized by reducing the gaps between aft inner labyrinth threads492and aft inner running surface499, first labyrinth threads431and bore running surface439, and second labyrinth threads436and bore running surface439. Alternative seals reducing the flow of cooling air along path for cooling air59may also be used.

While a portion of the cooling air may travel along path for cooling air59, angled holes464may be configured such that a majority of the cooling air traveling from first cavity465to second cavity466may travel along path for cooling air58, which travels through turbine diaphragm460along angled holes464. With the use of angled holes464the amount of cooling air passing from first cavity465to second cavity466may be predicted. Angled holes464may not be as sensitive to thermal expansion as the labyrinth seals and may provide a more stable flow of cooling air to second cavity466. In one embodiment, fifty to one-hundred percent of the cooling air travels along path for cooling air58and zero to fifty percent of the cooling air travels along path for cooling air59. In another embodiment, fifty to seventy percent of the cooling air travels along path for cooling air58and thirty to fifty percent of the cooling air travels along path for cooling air59. In another embodiment, fifty-five to sixty-five percent of the cooling air travels along path for cooling air58and thirty-five to forty-five percent of the cooling air travels along path for cooling aft59. In another embodiment, approximately sixty-two percent of the cooling air travels along path tor cooling air58and approximately thirty-eight percent of the cooling air travels along path for cooling air59.

Path for cooling air58through angled holes464is much shorter and less tortuous than path for cooling air59through aft labyrinth seal490and the labyrinth seal formed by first labyrinth threads431, second labyrinth threads436, and bore running surface439. The longer, more tortuous path for cooling air59may result in an increase in temperature, as the cooling air may be in contact with hot gas turbine engine components for a longer time period prior to reaching second cavity466. A pressure drop in the cooling air may also occur due to the length and tortuous path of path for cooling air59. The temperature increase and pressure drop may reduce the effectiveness of the cooling air. Directing cooling air through path for cooling air58may result in more effective cooling and may result in an increase in gas turbine engine efficiency.

Angled holes464may be configured to direct cooling air into second cavity466with an angular velocity that matches the angular velocity of second turbine disk435. A matching angular velocity between the cooling air and second turbine disk435may reduce metal temperatures of the second turbine disk435, dampers437, and the disk-posts, which can result in extending the turbine field service life of the second disk435and dampers437.

Directing the cooling air with angled holes464may lead to increased stress in regions of turbine diaphragm460. It was determined that first stress relief region468and second stress relief region469may reduce stress concentrations in turbine diaphragm460. The use of angled holes464may lead to longer service life hours tor second turbine disk435, dampers437, and the disk-posts, as well as an efficient use of the cooling air bled from compressor200.

FIG. 5is a flowchart of a method for forming angled cooling holes in a turbine diaphragm. The method includes determining the amount of cooling air needed to cool components aft of the turbine diaphragm460at step610. The components aft of the turbine diaphragm460may include the second turbine disk435, dampers437, and the disk-posts. Step610is followed by sizing the radius for the holes to allow all or a portion of the determined amount of cooling air to pass through the holes at step620. Step620may be partially based on the allowable downstream disk stress requirements. The downstream disk may be second turbine disk435(shown inFIG. 2). Step620may be followed by selecting an angle for the holes at step630. The angle of the holes may be selected by starting with straight holes, the axis of each hole being in the axial direction99of the turbine diaphragm and skewing the axis of each hole toward the rotational direction as much as possible under the condition of satisfying all mechanical design requirements.

Step630is followed by sizing a labyrinth seal clearance to allow the determined amount of cooling air not passing through the angled holes to pass through the labyrinth seal at step640. The labyrinth seal clearance may be for aft labyrinth seal490and the labyrinth seal formed by first labyrinth threads431, second labyrinth threads436, and bore running surface439(shown inFIG. 2). Step640may be followed by performing an engine test validation at step645.

The method also includes forming holes in a turbine diaphragm with the selected radius and with the selected angle at step650. The holes formed at step650may be angled holes464. In one embodiment step650is completed by drilling.

It is understood that the steps disclosed herein (or parts thereof) may be performed in the order presented or out of the order presented, unless specified otherwise. For example, step620may be performed before, after, or concurrently with step630.

The preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. The described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. Hence, although the present disclosure, for convenience of explanation, depicts and describes particular diaphragms and associated processes, it will be appreciated that other diaphragms and processes in accordance with this disclosure can be implemented in various other turbine stages, configurations, and types of machines. Furthermore, there is no intention to be bound by any theory presented in the preceding background or detailed description. It is also understood that the illustrations may include exaggerated dimensions to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.