Patent Description:
A turbine is a mechanical device that obtains a rotational force by an impulsive force or reaction force using a flow of a compressible fluid such as steam or gas. A turbine may be realized, for example, as a steam turbine using steam or as a gas turbine using a high temperature combustion gas.

Among them, the gas turbine is mainly composed of a compressor, a combustor, and a turbine. The compressor is provided with an air inlet for introducing air, and a plurality of compressor vanes and compressor blades, which are alternately arranged in a compressor housing.

The combustor supplies fuel to the compressed air compressed in the compressor and ignites a fuel-air mixture using a burner to produce a high temperature and high pressure combustion gas.

The turbine has a plurality of turbine vanes and turbine blades disposed alternately in a turbine casing. Further, a rotor is arranged to pass through the center of the compressor, the combustor, the turbine and an exhaust chamber.

Both ends of the rotor are rotatably supported by bearings. A plurality of disks is fixed to the rotor and the respective blades are connected to the disks. A drive shaft may be connected to the rotor and to a generator, for example, at an end of the exhaust chamber.

Since these gas turbines have no reciprocating mechanism, such as a piston found in a <NUM>-stroke engine, and consequently have no frictional parts like piston-cylinder, they have several advantages. These include minimal consumption of lubricating oil, a significant reduction in amplitude, which is a characteristic of a reciprocating machine, and the ability to operate at high speed.

Briefly describing the operation of the gas turbine, the compressed air compressed in the compressor is mixed with fuel and combusted to produce a high-temperature combustion gas in the combustor, which is then injected toward the turbine. The injected combustion gas passes through the turbine vanes and the turbine blades to generate a rotational force, which causes the rotor to rotate.

The foregoing is intended merely to aid in the understanding of the background of the present invention, and is not intended to mean that the present invention falls within the purview of the related art that is already known to those skilled in the art.

Document <CIT> discloses a detachable moving blade fixing method comprising a step of embedding a key fittingly in advance in a recess provided at the top end part of a rotor disc, a step of inserting a moving blade in a disc groove, a step of screwing a screw in a threaded hole of the key, a step of pulling up the key by screwing the screw, and the key is fixed in order for engaging with both the recess at the disc side and the moving blade.

Document <CIT> discloses a turbine blade assembly. A turbine blade which is a stop blade is provided with a cutout portion which is formed in one circumferential side surface of a shank portion at the center in the axial direction of a turbine rotor, a cutout portion which is formed in one circumferential side surface of the shank portion from one end portion to the cutout portion in the axial direction of the turbine rotor, and a through passage which is formed to pass from the cutout portion to an effective blade part and in which a moving member, which moves a stopper member to the effective blade part in the cutout portion, is inserted.

Document <CIT> discloses a fixture for buckets for a turbine. The fixture includes a rotor wheel that includes a plurality of dovetail grooves. A platform seat is formed between the dovetail grooves and provided with a first insertion groove in a circumferential direction. A bucket includes a root that inserts into one of the dovetail grooves. A base platform disposed on an upper surface of the root includes a second insertion groove at a position facing the first insertion groove. A first fixture is inserted into an internal area defined by the first and second insertion grooves. A second fixture is inserted into the internal area defined by the first groove and the second insertion groove of a final bucket to fix the final bucket among buckets mounted on the rotor wheel. An auxiliary fixture is inserted into an opening hole formed in the base platform of the final bucket and facing the second fixture.

Document <CIT> <NUM> discloses an axial locking device of a bucket comprising a rotor rotationally installed on a casing, a plurality of dovetail grooves formed along an outer circumferential surface of the rotor to have a distance in a tangential direction of the rotor, a plurality of buckets provided with a wing part, a platform, and a dovetail, and to be fixated by inserting to the dovetail groove along an axial direction of the rotor. A fastening groove is formed by penetrating a portion of the platform, and a receiving part is formed on the rotor to correspond to a position of the fastening groove. Finally, a fastening block is inserted to the fastening groove and the receiving part.

Document <CIT> discloses a securing device for turbine rotor vane. The securing device is for improving the strength against the force in the axial direction of a vane wheel by inserting a stopper into a space between an insertion groove at a portion for planting a rotor vane and a notched groove in the outer circumference of the vane wheel than inserting a pin along the underface of the stopper and securing the stopper. The stopper is inserted into a space formed between an insertion groove made in a planting section of the last rotor vane and a notched groove made in the planting section of vane wheel. A pin is inserted into a pin hole in the axial direction of the vane wheel, then fitted in a groove of the stopper and the stopper is fitted in the insertion groove.

Document <CIT> discloses a system and method for attaching a rotating blade in a turbine. A system for attaching a rotating blade in a turbine includes a bush having an axial slot and a radial slot that intersects with the axial slot. A radial retention member fits within the radial slot, and an axial retention member fits within the axial slot and engages with the radial retention member. A method for attaching a rotating blade in a turbine includes a step of inserting a bush into an axial passage in a rotor wheel and inserting a radial retention member into a radial passage in the rotor wheel and through at least a portion of the bush. The method further includes a step of inserting the rotating blade in a slot in the rotor wheel, a step of inserting the radial retention member into a retention slot in the rotating blade, and a step of inserting an axial retention member into the bush.

Document <CIT> discloses a locking of side entry blades. In a turbine rotor axially extending grooves and airfoil blades supported in the grooved includes locking devices disposed in circumferential slots and keyways formed in the blade roots such that said locking devices extend to the adjacent blade root, whereby disengagement of said locking devices from the respective blade roots is prevented and at lest the locking device locking the last of said blades installed in the rotor consists of two parts with deformable portions, which overlap for installing the last blade and which are deformed so as to abut one another to prevent disengagement of said locking device.

Accordingly, the present invention has been made keeping in mind the above problems occurring in the related art, and an objective of the present invention is to provide a solution which is capable of fastening an axial position of a blade upon installation on a rotor disk and preventing the axial movement of the blade after the installation, and a gas turbine including the same.

To this end, the present invention provides a rotor assembly in accordance with claim <NUM>, a gas turbine in accordance with claim <NUM>, and a method in accordance with claim <NUM>.

According to a first aspect of the present invention, there is provided a rotor assembly. The rotor assembly includes a blade, a rotor disk including a disk slot into which a root member of the blade is inserted, and a blade fastening assembly. The blade fastening assembly comprises a fastening key having a through hole which may optionally include an internal thread; a first groove formed in the rotor disk with a shape corresponding to a portion of the fastening key, a second groove at least partially formed in a platform of the blade in communication with the first groove and forming a shape corresponding to a remaining portion of the fastening key in a state of being inserted into the disk slot, a pin groove connected with the first groove and formed in the rotor disk, and a fastening pin inserted into the through hole and the pin groove.

The first groove, generally, may extend along an axial direction, and the second groove may extend transverse to the first groove, in particular, along a circumferential direction.

The first groove may be formed in an outer circumference or an outer circumferential surface of the rotor disk. Similar, the disk slot may be formed in the outer circumference of the rotor disk and extend along axial direction.

The second groove may include a first portion formed in the rotor disk in communication with the first groove, and a second portion formed in the platform in communication with the first portion of the second groove and the first groove. The second portion is adjacent to the first portion when the root member of the blade is received in the disk slot. Since the first portion of the second groove is formed in the rotor disk, i.e., in the outer circumference thereof, the fastening key may be supported with respect to the axial direction on the surfaces of the first portion of the second groove and, thus, by the rotor disk.

The fastening key may include a body part in which the through hole is formed, and a wing part extending from the body part toward the platform.

The body part may have a first height and the wing part may have a second height smaller than the first height. The body part may have an upper surface and an opposite lower surface, and the first height may be measured from the upper to the lower surface. Similar, the wing part may have an upper surface and an opposite lower surface, and the second height may be measured from the upper to the lower surface of the wing part.

The wing part may extend toward the platform or, generally, protrude from a side surface of the body part. The side surface may extend between the upper surface and the lower surface of the body part. The lower surface of the wing part may be positioned spaced from the lower surface of the body part. Additionally, or alternatively, the upper surface of the body part and the upper surface of the wing part may extend flush with each other or form a continuous surface.

The wing part may be inserted into the second groove, e.g., through both the first portion of the second groove and the second portion of the second groove.

The fastening pin may be in a cylindrical shape.

The fastening pin may include a first insertion part having an end inserted into the pin groove and a second insertion part connected to the insertion part inserted into the through hole. The second insertion part may comprise an external thread on an outer circumferential surface thereof. The external thread may be engaged with an optional internal thread provided in the through hole.

The second insertion part may be provided on an end face thereof with a fastening structure, e.g., in the form of a polygonal recess. The fastening structure is generally configured for being coupled with a fastener to rotate the fastening pin.

In a second aspect of the present invention, there is provided a gas turbine including: a compressor configured to compress air; a combustor configured to mix the compressed air from the compressor with fuel and combust a compressed air-fuel mixture; a turbine section rotated by combustion gases from the combustor to generate power; and a rotor assembly of the first aspect of the invention. The rotor assembly may be formed in at least one of the compressor and the turbine section to fasten an axial position of a blade and prevent the axial movement of the blade. As described above, the blade fastening assembly of the rotor assembly may include: a fastening key having a threaded hole; a rotor disk including a disk slot into which a root member is inserted, the rotor disk having a first groove with a shape corresponding to a portion of the fastening key, a first portion of a second groove adjacent to the first groove, and a pin groove connected with the first groove; a platform having a second portion of the second groove formed in communication with the first portion of the second groove and the first groove and forming a shape corresponding to a remaining portion of the fastening key in a state of being inserted into the disk slot; and a fastening pin inserted into the threaded hole and the pin groove.

The fastening key may include a body part in which the threaded hole is formed, and a wing part extending from the body part toward the platform.

The body part may have a first height and the wing part may have a second height smaller than the first height.

The wing part may extend toward the platform from a side surface of the body part extending downwards the second height from an upper surface of the body part.

The wing part may be inserted through both the first portion of the second groove and the second portion of the second groove.

The fastening pin may include an insertion part having an end inserted into the pin groove and a threaded part connected to the insertion part and threaded on an outer circumferential surface thereof.

The threaded part may be provided on an upper surface thereof with a fastening member into which a fastener is inserted to rotate the fastening pin.

According to a third aspect of the invention, a method for assembling the rotor assembly of the first aspect includes inserting the root member of the blade into the disk slot of the rotor disk, inserting the fastening key is inserted into the first groove and the second groove, and inserting the fastening pin into the through hole and the pin groove.

Optionally, the method may further include fastening the fastening pin in at least one of the through hole and the pin groove, for example, by engaging the optional external thread provided on the fastening pin with the optional internal thread provided in the through hole and/or the fastening groove, as described above.

Details of other implementations of various aspects of the present invention will be described in the following detailed description.

According to the present invention, the installation of the blade on the rotor disk can be structurally safe and easily performed. Furthermore, after the installation, the axial movement of the blade can be prevented. Accordingly, wear caused by friction between the blades and the rotor disk can be prevented, thereby improving the durability of a gas turbine.

However, it should be noted that the present invention is not limited thereto, and may include all of modifications, equivalents or substitutions within the scope of the present invention as defined in the appended claims.

Terms used herein are used to merely describe specific embodiments, and are not intended to limit the present invention. As used herein, an element expressed as a singular form includes a plurality of elements, unless the context clearly indicates otherwise. Further, it will be understood that the term "comprising" or "including" specifies the presence of stated features, numbers, steps, operations, elements, parts, or combinations thereof, but does not preclude the presence or addition of one or more other features, numbers, steps, operations, elements, parts, or combinations thereof.

It is noted that like elements are denoted in the drawings by like reference symbols as whenever possible. Further, the detailed description of known functions and configurations that may obscure the gist of the present invention will be omitted. For the same reason, some of the elements in the drawings are exaggerated, omitted, or schematically illustrated.

<FIG> is a partially cut-away perspective view of a gas turbine, <FIG> is a cross-sectional view illustrating a schematic structure of a gas turbine, and <FIG> is a partial cross-sectional view illustrating an internal structure of the gas turbine.

As illustrated in <FIG>, a gas turbine <NUM> includes a compressor <NUM>, a combustor <NUM>, and a turbine <NUM>. The compressor <NUM> includes a plurality of blades <NUM> radially installed. The compressor <NUM> rotates the blade <NUM> so that air flows while being compressed by the rotation of the blade <NUM>. The size and installation angle of the blade <NUM> may vary depending on the installation location. In one embodiment, the compressor <NUM> is connected directly or indirectly to the turbine <NUM>, and receives a portion of the power generated from the turbine <NUM> to rotate the blade <NUM>.

Air compressed by the compressor <NUM> flows to the combustor <NUM>. The combustor <NUM> includes a plurality of combustion chambers <NUM> and a fuel nozzle module <NUM> arranged in an annular shape.

The gas turbine <NUM> includes a housing <NUM> and a diffuser <NUM> which is disposed on a rear side of the housing <NUM> and through which a combustion gas passing through a turbine is discharged. A combustor <NUM> is disposed in front of the diffuser <NUM> so as to receive and burn compressed air.

Referring to the flow direction of the air, a compressor <NUM> is located on the upstream side of the housing <NUM>, and a turbine <NUM> is located on the downstream side of the housing. A torque tube <NUM> is disposed as a torque transmission member between the compressor <NUM> and the turbine <NUM> to transmit the rotational torque generated in the turbine to the compressor.

The compressor <NUM> is provided with a plurality (for example, <NUM>) of compressor rotor disks <NUM>, which are fastened by a tie rod <NUM> to prevent axial separation thereof.

Specifically, the compressor rotor disks <NUM> are axially arranged, wherein the tie rod <NUM> constituting a rotary shaft passes through substantially a central portion of the compressor rotor disks <NUM>. Here, the neighboring compressor rotor disks <NUM> are disposed so that opposed surfaces thereof are pressed against each other by being axially clamped between the torque tube <NUM> and a counter member fastened to an end of the tie rod <NUM> and the neighboring compressor rotor disks do not rotate relative to each other.

A plurality of blades <NUM> are radially coupled to an outer circumferential surface of each compressor rotor disk <NUM>. Each of the blades <NUM> is fastened to the respective compressor rotor disk <NUM>.

Vanes (not shown) fixed to the housing are respectively positioned between the compressor rotor disks <NUM>. Unlike the rotor disks, the vanes are fixed to the housing and do not rotate. The vane serves to align a flow of compressed air that has passed through the blades <NUM> of the compressor rotor disk <NUM> and guide the air to the blades <NUM> of the rotor disk <NUM> located on the downstream side.

As schematically shown in <FIG>, the tie rod <NUM> is arranged to pass through the center of the compressor rotor disks <NUM> and turbine rotor disks <NUM>. To a first end of the tie rod <NUM>, a counterpart is fastened which presses against the compressor rotor disk located on the most upstream side, and to a second end of the tie rod <NUM>, a fixing nut <NUM> is fastened which presses against a turbine rotor disk located on the most downstream side. Generally, the turbine rotor disks <NUM> are axially clamped between the torque tube <NUM> and the fixing nut <NUM>. The tie rod <NUM> may be composed of a single tie rod or a plurality of tie rods.

The shape of the tie rod <NUM> is not limited to that shown in <FIG>, but may have a variety of structures depending on the gas turbine. That is, as exemplarily shown in <FIG>, one single tie rod having a shape passing through a central portion of the rotor disk may be provided, or a plurality of tie rods may be arranged in a circumferential manner, or a combination thereof may be used.

Although not shown, the compressor of the gas turbine may be provided with a guide vane serving as a guide element at a position adjacent to the diffuser in order to adjust a flow angle of a pressurized fluid entering a combustor inlet to a designed flow angle. The guide vane may be referred to as a deswirler.

The combustor <NUM> mixes the introduced compressed air with fuel and combusts the air-fuel mixture to produce a high-temperature and high-temperature and high-pressure combustion gas. With an isobaric combustion process in the compressor, the temperature of the combustion gas is increased to the heat resistance limit that the combustor and the turbine components can withstand.

The combustor <NUM> may comprise a plurality of combustors, which is arranged in the housing formed in a cell shape, and includes a burner having a fuel injection nozzle and the like, a combustor liner forming a combustion chamber, and a transition piece as a connection between the combustor and the turbine, thereby constituting a combustion system of a gas turbine.

Specifically, the combustor liner provides a combustion space in which the fuel injected by the fuel nozzle is mixed with the compressed air of the compressor and the fuel-air mixture is combusted. Such a liner may include a flame canister providing a combustion space in which the fuel-air mixture is combusted, and a flow sleeve forming an annular space surrounding the flame canister. A fuel nozzle is coupled to the front end of the liner, and an igniter plug is coupled to the side wall of the liner.

On the other hand, a transition piece is connected to a rear end of the liner so as to transmit the combustion gas combusted by the igniter plug to the turbine side. An outer wall of the transition piece is cooled by the compressed air supplied from the compressor so as to prevent thermal breakage due to the high temperature combustion gas.

To this end, the transition piece is provided with cooling holes through which compressed air is injected into and cools the inside of the transition piece and flows towards the liner.

The air that has cooled the transition piece flows into the annular space of the liner and compressed air is supplied as a cooling air to the outer wall of the liner from the outside of the flow sleeve through cooling holes provided in the flow sleeve so that both air flows may collide with each other.

In the meantime, the high-temperature and high-pressure combustion gas from the combustor is supplied to the turbine <NUM>. The high-temperature and high-pressure combustion gas expands and collides with the rotating blades of the turbine, generating a reaction force that imparts a rotational torque. This torque is subsequently transferred to the compressor via the torque tube <NUM>. Any surplus power not needed to drive the compressor may be used to drive a generator or similar equipment.

The turbine <NUM> is basically similar in structure to the compressor. That is, the turbine <NUM> is also provided with a plurality of turbine rotor disks <NUM> similar to the compressor rotor disks of the compressor. Thus, each turbine rotor disk <NUM> also includes a plurality of turbine blades <NUM> extending radially from the turbine rotor disk <NUM>. The respective turbine blade <NUM> may be coupled to the turbine rotor disk <NUM> in a dovetail coupling manner, for example. Between the axially spaced turbine blades <NUM>, a turbine vane <NUM> fixed to a turbine casing <NUM> is provided to guide a flow direction of the combustion gas passing through the turbine blades <NUM>.

The turbine vane <NUM> is fixedly mounted within the turbine casing <NUM> by a turbine vane platform <NUM> coupled to radially inner and outer ends of the turbine vane <NUM>. On the other hand, in a position facing the radially outer end of the rotating turbine blade <NUM> on the inner side of the turbine casing <NUM>, a ring segment <NUM> is mounted to form a predetermined gap with the radially outer end of the turbine blade <NUM>.

The blade <NUM>, <NUM> may be installed on the rotor disk <NUM>, <NUM> by T-type pin fitting, wedge insertion, caulking, or the like. The T-type pin fitting and the wedge insertion are limited in shape when installation space is tight or the disk size is small, and the caulking has a problem of causing permanent damage to the disk. Therefore, provisions that are structurally safe and have ready-to-improve installation workability have been sought in installing the blade <NUM>, <NUM> on the rotor disk <NUM>, <NUM>.

Further, the engagement of the blades <NUM>, <NUM> with the rotor disk <NUM>, <NUM> is somewhat loosely performed due to assembly and tolerance. Accordingly, as a result of frequent shutdowns of a gas turbine or vibrations generated during operation of a gas turbine, the blade <NUM>, <NUM> may move in a direction parallel to the center tie rod <NUM> (also referred to herein as an axial direction). This axial movement of the blade <NUM>, <NUM> may cause wear due to friction between the blade <NUM>, <NUM> and the rotor disk <NUM>, <NUM>. Therefore, there is a need for a provision to prevent the axial movement of the blade <NUM>, <NUM>.

In the following, a description will illustrate the case where a blade and a rotor disk refer to a turbine blade <NUM> and a turbine rotor disk <NUM>, but the present invention may also be applied to the case where a blade and a rotor disk are a compressor blade <NUM> and a compressor rotor disk <NUM>.

<FIG> is a perspective view illustrating a turbine blade and a turbine rotor disk of a rotor assembly.

Referring to <FIG>, the turbine blade <NUM> includes an airfoil <NUM>, a platform <NUM>, and a root member <NUM>.

The airfoil <NUM> is formed on an upper or first surface of the platform <NUM>. Generally, the airfoil <NUM> may protrude from the first surface of the platform <NUM> along a radial direction Y. The airfoil <NUM> may be formed to have an optimized aerodynamic shape depending on the specifications of a gas turbine. Generally, the airfoil <NUM> includes a leading edge LE disposed on or defining an upstream side and a trailing edge TE disposed on or defining a downstream side based on a flow direction of combustion gases.

The platform <NUM> is disposed at the lower end or platform end of the airfoil <NUM>. The platform <NUM> may be formed in a substantially rectangular plate shape. Cooling flow paths may be formed inside the airfoil <NUM> such that cooling air can flow therethrough via the platform <NUM>.

The root member <NUM> is disposed on a radially inner side or second surface of the platform <NUM>. In other words, the root member <NUM> and the platform <NUM> extend on opposite sides of the platform <NUM> with respect to the radial direction Y. The radially inner side of the platform <NUM> may be a lower portion of the platform <NUM>. The root member <NUM> may be formed to taper radially inwardly in width. A dovetail may be formed on both circumferential sides of the root member <NUM>. The dovetail may be fir-tree shaped in cross-section and may be formed in plurality.

The turbine rotor disk <NUM> may be generally disk shaped, and includes a plurality of disk slots <NUM> that are formed on an outer circumference of the turbine rotor disk <NUM>. The disk slots <NUM> may be formed to have a fir-tree shaped cross-section. Generally, the disk slots <NUM> and the root member <NUM> of a respective turbine blade <NUM> may be formed to have corresponding cross-sections. The disk slots <NUM> may extend along an axial direction X. The turbine blades <NUM> may be fastened and installed in the disk slots <NUM>, in particular, by introducing the root members <NUM> in the disk slots <NUM>.

Referring to <FIG>, the direction X is an axial direction, and the upstream side and the downstream side are defined based on the flow direction of the combustion gases along the axial direction X. The upstream end and the downstream end may be referred to as a forward end and a rearward end, respectively. The direction Y is a radial direction, along which the airfoil <NUM> extends from the platform <NUM>. The direction Z is a circumferential direction, along which the outer circumference of the disk-shaped turbine rotor disk <NUM> is defined.

When the turbine blade <NUM> is installed in the disk slot <NUM>, a blade fastening assembly <NUM> is utilized to secure the axial position of the blade and prevent the axial movement of the blade once installed. The blade fastening assembly <NUM> includes a first groove <NUM>, a second groove <NUM>, a pin groove <NUM>, a fastening key <NUM>, and a fastening pin <NUM>.

<FIG> is an enlarged perspective view illustrating the turbine blade being fastened to the turbine rotor disk, and <FIG> is a perspective view illustrating a fastening key <NUM> and a fastening pin <NUM> being engaged in a state of turbine blade of <FIG>.

Referring to <FIG>, the first groove <NUM> and the pin groove <NUM> are formed in an axially forward end of the turbine rotor disk <NUM> or generally in an axial end of the turbine rotor disk <NUM>, and the second groove <NUM> is formed at least partially in the platform <NUM>,specifically, in a lateral end portion of the platform <NUM>.

The first groove <NUM>, generally, is formed in the outer circumference of the turbine rotor disk <NUM> and extends along the axial direction X. For example, the first groove <NUM> may be formed by machining or cutting a portion of the axially forward end and the upper end of the turbine rotor disk <NUM>. The first groove <NUM> may be substantially in a cuboidal shape.

The pin groove <NUM> may be formed in an axial end surface of the first groove <NUM>, wherein the axial end surface limits the first groove <NUM> in the axial direction X. For example, the pin groove <NUM> may be formed by machining or cutting the end wall, i.e., the rearward end surface of the first groove <NUM>. The pin groove <NUM> may be in a cylindrical hole shape, as exemplarily shown in <FIG>. The pin groove <NUM> may extend along the axial direction X, e.g., toward the rearward rear end from the rearward end surface of the first groove <NUM>.

The second groove <NUM> may include a first portion, also referred to as first wing groove portion 2002a, and a second portion, also referred to as a second wing groove portion 2002b. The second groove <NUM> may extend along the circumferential direction or, generally, in a direction transverse to the extension direction of the first groove <NUM>. As exemplarily shown in <FIG>, the second groove <NUM> extends from or is connected to the first groove <NUM> and extends at least partially within the platform <NUM>. Optionally, the first wing groove portion 2002a may extend in the turbine rotor disk <NUM>, and the second wing groove portion 2002b may extend in the platform <NUM>. The second groove <NUM> may be formed, for example, by machining or cutting a portion of the turbine rotor disk <NUM> adjacent to the first groove <NUM> (i.e., the first wing groove portion 2002a) and a side portion of the platform <NUM> (i.e., the second wing groove portion 2002b). The first wing groove portion 2002a is extended from and in communication with the first groove <NUM> and may be formed by machining or cutting a portion of the turbine rotor disk <NUM> adjacent to the first groove <NUM> toward the platform <NUM>. The first wing groove portion 2002a may be formed such that its rearward end surface is in a same axial position with the rearward end surface of the first groove <NUM> and its forward end surface is located distanced from the forward end surface of the turbine rotor disk <NUM> in the axial direction X. Further optional, the first wing groove portion 2002a may be formed to have a smaller depth than a depth of the first groove <NUM> from the upper end surface of the turbine rotor disk <NUM> in the radial direction Y. The second wing groove portion 2002b may formed at the platform <NUM> such that the locations of its rearward end surface, its forward end surface and its lower end surface matches the ones of the rearward end surface, the forward end surface and the lower end surface of the first wing groove portion 2002a when the turbine blade <NUM> is inserted into the disk slot <NUM>. Thereby, the second groove <NUM>, which is a combined groove of the first wing groove portion 2002a and the second wing groove portion 2002b, may be substantially in a cuboidal shape.

The first groove <NUM> is formed in a shape matching and corresponding to a portion of the fastening key <NUM> (i.e., body part <NUM>) described below, and the second groove <NUM> is formed in a shape corresponding to the remainder of the fastening key <NUM> (i.e., the wing part <NUM>). In other words, the first groove <NUM> and the second groove <NUM> are formed such that integral shape of them matches a shape of the fastening key <NUM>, and such that the fastening key <NUM> may be matingly inserted into the combined and integral groove formed by the first groove <NUM> and the second groove <NUM>. The first groove <NUM> and the second groove <NUM> are formed in communication with each other when the turbine blade <NUM> is inserted into the disk slot <NUM>. The first groove <NUM> and the second groove <NUM> are connected to each other to form a shape corresponding to the shape of the fastening key <NUM>.

The pin groove <NUM> is formed in connection with the first groove <NUM>. The pin groove <NUM> is formed by extending axially rearwards from an inner surface (i.e., the rearward surface which is also referred to as axial and surface herein) of the first groove <NUM>. In the pin groove <NUM>, an end of the fastening pin <NUM> described later is inserted.

Referring to <FIG>, the fixing key <NUM> is inserted into the first groove <NUM> and the second groove <NUM>, and the fixing pin <NUM> is inserted into the pin groove <NUM> through a through hole <NUM> formed in the fixing key <NUM>. This allows the axial position of the turbine blade to be fixed using the fixing key <NUM> and the fixing pin <NUM> when the turbine blade <NUM> is matingly installed in the turbine disk slot <NUM>. Further, after installation, the fastening key <NUM> and fastening pin <NUM> can prevent axial movement of the turbine blade <NUM> within the turbine disk slot <NUM>.

<FIG> is a perspective view illustrating the fastening key and the fastening pin constituting a blade fastening assembly according to an embodiment of the present invention.

Referring to <FIG>, the fastening key <NUM> includes a body part <NUM> and a wing part <NUM>.

The body part <NUM> may be formed in a predetermined shape, such as a cuboidal shape having a first height H1 in the radial direction Y when it is assembled with the turbine rotor disk <NUM> and the turbine blade <NUM>. The body part <NUM>, generally, may have a shape corresponding to the shape of the first groove <NUM> so that the body part <NUM> can be matingly received in the first groove <NUM>.

A through hole <NUM> is formed through the body part <NUM> along the axial direction X. The through hole <NUM> is formed to be in communication with the pin groove <NUM> when the body part <NUM> is assembly with the turbine rotor disk <NUM>. For example, the through hole <NUM> may be positioned coaxially with pin groove <NUM> when the body part <NUM> is assembly with the turbine rotor disk <NUM>, i.e., when the body part <NUM> is received in the first groove <NUM>. A thread, in particular, an internal thread may be formed on an inner wall of the through hole <NUM> as exemplarily shown in <FIG>. The body part <NUM> may be inserted into the first groove <NUM>.

The wing part <NUM> is formed to extend transverse to the body part <NUM>. The wing part <NUM> may protrude from the wing part <NUM> with a predetermined length. When being assembled with the turbine rotor disk <NUM>, the wing part <NUM> extends toward the platform <NUM> from the body part <NUM> in the circumferential direction Z. The wing part <NUM> is formed to have a second height H2 that is smaller than the first height H1 in the radial direction Y when it is assembled with the turbine rotor disk <NUM> and the turbine blade <NUM>. The wing part <NUM> is formed to extend toward the platform <NUM> from a side surface of the body part in the circumferential direction Z. An upper surface of the wing part <NUM> may be flush with an upper surface of the body part <NUM>. Hence, the second height H2 may be measured from the upper surface of the body part <NUM>. Since the second height H2 is smaller than the first height H1, a lower surface of the wing part <NUM> may be positioned spaced from a lower surface of the body part <NUM>. A length of the wing part <NUM> in the circumferential direction Z may be equal to the length of the second groove <NUM> (i.e., the combined width of the first wing groove portion 2002a and the second wing groove portion 2002b). Generally, the length of the body part <NUM> is at least such that the wing part <NUM> may be inserted into the second groove <NUM> so as to protrude into the portion of the second groove <NUM> formed in the platform <NUM>.

Referring to <FIG>, the fastening pin <NUM> may be in the form of a cylinder extending a predetermined length, and the fastening pin <NUM> may be inserted into and secured in the through hole <NUM> and the pin groove <NUM>. The fastening pin <NUM> may include a first insertion part <NUM> and a second insertion part <NUM>.

The outer circumferential surface of the first insertion part <NUM> may be formed in a smooth shape, allowing first insertion part <NUM> of the pin <NUM> to be inserted into the pin groove <NUM>. The outer circumferential surface of the second insertion part <NUM> may be threaded to engage with the thread optionally provided in the through hole <NUM> to allow the fastening pin <NUM> to be securely fastened to the fastening key <NUM>. Alternatively, or additionally to the thread in the through hole <NUM>, it may also be provided that the pin groove <NUM> is provided with an internal thread (not shown). In this case also the first insertion part <NUM> may be provided with a thread to be engaged with the thread of the pin groove. Optionally, the thread in the through hole <NUM> and the thread on the second insertion part <NUM> may be omitted. Although not shown, a fastening structure <NUM> may be formed on an forward end surface (externally exposed side) of the second insertion part <NUM> such that a separate fastener <NUM> may be inserted therethrough to rotate the fastening pin <NUM>. When the fastening pin <NUM> is inserted, the fastener <NUM> can be inserted into the fastening structure <NUM> to rotate the fastening pin <NUM>, so that the fastening pin <NUM> can be firmly fastened with the thread provided in the through hole <NUM> and/or the pin groove <NUM>. The fastener <NUM> and the fastening structure <NUM> may also be used for disassembling the fastening pin <NUM> from the through hole <NUM> and the pin groove <NUM>.

Next, a process of installing a turbine blade into a turbine disk slot will be described with reference to <FIG> and <FIG>. <FIG> and <FIG> are views illustrating a process of installing a turbine blade on a turbine rotor disk using a blade fastening <NUM> as described above. On the other hand, <FIG> and <FIG> may be understood as illustrating a process of manufacturing and assembling a gas turbine.

First, as illustrated in <FIG>, the first groove <NUM> and the pin groove <NUM> are formed in the turbine rotor disk <NUM>, and the second groove <NUM> is formed adjacent to the first groove <NUM> on the turbine rotor disk <NUM> and on the platform <NUM>. Next, the fastening key <NUM> is inserted into the first groove <NUM> and the second groove <NUM>. At this time, the body part <NUM> of the fastening key <NUM> is inserted into the first groove <NUM>, and the wing part <NUM> is inserted into the second groove <NUM>.

Next, as illustrated in <FIG>, the fastening pin <NUM> is inserted into the through hole <NUM>, and then rotated so that the fastening pin <NUM> is firmly fastened with the thread formed in the through hole <NUM> and/or in the pin groove <NUM>.

In this way, by inserting the fastening key <NUM> into the first groove <NUM> and the second groove <NUM> and the fastening pin <NUM> into the through hole <NUM> and the pin groove <NUM>, the installation of the blade on the rotor disk is structurally safe and improves installation workability. Furthermore, preventing the axial movement of the blade after installation to minimize wear resulting from friction between the blade and the rotor disk can enhance the overall durability of a gas turbine.

The blade fastening assembly <NUM> according to the present invention is particularly advantageous in cases where <NUM>) the working space on both axial sides the rotor disk is narrow, and <NUM>) the shape of the fastening pin or key is limited due to the small size of the blade and rotor disk.

Claim 1:
A rotor assembly for a gas turbine, comprising:
a blade (<NUM>, <NUM>) comprising a platform (<NUM>) and a root member (<NUM>);
a rotor disk (<NUM>, <NUM>) including a disk slot (<NUM>) into which the root member (<NUM>) is inserted; and
a blade fastening assembly (<NUM>), comprising:
a fastening key (<NUM>) having a through hole (<NUM>);
a first groove (<NUM>) formed in the rotor disk (<NUM>, <NUM>) with a shape corresponding to a portion of the fastening key (<NUM>),
a second groove (<NUM>) formed in communication with the first groove (<NUM>) and at least partially extending in the platform (<NUM>) of the blade (<NUM>, <NUM>), the second groove (<NUM>) having a shape corresponding to a remaining portion of the fastening key (<NUM>) in a state in which the root member (<NUM>) is inserted into the disk slot (<NUM>), wherein the fastening key (<NUM>) is positioned within the first groove (<NUM>) and the second groove (<NUM>),
a pin groove (<NUM>) connected with the first groove (<NUM>); and
a fastening pin (<NUM>) inserted into the through hole (<NUM>) and the pin groove (<NUM>),
wherein the first groove (<NUM>) and the pin groove (<NUM>) are formed in an axial end of the rotor disk (<NUM>, <NUM>), and the second groove (<NUM>) is formed to extend in a side portion of the platform (<NUM>).