Method and system for rotor overspeed protection

A gas turbine engine, system, and method of severing at least one blade of a row of blades spaced about a rotor during an over speed condition are provided. The gas turbine engine includes a shaft rotatable about an axis of the shaft, a rotor coupled to the shaft and having a radially outer surface, and one or more rows of blades spaced axially along the outer portion. Each row of blades includes a plurality of blades spaced circumferentially around the radially outer surface. A gas turbine engine casing at least partially surrounds the one or more rows of blades. One or more rows of stationary vanes are spaced axially along the casing. Each row of stationary vanes includes a plurality of stationary vanes extending radially between the shaft and the gas turbine engine casing. The stationary vanes are coupled to the gas turbine engine casing using a fuse element.

BACKGROUND

The field of the disclosure relates generally to gas turbine engines and, more particularly, to rotor over speed protection.

Gas turbine engines include a rotor assembly which is rotatable relative to stationary engine structures. The engine may include a plurality of stages. Each stage includes a rotor blade assembly including a plurality of circumferentially spaced airfoils or blades, and a stationary vane assembly including a plurality of circumferentially spaced airfoils or vanes. Each rotor blade assembly is coupled to a rotor on a shaft which is configured to drive the gas turbine engine. Hot combustion gases interact with the blades to drive the rotor assembly. In some known rotor assemblies, disengagement of a portion of the rotor may result in an undesirable over speed condition. For example, disengagement of aft stages occurs when an aft rotor blade assembly stage separates from the rotor. The separated stage is still powered by the hot combustion gases, but the separated stage can no longer transfer that power to the shaft. The separated stage may rotate significantly faster than it rotates during normal operations because it is unloaded and still receiving power input from the combustion gases, causing the over speed condition.

BRIEF DESCRIPTION

In one aspect, a gas turbine engine includes a shaft rotatable about a longitudinal axis of the shaft, a rotor coupled to the shaft and having a radially outer surface, and one or more rows of blades spaced axially along the outer portion. Each of the one or more rows of blades includes a plurality of blades spaced circumferentially around the radially outer surface. The gas turbine engine also includes a gas turbine engine casing at least partially surrounding the one or more rows of blades, and one or more rows of stationary vanes spaced axially along the gas turbine engine casing. Each of the one or more rows of stationary vanes includes a plurality of stationary vanes extending radially between the shaft and the gas turbine engine casing. The stationary vanes are coupled to the gas turbine engine casing using a fuse element.

Optionally, the plurality of stationary vanes are spaced circumferentially about a radially inner surface of the gas turbine engine casing and the one or more rows of stationary vanes may be interdigitated axially with the one or more rows of blades. Also optionally, the gas turbine engine casing includes a railing extending circumferentially about a radially inner surface of the gas turbine engine casing, and an upstream end of a stationary vane of the plurality of stationary vanes includes a hook configured to engage the railing. At least one of the railing and the hook forms the fuse element. A pin element may extend through the railing and the hook, and the pin element also may form the fuse element.

Optionally, the gas turbine engine casing includes a railing extending circumferentially about a radially inner surface of the gas turbine engine casing, and at least some of the plurality of stationary vanes are ganged together in a vane sector assembly. An upstream end of the vane sector assembly includes a hook configured to engage the railing. At least one of the railing and the hook forms the fuse element. A pin element may extend through the railing and the hook, and the pin element also may form the fuse element.

The gas turbine engine casing may include an upstream extending lip extending circumferentially about a radially inner surface of the gas turbine engine casing, and a downstream end of a stationary vane of the plurality of stationary vanes may include a ledge configured to engage the lip. At least one of the lip and the ledge may form the fuse element. Optionally, the pin element extends through the lip and the ledge, and the pin element also may form the fuse element.

The gas turbine engine casing may include an upstream extending lip extending circumferentially about a radially inner surface of the gas turbine engine casing, and at least some of the plurality of stationary vanes may be ganged together in a vane sector assembly. A downstream end of the vane sector assembly may include a ledge configured to engage the lip, and at least one of the ledge, the lip, and a pin element extending through the ledge and the lip form the fuse element. A pin element may extend through the lip and the ledge, and the pin element also may form the fuse element.

Optionally, one or more rows of stationary vanes may be spaced uniformly axially along the gas turbine engine casing.

In another aspect, a method of severing at least one blade of a row of blades spaced circumferentially about a rotor of a turbine during an over speed condition of the rotor is provided. The method includes driving the rotor with engine combustion gases after a separation of the rotor from a shaft of the turbine, and translating the rotor axially along the shaft using an axial force of the rotor. The method also includes engaging the at least one blade of the row of blades with at least one stationary vane of a plurality of stationary vanes, and severing at least a portion of the at least one blade of the row of blades from the rotor using the at least one stationary vane of the plurality of stationary vanes.

The method optionally includes severing at least the portion of the at least one blade of the row of blades from the rotor using at least one of a root, a shank, an airfoil, a tip, and a shroud of the at least one stationary vane of the plurality of stationary vanes. The method may also include the step of imparting sufficient force on a fuse element to cause the fuse element to part, which may include tilting the at least one stationary vane of the plurality of stationary vanes into a rotational path of the at least one blade of the row of blades. Optionally, the method includes fracturing the at least one blade of the row of blades into a plurality of pieces of blade debris. Optionally, the rotor may include a plurality of rows of blades spaced axially along the shaft of the turbine, and the method may also include the step of directing the blade debris into a downstream row of blades of the plurality of rows of blades, such that the blade debris causes fracturing of at least some of the blades of the downstream row of blades of the plurality of rows of blades.

In yet another aspect, a hanger system includes one or more rows of airfoils spaced axially along a machine casing. Each of the one or more rows of airfoils includes a plurality of airfoils extending from the machine casing. At least one of the airfoils in a row being coupled to the machine casing by a hangar that includes a fuse element. Optionally, the machine casing includes a railing extending about a surface of the machine casing. A first end of an airfoil of the plurality of airfoils includes a hook configured to engage the railing. At least one of the railing and the hook includes the fuse element. The hanger system may further include a pin element extending through the railing and the hook wherein the pin element includes the fuse element.

Also optionally, the machine casing includes a lip extending about the surface of the machine casing, and wherein a second end of the airfoil of the plurality of airfoils includes a ledge configured to engage the lip. At least one of the lip and the ledge includes the fuse element. The hanger system may further include a pin element extending through the lip and the ledge wherein the pin element includes the fuse element.

DETAILED DESCRIPTION

The following detailed description illustrates embodiments of the disclosure by way of example and not by way of limitation. It is contemplated that the disclosure has general application to mitigating damage in a rotor overspeed event in industrial and commercial applications.

Embodiments of a system for mitigating damage resulting from an overspeed condition of a portion of a machine are described herein. In some embodiments, the machine is a rotatable machine that has a rotatable portion, such as, but not limited to a rotor configured to rotate about a longitudinal axis or axis of rotation. The machine may include rotatable machines, such as, but not limited to gas turbine engines, compressors, motors, steam turbines, wind turbines, water turbines, and the like. The embodiments include a clashing mechanism that uses a momentum or torque of the overspeeding machine to cause a controlled mitigating event that eventually stops the overspeed. The clashing mechanism facilitates inhibiting energy sources from supplying the rotor, and converts the energy to another form that facilitates stopping the overspeed. Additionally or alternatively, the clashing mechanism can damage the ability of rotor to generate work. More specifically, even if only a portion of the blades are lost or damaged, the rotor will become less efficient in generating work. The reduction of work extraction from the energy source reduces the acceleration of the rotor. The embodiments described herein reduce the terminal speeds reached by a rotor during an over speed condition. Reduced terminal speeds result in an ability to design lighter-weight rotors.

For example, mid-shaft events in turbofan engines result in loss of load on the low pressure turbine, causing the turbine to accelerate. In some embodiments, the rotating blades of the turbine are clashed with stationary vanes, causing damage to the blades. The loss of blades results in a loss of extraction of work from the hot gases of combustion, resulting in limiting the over speed of the turbine. The clash may be caused by an axial force acting on the rotor, resulting in the rotor translating aftward into the stationary nozzle vanes. The clashing of the turbine rotor with the stationary vanes results in higher axial loads on the stationary vanes, as compared to normal operating loads. As described herein, a mechanical fuse element is positioned in the stationary vane, such that the stationary vane liberates during the overspeed condition due to an excessive axial and/or circumferential load. The liberated stationary vane will move aft due the force exerted by the hot combustion gases and/or by the clashing force. The liberated stationary vane will then move into a plane of rotor rotation and cause additional obstruction to the rotor. The obstruction to the rotor blades in the path of rotation accelerates the damage of blades and results in a quick loss of the ability of the turbine to extract work, thereby limiting the excessive rotor speeds in the turbine. Such a clashing mechanism can result in lighter turbine rotor designs, as the over speed resulting from the event will be limited.

During a low pressure turbine clash event, as described above, an axial force acting on the rotor will move the rotor aftward to clash against the stationary vane casing assembly. The axial rotor-to-stationary-vane gaps at different stages of the rotor are different, which causes the rotor to clash against the stationary vanes in the stage having the smallest axial gap, resulting in the stationary vanes of that stage taking the entire axial load of the rotor. This axial load imparted to the stationary vanes (typically loading at the inner diameter of the stationary vanes due to a minimum clearance adjacent to angel wings of the blades) is reacted by a bending moment hinging on the casing assembly. When such axial load is applied, embodiments of the present disclosure permit at least a portion of the stationary vane support to decouple, detaching the stationary vane(s) from the casing assembly. The detached stationary vane is liberated in the hot gas flow path, and the fluid force from the combustion gases and/or the force imparted due to clashing cause the stationary vane to move aftward and to come into a path of rotation of rotor blades of downstream stages. Impacting the downstream rotor blades causes significant damage to the rotor blades. This means of rotor blade destruction is relatively fast and facilitates slowing down the acceleration of the rotor, thereby reducing the terminal speed of the over speed condition. This liberation of stationary vanes into the hot gas flow path has demonstrated significant damage to the rotor blades during evaluation.

The technical advantages of the described apparatus and method over the present state of the art include improved efficiency of the low pressure turbine due to elimination of banana vanes commonly used for over speed blade destruction, more efficient axial gaps between the rotor blades and stationary vanes in each stage, and an elimination of the need to compromise on other low pressure turbine axial clearances, which results in efficient secondary flow systems and mechanical designs. Elimination of the banana vanes also permits a reduction in engine weight due to an axial length reduction and a reduction in a weight of the rotor system, as the design may be limited to lower overspeed conditions.

In addition, using fuse elements in clashing components avoids a need to have a three dimensional (3-D) shape of the clashing vane. Known clashing vanes include a 3-D airfoil with forward sweep in the midspan region. With the fuse element clashing component technology described herein, the clashing vane need not be forward swept and can also be straight. Also, because the sweep is straight at the leading edge and trailing edge of the clashing vane, the axial gaps between the clashing vanes and the upstream/downstream blade rows are uniform, with less variation from hub to tip. Due to this, mixing losses in the upstream/downstream blade rows, and unsteady flow losses associated with the mixing losses, are reduced significantly. Further, with the removal of the banana-shaped clashing vanes, the non-uniformity in the total pressure profile downstream of the clashing vane is avoided, which is beneficial for the reduction in vorticity in the downstream blades and vanes.

The following description refers to the accompanying drawings, in which, in the absence of a contrary representation, the same numbers in different drawings represent similar elements.

FIG. 1is a perspective view of an aircraft100. In the example embodiment, aircraft100includes a fuselage102that includes a nose104, a tail106, and a hollow, elongate body108extending therebetween. Aircraft100also includes a wing110extending away from fuselage102in a lateral direction112. Wing110includes a forward leading edge114in a direction116of motion of aircraft100during normal flight and an aft trailing edge118on an opposing edge of wing110. Aircraft100further includes at least one gas turbine engine120configured to drive a bladed rotatable member122or fan to generate thrust. An engine pylon124, for example, may couple the at least one gas turbine engine120to at least one of wing110and fuselage102, for example, in a pusher configuration proximate tail106.

FIG. 2is a schematic cross-sectional view of a rotatable machine, such as, gas turbine engine120in accordance with an exemplary embodiment of the present disclosure. In the example embodiment, gas turbine engine120is embodied in a high bypass turbofan jet engine. As shown inFIG. 2, gas turbine engine120defines an axial direction A (extending parallel to a longitudinal axis202provided for reference) and a radial direction R. In general, gas turbine engine120includes a fan assembly204and a core turbine engine206disposed downstream from fan assembly204.

In the example embodiment, core turbine engine206includes an approximately tubular gas turbine engine casing208that defines an annular core engine inlet220. Gas turbine engine casing208encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor222and a high pressure (HP) compressor224; a combustor226; a turbine section including a high pressure (HP) turbine228and a low pressure (LP) turbine230; and a jet exhaust nozzle232. A high pressure (HP) spool or shaft234drivingly connects HP turbine228to HP compressor224. A low pressure (LP) spool or shaft236drivingly connects LP turbine230to LP compressor222. The compressor section, combustor226, turbine section, and jet exhaust nozzle232together define a core air flowpath237.

In the example embodiment, fan assembly204includes a variable pitch fan238having a plurality of fan blades240coupled to a fan disk242in a spaced apart relationship. Fan blades240extend radially outwardly from fan disk242. Each fan blade240is rotatable relative to fan disk242about a pitch axis P by virtue of fan blades240being operatively coupled to a suitable pitch change mechanism (PCM)244configured to vary the pitch of fan blades240. In other embodiments, pitch change mechanism244is configured to collectively vary the pitch of fan blades240in unison. Fan blades240, fan disk242, and pitch change mechanism244are together rotatable about longitudinal axis202by LP shaft236across a power gear box246. Power gear box246includes a plurality of gears for adjusting the rotational speed of fan238relative to LP shaft236to a more efficient rotational fan speed.

Fan disk242is covered by rotatable front hub248aerodynamically contoured to promote an airflow through the plurality of fan blades240. Additionally, fan assembly204and at least a portion of core turbine engine206are surrounded by a nacelle assembly249, which may include an annular fan casing or outer nacelle250that circumferentially surrounds fan238and/or at least a portion of core turbine engine206. In the example embodiment, outer nacelle250is configured to be supported relative to core turbine engine206by a plurality of circumferentially-spaced outlet guide vanes252. Moreover, a downstream section254of outer nacelle250may extend over an outer portion of core turbine engine206so as to define a bypass duct256therebetween.

Nacelle assembly249is a system of components or structures attached to gas turbine engine120and/or engine pylon124, which provides aerodynamic surfaces around gas turbine engine120, defines a portion of bypass duct256, defines core engine inlet220and bypass duct256, defines appropriate nozzles for the exhaust of bypass duct256and a core exhaust257, and houses or contains auxiliary devices for the engine and other components for the aircraft including various ducts, lines, pipes and wires.

Nacelle assembly249may be subdivided into outer nacelle250and a core engine cowl or inner nacelle259generally separated by bypass duct256. Outer nacelle250may include an inlet opening260. Outer nacelle250may also partially overlap a forward portion261of inner nacelle259, with outer nacelle250providing a radially outer wall for bypass duct256and inner nacelle259providing a radially inner wall.

Inner nacelle259forms at least a part a generally cylindrical or barrel-shaped cowl formed around the gas turbine engine casing208. Inner nacelle259houses and is configured to provide an aerodynamic cover for gas turbine engine casing208.

During operation of gas turbine engine120, a volume of air258enters gas turbine engine120through associated inlet opening260of outer nacelle250and/or fan assembly204. As volume of air258passes across fan blades240, a bypass portion262of volume of air258is directed or routed into bypass duct256and a core engine portion264of volume of air258is directed or routed into core air flowpath237, or more specifically into LP compressor222. A ratio between bypass portion262and core engine portion264is commonly referred to as a bypass ratio. The pressure of core engine portion264is then increased as it is routed through high pressure (HP) compressor224and into combustor226, where it is mixed with fuel and burned to provide combustion gases266.

Combustion gases266are routed through HP turbine228where a portion of thermal and/or kinetic energy from combustion gases266is extracted via sequential stages of HP turbine stationary airfoils or vanes268that are coupled to gas turbine engine casing208and HP turbine rotor blades270that are coupled to HP spool or shaft234, thus causing HP spool or shaft234to rotate, which then drives a rotation of HP compressor224. Combustion gases266are then routed through LP turbine230where a second portion of thermal and kinetic energy is extracted from combustion gases266via sequential stages of stationary airfoils or vanes272that are coupled to a radially inner surface273of gas turbine engine casing208and LP turbine rotor blades274that are coupled to LP spool or shaft236through a rotor235of LP turbine230having a radially outer surface239. LP spool or shaft236drives a rotation of LP compressor222and/or rotation of fan238. In various embodiments, sequential stages of stationary vanes272include a plurality of circumferentially spaced individual vanes, each attached separately through a hangar system306(shown inFIG. 3). In other embodiments, sequential stages of stationary vanes272are attached to gas turbine engine casing208in a gang with others of the plurality of circumferentially spaced vanes, which form a vane sector400(shown inFIG. 4), sometimes referred to as a doublet if two vanes are ganged together, a triplet if three vanes are ganged together, and so on.

Combustion gases266are subsequently routed through jet exhaust nozzle232of core turbine engine206to provide propulsive thrust. Simultaneously, the pressure of bypass portion262is substantially increased as bypass portion262is routed through bypass duct256before it is exhausted from a fan exhaust nozzle276of gas turbine engine120, also providing propulsive thrust. HP turbine228, LP turbine230, and jet exhaust nozzle232at least partially define a hot gas path278for routing combustion gases266through core turbine engine206.

Gas turbine engine120is depicted in the figures by way of example only. In other exemplary embodiments, gas turbine engine120may have any other suitable configuration including for example, a turboprop engine, a turboshaft engine, a military purpose engine, and a marine or land-based aero-derivative engine.

FIG. 3is a side elevation view of a stationary airfoil or vane300in accordance with an example embodiment of the present disclosure. In the example embodiment, stationary vane300is implemented as one of LP stationary vanes272of gas turbine engine120(shown inFIG. 2). More specifically, in the example embodiment, gas turbine engine casing208at least partially surrounds LP turbine230(shown inFIG. 2). Gas turbine engine casing208includes a circumferentially extending railing302and a circumferentially extending ledge304axially spaced from railing302. Stationary vane300includes a hanger system306coupled to casing208. More specifically, hanger system306includes a first hanger or hook308configured to engage railing302, and a second hanger or lip310configured to engage ledge304. In various embodiments, the engagement of lip310and ledge304may hold lip310and ledge304fixed in one axial position, while the engagement of hook308and railing302may permit sliding axial movement of hook308with respect to railing302to accommodate expansion and contraction during heating and cooling of LP turbine230. Hanger system306further includes a fuse element312formed in at least one of railing302, hook308, lip310, ledge304, a pin314extending through hook308and railing302, and a pin316extending through lip310and ledge304.

With reference toFIGS. 2 and 3, gas turbine engine casing208includes one or more rows or stages of stationary vanes272of LP turbine230that are interdigitated with one or more rows or stages of LP turbine rotor blades274. In various embodiments, at least one stationary vane300of one or more rows or stages of stationary vanes272are coupled to gas turbine engine casing208via hanger system306. Fuse element312is configured to decouple when a blade of the one or more rows of blades274clashes with the one or more rows of vanes272. Fuse element312may be formed of a reduced or “necked-down” dimensional feature to provide a controlled decoupling location. For example, pins314and316may include an area of reduced diameter along a length of the pin. Preferably, the area of reduced diameter of pin314is positioned where railing302and hook308engage, and the area of reduced diameter of pin316is positioned where lip310and ledge304engage. Additionally or alternatively, fuse element312may include a treated portion such that the treated portion is weakened, relative to other portions of fuse element312, by the treatment. A laser, chemical, heat, mechanical, or other treatment may be used to weaken the treated portion of fuse element312without affecting its physical dimensions. Fuse element312may also be formed of a material than is different from a material of the components which fuse element312is in contact with. If fuse element312is formed in an additive manufacturing process, fuse element312may be built up of a different material than the surrounding material. As used herein, “additive manufacturing” refers to any process which results in a three-dimensional object and includes a step of sequentially forming the shape of the object one layer at a time. Additive manufacturing processes include, for example, three dimensional printing, laser-net-shape manufacturing, direct metal laser sintering (DMLS), direct metal laser melting (DMLM), selective laser sintering (SLS), plasma transferred arc, freeform fabrication, and the like. One exemplary type of additive manufacturing process uses a laser beam to sinter or melt a powder material. Additive manufacturing processes can employ powder materials or wire as a raw material. Moreover, additive manufacturing processes can generally relate to a rapid way to manufacture an object (article, component, part, product, etc.) where a plurality of thin unit layers are sequentially formed to produce the object. For example, layers of a powder material may be provided (e.g., laid down) and irradiated with an energy beam (e.g., laser beam) so that the particles of the powder material within each layer are sequentially sintered (fused) or melted to solidify the layer.

A radially inner end318of stationary vane300may be independent of the radially inner ends of adjacent vanes or may be shrouded or ganged with radially inner ends318of adjacent vanes, such as with the radially inner ends of doublet or triplet vanes.

FIG. 4is a planform section view of a vane sector400in accordance with an example embodiment of the present disclosure. In the view ofFIG. 4, hanger system306is omitted for clarity. Stationary vanes300are positioned side-by-side with their respective radially inner ends318(shown inFIG. 3) coupled to a radially inner shroud402to form vane sector400. In the example embodiment, vane sector400is a triplet, that is, vane sector400includes three stationary vanes300. Alternatively, vane sector400includes any suitable number of stationary vanes300.

FIG. 5is a side elevation view of another example embodiment of hanger system306. In the example embodiment, fuse element312may be formed at any of several locations on railing302, hook308, lip310, and ledge304. For example, fuse element312may include a first railing fuse element502formed axially across a radially inwardly extending portion504of railing302. Similarly, fuse element312may include a second railing fuse element506formed radially across an axially extending portion508of railing302.

Other embodiments of fuse element312may be formed on hook308. For example, fuse element312may include a first hook fuse element510formed axially across a radially outwardly extending portion512of hook308. Similarly, fuse element312may include a second hook fuse element514formed radially across an axially extending portion516of hook308.

Additionally or alternatively, fuse element312may include a first ledge fuse element518formed axially across a radially inwardly extending portion520of ledge304. Similarly, fuse element312may include a second ledge fuse element522formed radially across an axially extending portion524of ledge304.

Still other embodiments of fuse element312may be formed on lip310. For example, fuse element312may include a first lip fuse element526formed axially across a radially outwardly extending portion528of lip310. Similarly, fuse element312may include a second lip fuse element530formed radially across an axially extending portion532of lip310.

FIG. 6is a side elevation view of another example embodiment of hanger system306. In the example embodiment, fuse element312may be formed from, for example, pin314extending through hook308and railing302, and/or pin316extending through lip310and ledge304. In each of pin314and pin316, a respective weakened area602and weakened area604is formed by dimensional reduction or by chemical, mechanical, or other treatment, to alter the structure of weakened area602and/or weakened area604.

FIG. 7is a flowchart of a method700of severing at least one blade of a row of blades spaced circumferentially about a rotor of a turbine, such as at least one of blades274of rotor235of gas turbine engine120(shown inFIG. 2), during an over speed condition of the rotor. With reference also toFIG. 2, in the example embodiment, method700includes driving702the rotor with engine combustion gases after a separation of the rotor from a shaft of the turbine, such as shaft236. Method700also includes translating704the rotor axially along the shaft using an axial force of the rotor, and engaging706the at least one blade of the row of blades with at least one stationary vane of a plurality of stationary vanes, such as one of stationary vanes272. Method700further includes severing708the at least one blade of the row of blades from the rotor using the at least one stationary vane of the plurality of stationary vanes.

The above described embodiments of a method and system of clashing rotor blades on an overspeeding rotor with stationary vanes coupled to a fixed casing provides a cost-effective and reliable means for reducing damage in an overspeed event More specifically, the methods and systems described herein facilitate providing an overspeed mitigation system without the use of a banana vane or a control mechanism to limit overspeed events such as, but not limited to a sensing system and fuel controls. Reducing the speed the overspeed can attain also permits reducing the design parameters to which the rotor assembly is formed, which in turn allows for additional weight reduction In addition, the above described methods and systems facilitate providing additional room in the flowpath of the machine due to the absence of the banana vane. As a result, the methods and systems described herein facilitate improving the efficiency and manufacturability of a machine in a cost-effective and reliable manner.