GAS TURBINE ENGINE COMPRESSOR UNDERCUT SPACER

Technical Field

The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a compressor undercut spacer of a gas turbine engine.

Background

Gas turbine engines include compressor, combustor, and turbine sections. Components of the gas turbine engine sections are subject to high temperatures and pressures. These temperatures and pressures may vary during transients of the gas turbine engine, especially during start up and shut down of the gas turbine engine. The components may thermally expand at different rates generating thermal stresses and strains within the components.

U.S. Patent Application Pub. No. 2009/0297350, to S. Augustine discloses a mount interface between two rotating gas turbine engine components including a rigid ring to provide radial deflection restraint. According to

Augustine a hoop snap spacer or rigid ring could be used as a spacer providing connection between rotor stages near a rim where the blades are supported. The rigid ring ties the rotor stages together via friction rather than bolts.

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 compressor spacer ring configured for mounting between compressor rotor disks has a body with a hollow cylinder shape. The body includes an outer surface. The outer surface is the radially outermost circumferential surface of the body. The outer surface includes an axially forward end, and an axially aft end. The body also includes a forward face extending radially inward from the axially forward end of the outer surface. A forward lip extends axially from the body. The forward lip includes a forward lip surface located radially inward from the forward face. The forward lip surface is the radially outer circumferential surface of the forward lip. The spacer also includes a forward undercut. The forward undercut is located between the forward face and the forward lip surface.

Brief Description of the Drawings

FIG. 1 is a schematic illustration of an exemplary gas turbine engine.

FIG. 2 is a perspective view an aft portion of a compressor rotor assembly of an exemplary gas turbine engine.

FIG. 3 is a cross-sectional view of a portion of a compressor of an exemplary gas turbine engine.

FIG. 4 is a cross-sectional view of a portion of a spacer for a compressor of an exemplary gas turbine engine.

Detailed Description

FIG. 1 is a schematic illustration of an exemplary gas turbine engine. A gas turbine engine 100 typically includes a compressor 200, a combustor 300, a turbine 400, and a shaft 120. The gas turbine engine 100 may have a single shaft or a dual shaft configuration. For convention in this disclosure all references to radial, axial, and circumferential directions and measures refer to center axis 95 unless otherwise specified. Center axis 95 may generally be defined by the longitudinal axis of shaft 120. Center axis 95 may be common to or shared with other gas turbine engine concentric components.

Air 10 enters an inlet 15 as a "working fluid" and is compressed by the compressor 200. Fuel 35 is added to the compressed air in the combustor 300 and then ignited to produce a high energy combustion gas. Energy is extracted from the combusted fuel/air mixture via the turbine 400 and is typically made usable via a power output coupling 5. The power output coupling 5 is shown as being on the forward side of the gas turbine engine 100, but in other configurations it may be provided at the aft end of gas turbine engine 100.

Exhaust 90 may exit the system or be further processed (e.g., to reduce harmful emissions or to recover heat from the exhaust).

The compressor 200 includes a compressor rotor assembly 210 mechanically coupled to shaft 120. As illustrated, the compressor rotor assembly 210 is an axial flow rotor assembly. The compressor rotor assembly 210 includes one or more compressor disk assemblies 220. Spacers 230 span between adjacent compressor disk assemblies 220. Compressor stationary vanes ("stators") 250 axially precede each or the compressor disk assemblies 220.

The turbine 400 includes a turbine rotor assembly 410 mechanically coupled to the shaft 120. As illustrated, the turbine rotor assembly 410 is an axial flow rotor assembly. The turbine rotor assembly 410 includes one or more turbine disk assemblies 420. Each turbine disk assembly 420 includes a turbine rotor disk that is circumferentially populated with turbine rotor blades. Turbine nozzles 450 axially precede each of the turbine rotor assemblies 420. The turbine nozzles 450 have circumferentially distributed turbine nozzle vanes. The turbine nozzle vanes helically reorient the combustion gas that is delivered to the turbine rotor blades where the energy in the combustion gas is converted to mechanical energy and rotates the shaft 120.

The various components of the compressor 200 are housed in a compressor case 201 that may be generally cylindrical. The various components of the combustor 300 and the turbine 400 are housed, respectively, in a combustor case 301 and a turbine case 401.

FIG. 2 is a perspective view of an aft portion of a compressor rotor assembly of an exemplary gas turbine engine which may be used in the gas turbine engine 100 of FIG. 1. The compressor rotor assembly 210 includes compressor disk assemblies 220, spacers 230, and rear hub 245. Each compressor disk assembly 220 includes a compressor rotor disk ("disk") 221 and one or more compressor rotor blades ("airfoils") 227. Disks 221 are coupled or welded together when forming the compressor rotor assembly 210. In the embodiment shown, disks 221 are coupled together with curvic teeth 219. Each disk 221 is circumferentially populated with airfoils 227. Each disk 221 may be formed from materials such as incoloy 901 or 410 stainless steel.

Rear hub 245 may be located aft of disks 221 and is generally the most aft component of compressor rotor assembly 210. Rear hub 245 may have a disk shape. Shaft interface 248 extends aft from the disk shape of rear hub 245 with a cylindrical shape. Shaft interface 248 may be tapered for coupling to a portion of shaft 120.

FIG. 3 is a cross-sectional view of a portion of the compressor 200 of a gas turbine engine which may be used in the gas turbine engine 100 of FIG. 1. Disk 221 of each compressor disk assembly 220 includes a rim 222, a forward arm 225, and an aft arm 226. Rim 222 is located at the radial outermost portion of the disk 221 and may be located at a radially outer circumference of disk 221. In one embodiment rim 222 circumferentially extends completely around disk 221. Generally, each rim 222 includes forward extension 223 extending axially forward and aft extension 224 extending axially aft. In one embodiment both forward extension 223 and aft extension 224 circumferentially extend completely around disk 221.

Forward extension 223 includes forward extension face 215 and forward extension surface 217, and aft extension 224 includes aft extension face 216 and aft extension surface 218. Forward extension face 215 is the axially forward most surface of forward extension 223. Forward extension face 215 is generally normal to the axis of disk 221. Aft extension face 216 is the axially aft most surface of aft extension 224. Aft extension face 216 is generally parallel to forward extension face 215. Forward extension surface 217 is the radially inner circumferential surface of forward extension 223. Aft extension surface 218 is the radially inner circumferential surface of aft extension 224.

Forward arm 225 and aft arm 226 are located radially inward from rim 222 and radially outward from the axis of disk 221. Forward arm 225 and aft arm 226 may be used to couple adjacent disks 221 together. In one embodiment forward arm 225 and aft arm 226 circumferentially extend completely around disk 221. Forward arm 225 extends axially forward and aft arm 226 extends axially aft. For ease of explanation the disks 221, forward arms 225, and the aft arms 226 shown in FIG. 3 are labeled with a, b, c, or d based on their relative positions in the compressor with the forward most disk 221 shown denoted as 221a and the aft most disk 221 shown denoted as 22 Id.

Each disk 221 couples to an adjacent disk 221. For example, disk 221b is located adjacent and axially forward of disk 221c. Disk 221b couples to disk 221c. The forward arm 225c of disk 221c radially aligns with the aft arm 226b of disk 221b. Forward arm 225c is coupled to aft arm 226b. In one embodiment each forward arm 225 and each aft arm 226 may include curvic teeth 219.

Airfoils 227 couple to disks 221 at rim 222. Each airfoil 227 includes a base (not shown) with a retaining feature such as a fir tree or a dovetail. Slots (not shown) in rim 222 have a corresponding retaining feature that secures each airfoil 227 to disk 221.

Rear hub 245 includes hub arm 246. Hub arm 246 extends axially forward from and extends circumferentially around rear hub 245. Hub arm 246 radially aligns with and is configured to couple to rear arm 226d of disk 22 Id. In one embodiment hub arm 246 includes curvic teeth 219 to form the coupling between rear hub 245 and disk 22 Id. Rear hub 245 is configured to form shaft cavity 247. Shaft cavity 247 is a cylindrical cavity at the axis of rear hub 245 configured to receive shaft 120. Each spacer 230 is shaped generally as a hollow cylinder or annular ring. Spacers 230 span between adjacent disks 221 and couple to adjacent rims 222 with a press fit, slip fit, or interference fit. In one embodiment, the forward end of the spacers 230 couple to an adjacent disk 221 with a slip fit, while the aft end of the spacers couple to an adjacent disk 221 with a press fit. In another embodiment, the forward end of the spacers 230 couple to an adjacent disk 221 with a press fit, while the aft end of the spacers couple to an adjacent disk 221 with a slip fit. Spacers 230 are located radially inward from stators 250. In FIG. 3 one of the spacers 230 is shown as an undercut spacer 231. However, others of spacers 230 may also be undercut spacers 231.

FIG. 4 is a cross-sectional view of a portion of the undercut spacer

231. Referring now to FIG. 3 and FIG. 4, all references to radial, axial, and circumferential directions and measures defining elements of undercut spacer 231 refer to the axis of undercut spacer 231 which is concentric to center axis 95. Undercut spacer 231 is generally shaped as a hollow cylinder or annular ring. Undercut spacer 231 spans between adjacent disks 221 and couples to adjacent rims 222 with a press fit or interference fit. Undercut spacer 231 is located radially inward from stators 250. In one embodiment, each of the features of undercut spacer 231 described below circumferentially extends completely around undercut spacer 231.

Undercut spacer 231 includes cylindrical body 229, outer surface

232, forward face 233, aft face 234, and inner surface 243. Body 229 may be a hollow cylinder or annular ring. Outer surface 232 is the radially outermost circumferential surface of body 229. Forward face 233 is a ring like surface extending radially inward from an axially forward edge of outer surface 232 towards the axis of undercut spacer 231. Forward face 233 may be perpendicular to the axis of undercut spacer 231. Aft face 234 is a ring like surface extending radially inward from an axially aft edge of outer surface 232. Aft face 234 is distal to forward face 233. Aft face 234 may be perpendicular to the axis of undercut spacer 231. Inner surface 243 is located radially inward from outer surface 232. Inner surface 243 may also be located radially inward from a radially innermost edge of forward face 233, and a radially innermost edge of aft face

234. Inner surface 243 may be the radially innermost circumferential surface of body 229.

Undercut spacer 231 may also include thickened portion 241, forward lip 237, and aft lip 238. Thickened portion 241 may protrude radially inward from a forward portion of body 229. In one embodiment, thickened portion 241 extends from a forward half of body 229. Thickened portion 241 may be axially forward of inner surface 243 and extend radially inward beyond inner surface 243. In the embodiment shown in FIG. 3 and FIG. 4, the interface between thickened portion 241 and inner surface 243 includes fillet 242 at the axially aft end of thickened portion 241. In another embodiment, inner surface 243 spans the entirety of undercut spacer 231 with the entirety of undercut spacer 231 thickened.

Forward lip 237 extends axially forward from body 229. A portion of forward lip 237 may extend from thickened portion 241. Forward lip 237 may include forward lip surface 239, the radially outer circumferential surface of forward lip 237. Forward lip 237 and forward lip surface 239 may be radially inward and may be axially forward of forward face 233.

Undercut spacer 231 is configured to include forward undercut

235, a circumferential recess. The circumferential recess may have a curved or arc shaped profile. Forward undercut 235 recedes aft into body 229 from forward face 233. In one embodiment, forward undercut 235 recedes from the radially inner edge of forward face 233 and the axially aft edge of forward lip surface 239. In another embodiment, forward undercut 235 also recedes radially inward from forward lip surface 239. In the embodiment depicted in FIG. 3 and FIG. 4, forward lip surface 239 is tangent to the arc of forward undercut 235. Forward undercut 235 may be of various heights, depths, and radii. The height of forward undercut 235 may be equal to the radial distance from the radial inner edge of forward face 233 to forward lip surface 239. In one embodiment, the height of forward undercut 235 is between 0.2 inches and 0.3 inches. In another embodiment, the height of forward undercut 235 is 0.24445 inches. In yet another embodiment, the ratio between a height of forward undercut and the thickness of the forward end of the spacer is between forty percent and fifty percent.

The profile of forward undercut 235 may include multiple radii. In one embodiment, forward undercut 235 includes a first radius and a second radius. The first radius may be 0.050 inches and the second radius may be 0.225 inches. Forward lip 237 may be tangent to the second radius.

Aft lip 238 extends axially aft from body 229. Aft lip surface 240 is the radially outer circumferential surface of aft lip 238. Aft lip 238 and aft lip surface 240 are located radially inward from aft face 234.

Undercut spacer 231 is configured to include aft undercut 236, a circumferential recess. The circumferential recess may have a curved or arc shaped profile. Aft undercut 236 recedes forward into body 229 from aft face 234. In one embodiment, aft undercut 236 recedes from the radially inner edge of aft face 234 and the axially forward edge of aft lip surface 240. In another embodiment, aft undercut 236 also recedes radially inward from aft lip surface 240. In the embodiment depicted in FIG. 3 and FIG. 4, aft lip surface 240 is tangent to the arc of aft undercut 236, and aft undercut 236 is smaller than forward undercut 235. In one embodiment, aft undercut 236 is between ten percent and twenty percent the size of forward undercut 235. In another embodiment, aft undercut 236 is between sixteen percent and seventeen percent the size of forward undercut 235.

Aft undercut 236 may be of various heights, depths, and radii. The height of aft undercut 236 may be equal to the radial distance from the radial inner edge of aft face 234 to aft lip surface 240. In one embodiment, the height of aft undercut 236 is between 0.02 inches and 0.06 inches. In another embodiment, the height of aft undercut 236 is 0.040 inches. The profile of aft undercut 236 may also include multiple radii. In yet another embodiment, the ratio between a height of aft undercut 236 and the thickness of the aft end of the spacer is between five percent and fifteen percent.

In the embodiment depicted in FIG. 3 and FIG. 4, corners of forward lip 237, aft lip 238, and thickened portion 241 may be chamfered. Outer surface 232 may be tapered with the diameter of outer surface 232 gradually increasing while moving along outer surface 232 axially from forward face 233 to aft face 234.

When undercut spacer 231 is installed into compressor 200, outer surface 232 is adjacent to stators 250. In one embodiment, the forward end of undercut spacer 231 is installed with a slip fit and the aft end of undercut spacer 231 is installed with a radial press fit or interference fit. Forward face 233 is adjacent the aft extension face 216 and forward lip surface 239 is adjacent the aft extension surface 218 of the axially forward adjacent disk 221. Forward face 233 is generally parallel to aft extension face 216 with a small axial gap therebetween. Forward lip surface 239 may not contact aft extension surface 218 while the engine is cold. Aft face 234 contacts the forward extension face 215 and aft lip surface 240 contacts the forward extension surface 217 of the axially aft adjacent disk 221. Aft face 234 is generally parallel to forward extension face 215.

In another embodiment, the forward end of undercut spacer 231 is installed with a radial press fit or interference fit, and the aft end of undercut spacer 231 is installed with a slip fit. Forward face 233 contacts the aft extension face 216 and forward lip surface 239 contacts the aft extension surface 218 of the axially forward adjacent disk 221. Forward face 233 is generally parallel to aft extension face 216. Aft face 234 is adjacent the forward extension face 215 and aft lip surface 240 is adjacent the forward extension surface 217 of the axially aft adjacent disk 221. Aft face 234 is generally parallel to forward extension face 215 with a small axial gap therebetween. Aft lip surface 240 may not contact forward extension surface 217 while the engine is cold.

During operation of the gas turbine engine the radial fits of the forward and aft ends of undercut spacer 231 may vary between a slip fit and an interference fit. During operation of the gas turbine engine undercut spacer 231 and the adjacent disks 221 thermally expand. During thermal expansion the radial interference fits may increase, with a maximum radial interference fit being between 0.010" to 0.014". In another embodiment undercut spacer 231 is configured to have a maximum radial interference fits of 0.012" with adjacent disks 221.

Anti-rotation pins (not shown) may be installed between spacers 230 and disks 221. Each anti-rotation pin protrudes axially into rim 222 and spacer 230. In one embodiment each anti-rotation pin protrudes through a forward extension face 215 into a forward extension 223 and into the adjacent spacer 230. The anti-rotation pins protruding into undercut spacer 231 may protrude through aft face 234.

In the embodiment depicted in FIG. 3, undercut spacer 231 is axially forward and adjacent to the penultimate disk 221 of compressor 200 with the aft end of undercut spacer 231 installed against the penultimate disk 221. This may be the twelfth stage of compressor 200. However, undercut spacer 231 may be installed between any two disks 221 within compressor 200, especially in the later stages of compressor 200 and the forward end of undercut spacer 231 or the aft end of undercut spacer 231 may be installed against the adjacent aft or forward disks 221.

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, and the transportation industry.

Gas turbine engines operate at high temperatures and pressures.

During operation of gas turbine engines the components of gas turbine engines thermally expand and experience stresses and strains due to the thermal expansion. The stresses and strains are highest during the transients of gas turbine engines. More particularly, the peak stresses and strains may occur during the start-up of a gas turbine engine.

During start-up the gas turbine engine heats up rapidly. Each component may thermally expand at a different rate due to the geometry of the component, the coefficient of thermal expansion of the material of the component, and the location of the component within the gas turbine engine. The different thermal expansion rates of adjacent components may lead to high stresses and strains. Varying centrifugal force loads may also contribute to the high stresses and strains.

One such instance may occur within the compressor 200. Spacers 230 may thermally expand faster than the disks 221. Through research and testing it was determined that stress concentrations may occur within the spacers 230 due to the increased contact and interference between spacers 230 and the adjacent disks 221. The stresses may cause deformation of the spacers 230. At start-up the stresses may exceed the yield strength of the spacers 230 which may lead to a shortened life, cracking, or failure of the spacers 230. These stresses may be more likely to occur in the later stages of compressor 200 where the temperatures and pressures may be higher.

Through research and testing it was determined that adding forward undercut 235 and aft undercut 236 may reduce the stress concentrations. As undercut spacer 231 thermally expands between adjacent disks 221 the stresses may be more evenly distributed about forward undercut 235 and aft undercut 236. This may reduce the strains in undercut spacer 231 and may reduce stress concentrations below the yield strength of undercut spacer 231 and may reduce stress concentrations low enough to prevent low cycle fatigue. It was also determined that undercut spacer 231 may be thickened radially, which may provide structural support within undercut spacer 231.

Transients where the gas turbine engine moves from a higher temperature to a lower temperature may present different problems due to varying thermal contraction of gas turbine engine components. This may be especially true at shut-down of a gas turbine engine. As previously mentioned, spacers 230 may be installed to disks 221 with an interference fit. The spacers

230 may cool down and thermally contract faster than the disks 221. This may cause a loss of the pilot or interference fit between spacers 230 and disks 221. Anti-rotation pins installed between spacers 230 and disks 221 may shear and spacers 230 may rub against disks 221 damaging spacers 230 or disks 221. Loss of pilot may also allow spacers 230 to shift. A shift may cause spacers 230 to rub against stators 250 damaging spacers 230 or stators 250. A shift may also cause an imbalance in the compressor 200.

It was determined that the interference fit between undercut spacer

231 and an adjacent disk 221 may be increased due to the possible stress reductions of forward undercut 235 and aft undercut 236. The increase in interference fit may reduce the possibility of a loss of pilot as undercut spacer 231 thermally contracts. Radially thickening undercut spacer 231 may slow down thermal contraction and may further reduce the possibility of a loss of pilot as the rate of thermal contraction of undercut spacer 231 may be closer to the rate of thermal contraction of the adjacent disks 221.

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 a particular spacer and associated processes, it will be appreciated that other spacers 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.