Gas turbine core tie rod with reduced span

Abstract

Structures are described for achieving reduced span of tie rods and improved vibration mode margins in gas turbine engines. A gas turbine engine includes a tie rod assembly, a plurality of coupling nuts, a front shaft, a bladed disc, a threaded joint coupled to a tapered shaft of the bladed disc, a high pressure compressor rotor, and a high pressure turbine rotor including the tapered shaft. The first coupling nut is coupled to a tapered shaft of a high pressure compressor rotor. The second coupling nut is coupled to the front axle. The third coupling nut is coupled to an aft end stage of the high pressure turbine rotor.

Description

Technical Field

[0001] The present disclosure relates to gas turbine engines, and more particularly, to gas turbine engines including tie rod assemblies having a reduced span. Background Art

[0002] Gas turbine engines for commercial aircraft typically include a fan and a turbine. The turbine (often referred to as the core) typically includes a compressor section, a combustion section, and a turbine section arranged in a series flow arrangement. The compressor section compresses air, which is directed to the combustion section, where the air is mixed with fuel. The mixture is then ignited to generate hot combustion gases. The combustion gases are directed to the turbine section, which extracts energy from the combustion gases to power the compressor section and to generate work, such as for propulsion of the aircraft in flight or for powering machines such as generators.

[0003] Current tie-rod architectures, including increasing the diameter of the tie-rods, can improve vibration mode margins. However, there are limits to the tie-rods and the added weight of the rotor. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] The embodiments illustrated in the drawings are illustrative and exemplary in nature and are not intended to limit the subject matter defined by the claims. The following detailed description of the illustrative embodiments may be understood when read in conjunction with the following drawings, in which like structures are represented by like reference numerals, and in which:

[0005] Figure 1 schematically depicts a cross-sectional view of a gas turbine engine according to one or more aspects described and illustrated herein;

[0006] Figure 2A Schematically depicts a device according to one or more aspects described and illustrated herein. Figure 1 a plurality of compressor stages of a gas turbine engine;

[0007] Figure 2B Schematically depicts a device according to one or more aspects described and illustrated herein. Figure 2A a detailed view of a portion of a compressor stage and a portion of a combustion section of a gas turbine engine;

[0008] Figure 3A Schematically depicts a system including a tie rod according to one or more aspects described and illustrated herein. Figure 1 a cross-sectional view of a gas turbine engine;

[0009] Figure 3B Schematically depicts a device according to one or more aspects described and illustrated herein. Figure 3A a portion of a cross-sectional view of a gas turbine engine;

[0010] Figure 4A Schematically depicts a vehicle including an integral front axle with a blisk according to one or more aspects described and illustrated herein. Figure 3A a cross-sectional view of a gas turbine engine;

[0011] Figure 4B Schematically depicts a device according to one or more aspects described and illustrated herein. Figure 4A a portion of a cross-sectional view of a gas turbine engine;

[0012] Figure 5A Schematically depicts a vehicle including an integral front axle with a tie rod according to one or more aspects described and illustrated herein. Figure 3A a cross-sectional view of a gas turbine engine;

[0013] Figure 5B Schematically depicts a device according to one or more aspects described and illustrated herein. Figure 5A a portion of a cross-sectional view of a gas turbine engine; and

[0014] Figure 6 Schematically depicts a system including a tie rod welded to a blisk according to one or more aspects described and illustrated herein. Figure 3A Cross-sectional view of a gas turbine engine. DETAILED DESCRIPTION

[0015] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and the description have been used to refer to like or similar parts of the disclosure.

[0016] Unless otherwise specified herein, the terms "coupled," "fixed," "attached," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0017] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0018] The term "at least one" in a context such as "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

[0019] The terms "fore" and "aft" refer to relative positions within a gas turbine engine, pump, or vehicle and refer to the normal operating attitude of the gas turbine engine, pump, or vehicle. For example, for a gas turbine engine, the front position refers to the position closer to the engine inlet, while the aft position refers to the position closer to the engine nozzle or exhaust.

[0020] The terms "upstream" and "downstream" refer to relative directions relative to the flow in a path. For example, with respect to fluid flow, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0021] As used in this application, stating that any portion (e.g., region) is in any manner on (e.g., positioned on, located on, disposed on, or formed on, etc.) another portion indicates that the referenced portion is in contact with the other portion, or that the referenced portion is located above the other portion with one or more intermediate portions located therebetween.

[0022] As used herein, unless otherwise indicated, connection references (e.g., attached, coupled, connected, and joined) may include intermediate members between the elements to which the connection reference refers and / or relative movement between those elements. Thus, connection references do not necessarily infer that two elements are directly connected and / or are in fixed relation to each other. As used herein, stating that any part is "in contact with" another part is defined to mean that there are no intermediate parts between the two parts.

[0023] Unless otherwise specifically stated, descriptors such as "first," "second," and "third" as used herein do not confer or otherwise indicate any meaning of priority, physical order, arrangement in a list, and / or ordering in any manner, but are merely used as labels and / or arbitrary names to distinguish elements to facilitate understanding of the disclosed examples. In some examples, the descriptor "first" may be used to refer to an element in the detailed description, while a different descriptor (such as "second" or "third") may be used in the claims to refer to the same element. In such cases, it should be understood that such descriptors are only used to clearly identify those elements that might otherwise share the same name, for example.

[0024] As used herein throughout the specification and claims, approximating language is applied to modify any quantitative representation that can permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, values ​​modified by terms such as "about" and "substantially" are not limited to the precise values ​​specified. In at least some instances, approximating language may correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximating language may mean within a 10% margin.

[0025] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0026] Some types of tie-rod architectures may include a configuration with tie-rods and cap-shaped springs. The cap-shaped springs may be configured to provide mid-span support to improve vibration mode margins. The cap-shaped springs may crack at one location and cause radially asymmetric loading, leading to operational issues such as rotor imbalance and rotor vibration during operation. Therefore, these types of tie-rod architectures may be configured to include tie-rods with thickened ends to improve vibration mode margins without the cap-shaped springs.

[0027] Furthermore, other types of tie-rod architectures may include configurations with tie rods having a higher length-to-diameter (L / D) ratio than the aforementioned types of tie-rod architectures. This higher L / D ratio reduces the tie-rod modal margin, dropping it below the 20% margin requirement at core redline speed. As with the aforementioned types of tie-rod architectures, the option of using thickened tie-rod ends does not improve modal margins, thus requiring mid-span bracing.

[0028] While some methods of improving vibration mode margins include increasing the tie rod diameter, doing so results in not only an increase in tie rod weight, but also an increase in the overall rotor weight of the high-pressure compressor rotor / high-pressure turbine rotor, or increasing the axial span of the final stage (such as multiple compressor stages) of the high-pressure compressor rotor to accommodate weld cleaning.

[0029] Referring now to the accompanying drawings, Figure 1 A schematic cross-sectional view of a turbofan engine 100 is provided in accordance with an exemplary embodiment of the present disclosure. Figure 1 In the depicted embodiment, the turbofan engine 100 is an aviation high-bypass turbofan engine configured to be mounted to an aircraft, for example, in an underwing configuration. As shown, the turbofan engine 100 defines an axial direction A, a radial direction R, and a circumferential direction C. The axial direction A extends parallel to or coaxially with a longitudinal centerline 102 defined by the turbofan engine 100.

[0030] The turbofan engine 100 includes a fan section 104 and a core turbine engine 106 disposed downstream of the fan section 104. The core turbine engine 106 includes a cowling 108 defining an annular core inlet 110. The cowling 108 encloses, in series flow relationship, a compressor section 112 including a first supercharger (e.g., LP compressor 114) and a second supercharger (e.g., HP compressor 116); a combustion section 118; a turbine section 120 including a first turbine (e.g., HP turbine 122) and a second turbine (e.g., LP turbine 124); and an exhaust section 126. The compressor section 112, the combustion section 118, the turbine section 120, and the exhaust section 126 together define a core air flow path 132 through the core turbine engine 106.

[0031] HP shaft 128 drivingly connects HP turbine 122 to HP compressor 116. LP shaft 130 drivingly connects LP turbine 124 to LP compressor 114. HP shaft 128, rotating components of HP compressor 116 mechanically coupled to HP shaft 128, and rotating components of HP turbine 122 mechanically coupled to HP shaft 128 collectively form a high-pressure spool, or HP spool 131. LP shaft 130, rotating components of LP compressor 114 mechanically coupled to LP shaft 130, and rotating components of LP turbine 124 mechanically coupled to LP shaft 130 collectively form a low-pressure spool, or LP spool 133.

[0032] Fan section 104 includes a fan assembly 138 having a fan 134 mechanically coupled to a fan rotor 140. Fan 134 has a plurality of fan blades 136 circumferentially spaced apart from one another. As depicted, fan blades 136 extend outwardly from fan rotor 140 in a radial direction R. A power gearbox 142 mechanically couples LP spool 133 and fan rotor 140. Power gearbox 142 may also be referred to as a main gearbox. Power gearbox 142 includes a plurality of gears for reducing the rotational speed of LP shaft 130 to provide a more efficient rotational fan speed for fan 134. In other exemplary embodiments, fan blades 136 of fan 134 may be mechanically coupled to a suitable actuating member configured to pitch fan blades 136 about respective pitch axes, for example, in unison. In some alternative embodiments, turbofan engine 100 does not include power gearbox 142. In such alternative embodiments, fan 134 may be mechanically coupled directly to LP shaft 130, for example, in a direct drive configuration.

[0033] Still refer to Figure 1, the fan rotor 140 and the hub of the fan blades 136 are covered by a rotatable spinner 144 that is aerodynamically shaped to facilitate airflow through the plurality of fan blades 136. In addition, the fan section 104 includes an annular fan casing 145 and an outer nacelle 146 connected to the fan casing 145. The fan casing 145 and the outer nacelle 146 both circumferentially surround the fan 134 and / or at least a portion of the core turbine engine 106. The fan casing 145 and the outer nacelle 146 are supported relative to the core turbine engine 106 by a plurality of circumferentially spaced outlet guide vanes 148. A downstream section 150 of the nacelle 146 extends above the outer portion of the core turbine engine 106 to define a bypass passage 152 therebetween.

[0034] During operation of turbofan engine 100, a volume of air 154 enters turbofan engine 100 through nacelle 146 and / or an associated inlet 156 of fan section 104. As volume of air 154 passes through fan blades 136, a first portion of air 158 is directed or channeled into bypass passage 152, and a second portion of air 160 is directed or channeled into annular core inlet 110. As second portion of air 160 flows downstream through LP compressor 114 and HP compressor 116, the pressure of second portion of air 160 gradually increases. Specifically, LP compressor 114 includes sequential stages of LP compressor stator vanes 182 and LP compressor blades 184 that progressively compress second portion of air 160. LP compressor blades 184 are mechanically coupled to LP shaft 130. Similarly, HP compressor 116 includes sequential stages of HP compressor blades 186 and HP compressor blades 188 that further progressively compress second portion of air 160. HP compressor blades 188 are mechanically coupled to HP shaft 128. Additional details regarding the various components of the LP compressor 114 and the HP compressor 116 are described in greater detail below. The compressed second portion of air 160 is then discharged from the compressor section 112 into the combustion section 118 .

[0035] The compressed second portion of air 160 discharged from compressor section 112 is mixed with fuel and combusted within the combustor of combustion section 118 to provide combustion gases 162. Combustion gases 162 are directed from combustion section 118 along a hot gas path 174 of core air flow path 132 through HP turbine 122, where a portion of the thermal and / or kinetic energy from combustion gases 162 is extracted via sequential stages of HP turbine stator vanes 164 and HP turbine blades 166. HP turbine blades 166 are mechanically coupled to HP shaft 128. Thus, as HP turbine blades 166 extract energy from combustion gases 162, HP shaft 128 rotates, which supports operation of HP compressor 116. Combustion gases 162 are directed through LP turbine 124, where a second portion of the thermal and kinetic energy is extracted from combustion gases 162 via sequential stages of LP turbine stator vanes 168 and LP turbine blades 170. LP turbine blades 170 are coupled to LP shaft 130. Thus, as the LP turbine blades 170 extract energy from the combustion gases 162 , the LP shaft 130 rotates and supports the operation of the LP compressor 114 and, through the power gearbox 142 , the fan 134 .

[0036] Combustion gases 162 exit LP turbine 124 and are exhausted from core turbine engine 106 through exhaust section 126 to provide propulsive thrust. Simultaneously, the pressure of first portion air 158 is significantly increased as it is directed through bypass passage 152 before being exhausted from fan nozzle exhaust section 172 of turbofan engine 100, also providing propulsive thrust. HP turbine 122, LP turbine 124, and exhaust section 126 at least partially define a hot gas path 174.

[0037] It should be understood that Figure 1 The turbofan engine 100 depicted in FIG is provided as an example, and in other example embodiments, the turbofan engine 100 has other configurations. Additionally or alternatively, aspects of the present disclosure may be used with other suitable aviation turbofan engines, turboshaft engines, and turboprop engines.

[0038] Now refer to Figure 2A , provides Figure 1 1 is a schematic cross-sectional view of a portion of the compressor section 112 and a portion of the combustion section 118 of the turbofan engine 100. More specifically, Figure 2A The aft end of the HP compressor 116 and a portion of the combustion section 118 of the compressor section 112 are depicted. However, it should be understood that the various components described herein may be included in other compressor sections of the turbofan engine 100, including the LP compressor 114 and / or the intermediate pressure (IP) compressor in a 3-spool gas turbine engine.

[0039] refer to Figure 1 and Figures 2A-2B As described above, during operation of the turbofan engine 100, the airflow through the core air flow path 132 of the turbofan engine 100 is compressed sequentially as it flows through the compressor section 112 (or, more specifically, as it flows through the LP compressor 114 and the HP compressor 116). The compressed air from the compressor section 112 is then provided to the combustion section 118, where at least a portion of the compressed air is mixed with fuel and combusted to produce combustion gases 162. The combustion gases 162 flow from the combustion section 118 to the turbine section 120, and more specifically, sequentially through the HP turbine 122 and the LP turbine 124, which, for the depicted embodiment, drive the HP turbine 122 and the LP turbine 124. The HP spool 131 is drivingly coupled to the HP turbine 122 and the HP compressor 116.

[0040] Special References Figure 2A , the HP compressor 116 includes a plurality of compressor stages 202a-202e (collectively, compressor stages 202), wherein each of the compressor stages 202 includes, for example, a plurality of HP compressor blades 188 and a rotor 206. Figure 2A Five compressor stages 202 are depicted in FIG. 1 , but in other embodiments, the HP compressor 116 includes more or fewer than five stages. Each of the compressor stages 202 is drivingly coupled to the HP spool 131 such that the HP turbine 122 ( Figure 1 ) may drive the HP compressor 116 via the HP spool 131 . Among the plurality of compressor stages 202 of the HP compressor 116 , a last stage 202 a is located at the rear end 200 of the HP compressor 116 .

[0041] The final stage 202a provides compressed air to the combustion section 118. More specifically, for Figure 2A In the embodiment depicted in FIG, the combustion section 118 includes a diffuser 230, an inner combustor casing 232, and a combustor assembly 234. In addition, the combustion section 118 defines a diffuser cavity 236, wherein the diffuser 230 is located downstream of the compressor stage 202 of the HP compressor 116 and upstream of the diffuser cavity 236, such that compressed air from the last stage 202a is provided to the diffuser cavity 236 through the diffuser 230. The compressed air within the diffuser cavity 236 is then provided to the combustor assembly 234, where the compressed air is mixed with fuel and combusted to generate the combustion gases 162. Figure 2A As depicted in FIG, combustor assembly 234 generally includes a fuel nozzle 240 , an inner liner 242 , and an outer liner 244 , where the inner liner 242 and the outer liner 244 together form a combustion chamber 250 .

[0042] It should be understood that the combustor assembly 234 is configured as a turbofan engine 100 ( Figure 1 For example, in some embodiments, the combustor assembly 234 is configured as an annular combustor assembly, a can combustor assembly, or a can-annular combustor assembly.

[0043] Still refer to Figure 2A As previously described, the HP spool 131 is drivingly connected to the HP compressor 116. For the depicted embodiment, the HP spool 131 generally includes a central spool section including a central spool member 208, which may also be referred to herein as an inner peripheral support structure. Figure 2A , the centerline shaft member 208 extends generally along the axial direction A at a location radially inward of the combustor assembly 234 of the combustion section 118. Additionally, the centerline shaft member 208 is coupled to or integrally formed with one or more spacer arms 210 located forward of the centerline shaft member 208. For the depicted embodiment, the one or more spacer arms 210 also extend generally along the axial direction A. The centerline shaft member 208 and the one or more spacer arms 210 may together form an inner peripheral support structure 209 for the HP compressor 116.

[0044] Still refer to Figure 2A , the last stage 202a of the HP compressor 116 represents the last stage of the HP compressor 116 when traversing the HP compressor 116 from a front position to a rear position in the axial direction A. One or more preceding stages 202b-202e located before the last stage 202a include, for example, a first preceding stage 202b, a second preceding stage 202c, a third preceding stage 202d, and a fourth preceding stage 202e. Each of the compressor stages 202a-202f includes corresponding compressor blades and compressor vanes in the HP compressor buckets 186 and the HP compressor blades 188. That is, the last stage 202a includes the last blade 186a (e.g., the first blade) and the first compressor blade 188a, the first front stage 202b includes the second blade 186b and the second compressor blade 188b, the second front stage 202c includes the third blade 186c and the third compressor blade 188c, the third front stage 202d includes the fourth blade 186d and the fourth compressor blade 188d, and the fourth front stage 202e includes the fifth blade 186e and the fifth compressor blade 188e, and so on (e.g., the sixth blade 186f and the sixth compressor blade 188f, etc.).

[0045] The HP compressor 116 also includes an outer shell 204, which may also be referred to herein as an outer peripheral support structure. The outer shell 204 may extend generally radially outward of the inner peripheral support structure 209 in the axial direction A. In some embodiments, the outer shell 204 and the inner peripheral support structure 209 are arranged around a central axis (e.g., the turbofan engine 100 ( Figure 1 ) is positioned along the longitudinal centerline 102). That is, the inner peripheral support structure 209 is positioned along the longitudinal centerline 102 ( Figure 1 ), and the outer shell 204 is spaced radially outward from the inner peripheral support structure 209, as shown Figure 2A Depicted in.

[0046] refer to Figure 2A and Figure 2B , each of the compressor's buckets 186 generally extends a distance inwardly from the outer casing 204 in a radial direction R. Each of the buckets 186 extends from the outer casing 204 at a location between adjacent compressor blades 188. For example, the last bucket 186a may extend from the outer casing 204 at a location between the first compressor blade 188a and the second compressor blade 188b. In addition, each bucket 186 extends toward the inner peripheral support structure 209, and in particular, toward one of the one or more spacer arms 210 thereof. In an embodiment, one or more components are disposed between the bucket 186 and the corresponding spacer arm 210, such as the inner platform 282, the seal support structure 284, the seal structure 286, and / or one or more seal teeth 260, as described in more detail herein.

[0047] Special References Figure 2B , which schematically depicts Figure 2AIn the enlarged view of portion 2B in FIG, each of the blades 186 (e.g., the last blade 186a, the second blade 186b, etc.) includes a root 262, a tip 264, a leading edge 268, and a trailing edge 266. The root 262 of each blade 186 represents the radially outward extension of the blade 186 at the connection point with the casing 204. In other words, the root 262 of each blade 186 is the portion (e.g., the end) of the blade 186 that contacts the casing 204. The tip 264 of each blade 186 represents the radially inward extension of the blade 186. In other words, the tip 264 of each blade 186 is the portion (e.g., the end) of the blade closest to the corresponding spacer arm 210. The leading edge 268 of each blade 186 represents the edge of the blade 186 that extends from the root 262 to the tip 264 and is the generally forward-most edge of the blade 186 in the axial direction (e.g., the edge that receives fluid flowing through the HP compressor 116, as described herein). Trailing edge 266 of each vane 186 represents the edge of vane 186 extending from root 262 to tip 264 and is the rearmost edge of vane 186 generally in the axial direction. Thus, trailing edge 266 and leading edge 268 are opposed to one another. In some embodiments, trailing edge 266 and leading edge 268 are parallel or substantially parallel to one another. In other embodiments, trailing edge 266 and leading edge 268 are not parallel to one another.

[0048] like Figure 2B , each of vanes 186 defines a first point 272 and a second point 274. First point 272 represents the intersection of tip 264 of vane 186 and trailing edge 266 of vane 186. Second point 274 represents the intersection of root 262 of vane 186 and trailing edge 266 of vane 186.

[0049] As described herein, one or more components may be disposed between the tip 264 of each bucket 186 and the corresponding spacer arm 210, including, for example, an inner platform 282, a seal support structure 284, a seal structure 286, and / or one or more seal teeth 260. In an embodiment, the inner platform 282, the seal support structure 284, the seal structure 286, and the one or more seal teeth 260 occur in a serial order from the tip 264 to the corresponding spacer arm 210, wherein the inner platform 282, the seal support structure 284, and the seal structure 286 are coupled to one another, and the tip 264 of each bucket 186 and the one or more seal teeth are disposed on a radially outer surface 294 of the spacer arm 210.

[0050] The inner platform 282 is a component that defines the flow path. That is, through the compressor stage 202 ( Figure 2A) occurs via a flow path defined by inner platforms 282. Inner platforms 282 are coupled to tips 264 of blades 186 and extend inwardly in radial direction R from tips 264 of blades 186. As will be appreciated, inner platforms 282 have shapes and surface features that are not necessarily limited to those disclosed in the examples. For example, inner platforms 282 may be shaped to correspond to the shape of tips 264 of blades 186 and / or may be shaped to flare outwardly in axial direction A relative to the width of blades 186 (e.g., the dimension extending from leading edge 268 to trailing edge 266 of blades 186). Each inner platform 282 may differ in shape, size, and configuration relative to the other inner platforms 282, or may be substantially identical in shape, size, and configuration to the other inner platforms 282.

[0051] The inner platform 282 also defines the core air flow path 132 ( Figure 1 As described in more detail herein, the specific dimensions of the inner platform 282 direct the air from the core air flow path 132 ( Figure 1 ) of air. While still maintaining the flow path hub, the angle of the high pressure rear cone arm relative to the longitudinal centerline 102 ( Figure 1 ) is reduced, which makes the life of each component longer.

[0052] The seal support structure 284 is generally a component coupled to the inner platform 282 and disposed inwardly in the radial direction R of the inner platform 282. The seal support structure supports a seal structure 286 thereon. The seal structure 286 is generally any component that prevents or minimizes leakage of fluid from the flow path defined by the inner platform 282. In other words, the seal structure 286 is used to maintain fluid flow within the flow path defined by the inner platform 282. Figure 2B In the depicted embodiment, the seal structure 286 is an abradable honeycomb seal. That is, the seal structure 286 is a machined component with separate chambers that create a pressure drop to mitigate leakage and / or damage to the HP shaft 128 ( Figure 1 The sealing structure 286 forms a seal with the sealing teeth 260 provided on the radially outer surface 294 of the spacer arm 210.

[0053] It should be understood that Figure 2B The seal structure 286 depicted in FIG2 is not limited to an abradable honeycomb seal. For example, in other embodiments, the seal structure 286 is a bridge seal, a rod seal, a cassette seal, an attached seal ring housing, a foil seal, a brush seal, an advanced air-breathing seal, etc. In some embodiments, the selection of the seal structure 286 depends on the size of the interstage seal (ISS) cavity defined by the spacer arm 210, the adjacent rotor 206, and the casing 204.

[0054] Reference again Figure 2A and Figure 2B , the spacer arms 210 are generally positioned a distance inward from the casing 204 in the radial direction R to define the space for each of the compressor stages 202 (including its buckets 186 and HP compressor blades 188). The spacer arms 210 of the final stage 202a define points 292 that are centrally located at the intersection of the spacer arms 210 and each rotor 206 that defines the final stage 202a. As will be described in greater detail herein, a first line 291 drawn through the two points 292 forms an angle θ with a second line 293 that is parallel to the longitudinal centerline 102 (e.g., in some embodiments, extending through at least one midpoint 290 positioned equidistant from the trailing edge 266 and the leading edge 268 at the root 262 of the bucket 186). The angle θ may be referred to as the spacing angle. It should be understood that because each spacer arm 210 may have a different slope, each compressor stage 202 may have a corresponding spacing angle that is different from the spacing angles of adjacent or nearby spacer arms. Thus, Figure 2B The angle θ depicted in φ is referred to as the interval angle of the last stage 202a.

[0055] As previously described herein, the spacer arm 210 includes a radially outer surface 294 and a radially inner surface 296. The radially inner surface 296 is opposite the radially outer surface 294. The radially outer surface 294 of the spacer arm 210 generally faces the bucket 186 and, in some embodiments, supports one or more seal teeth 260 coupled thereto. The spacer arm 210 generally defines a thickness in the radial direction R between the radially outer surface 294 and the radially inner surface 296. Additionally, the spacer arm 210 defines a midpoint 211 on the radially inner surface 296, the midpoint 211 being positioned equidistant between adjacent points 292, as shown in FIG. Figure 2B Depicted in.

[0056] As will be described in further detail herein, the first radial distance Ch is defined by the distance in the radial direction R between the first point 272 and the midpoint 211 on the radial inner surface 296 of the corresponding spacer arm 210. That is, the first radial distance Ch represents the distance including all components disposed between the tip 264 of the blade 186 and the corresponding spacer arm 210 (in some examples, including the inner platform 282, the seal support structure 284, the seal structure 286, one or more seal teeth 260, and the thickness of the spacer arm 210). This first radial distance Ch may also be referred to as the cavity height. As will be described in further detail herein, the second radial distance Vh is defined by the distance in the radial direction R between the first point 272 and the second point 274. The second radial distance Vh also represents the height of the blade 186 and may be referred to as the blade height. In addition, with reference to Figure 2A, the third radial distance Rh is defined by the distance in the radial direction R between the first point 272 and the longitudinal centerline 102 of the engine.

[0057] Now refer to Figure 3A , depicting the Figure 1 2. Another cross-sectional view of the turbofan engine 100 is shown with a pull rod 207. The turbofan engine 100 may include a plurality of compressor stages 202 (such as a plurality of compressor stages 202a-202h), a plurality of rotors 206, a plurality of spacer arms 210, a blisk 304, one or more seal teeth 260, and a plurality of airfoils 306, such as one or more buckets 186, each including a leading edge 268. Figure 3A Further depicted are a first nut 212 , a second nut 214 , a third nut 216 , a front shaft 218 , and a threaded engagement 220 . Figure 3A You can refer to and combine the above Figure 1 and Figures 2A-2B any component of the turbofan engine 100 as explained. Although depicted Figure 3A 1 is a single example of a component of turbofan engine 100, but it will be understood that any number of components may be included.

[0058] Figure 3A The dashed box 3B in the figure corresponds to Figure 3B In some examples, the threaded joint 220 can be configured to engage the internal threads of the front shaft 218 with the external threads of the tie rod 207. In some examples, the threaded joint 220 can be directly coupled to the tie rod 207 without any intermediate portions. In other examples, the threaded joint 220 can be indirectly coupled to the tie rod 207, such as via one or more intermediate portions. For example, dividing the tie rod 207 into three or more loops using the front shaft 218 can reduce the unsupported length of the tie rod 207, which helps improve the tie rod vibration mode margin for the first bending mode. For example, the three-loop configuration of the tie rod 207 includes a first nut 212 on the threaded joint 220 of the high-pressure compressor rotor tapered shaft and blisk 304 (such as the blisk tapered shaft 224), a second nut 214 on the front shaft 218, and a third nut 216 at the rear end of the high-pressure turbine rotor. In some examples, each of the first nut 212, the second nut 214, and the third nut 216 can include a coupling nut. As an example, the coupling nut may be circumferentially shaped, but is not limited to this configuration. The front shaft 218 may comprise IN718 alloy and may be configured to improve low cycle fatigue at the thread fillet.

[0059] Dividing the high pressure tie rod rotor assembly (such as tie rod 207) into multiple clamping rings (e.g., three rings) where the front shaft 218 is connected to the blisk 304 shortens the effective span of the tie rod. This division helps to reduce the tie rod 207 L / D ratio, thereby improving vibration mode margins. In addition, the division provides the turbofan engine 100 with higher interface loads and higher torque carrying capacity, and reduces the high pressure span, resulting in improved high pressure, low pressure dynamics. Further, the specially designed front shaft is constructed to improve low cycle fatigue at the thread fillet. In addition, the blisk 304, such as the high pressure compressor blisk, can be replaced without disassembling the core rotor.

[0060] Dividing the tie rod 207 in the high pressure compressor module generates higher clamping load and torque carrying capacity at the rear stage 202 of the high pressure turbine rotor. This enables the friction joint to be maintained, resulting in a reduced axial length of the high pressure compressor rotor. In some examples, the joint can include a friction joint, a curved coupling, an induction weld (IW), can be bolted, or other situations that can be used with a tie bolt rotor. The outer diameter of the shaft 219 of the tie rod 207 can be smaller than the inner diameter of the high pressure turbine rotor 221, so that the tie rod 207 enters the core from the front end.

[0061] Figure 3B At least one of the plurality of compressor stages 202 depicted in FIG includes a blisk 304. That is, at least one of the plurality of compressor stages 202 includes a disk with integral / welded blades, rather than other forms of blade-to-disk attachment, such as axial or circumferential dovetails, bolted connections, or pinned connections. These are different combinations / types of blade attachments that may be used interchangeably at at least one of the plurality of compressor stages 202 or any other stage of the compressor.

[0062] In some examples, airfoils 306 may be connected to blisks 304. Airfoils 306 may include a trapezoidal or similar shape. However, it should be understood that airfoils 306 are not limited to this shape, and any shape for airfoils 306 may be used.

[0063] The compressor of turbofan engine 100 may include a plurality of compressor stages 202. For example, the plurality of compressor stages 202 may include ten compressor stages. However, it should be understood that the plurality of compressor stages 202 is not limited to this number of compressor stages, and any number of compressor stages 202 may be used. As an example, a first portion (202a-202e) of the plurality of compressor stages 202 may include blisks 304, the first portion including five compressor stages. A second portion (202f-202j) may include circumferential dovetail bladed disks, the second portion including five compressor stages. It should be understood that the first predetermined number of disks of the first type and the second predetermined number of disks of the second type are not limited to these types and / or numbers of disks, and other types of disks may be used. For example, the first predetermined number of disks of the first type may be less than the second predetermined number of disks of the second type. In another example, the first predetermined number of disks of the first type may be equal to the second predetermined number of disks of the second type. In yet another example, the first predetermined number of disks of the first type may be greater than the second predetermined number of disks of the second type. Furthermore, the plurality of compressor stages 202 may be welded together.

[0064] Figure 4A Schematically depicts a front axle 218 having an integral blade disk 304. Figure 3A A cross-sectional view of a turbofan engine 100 . Figure 4A You can refer to and combine the above Figure 1 、 Figures 2A-2B and Figures 3A-3B any component of the turbofan engine 100 as explained. Although depicted Figure 4A 1 is a single example of a component of turbofan engine 100, but it will be understood that any number of components may be included. Figure 4A An integrated front axle 218 is depicted with a blisk 304 together as a single-piece unitary component, which in certain embodiments may be formed, but is not limited to, by heat application, pressure application, additive manufacturing, or any combination thereof. Figure 4A The dashed box 4B in the figure corresponds to Figure 4B . The multiple compressor stages 202 ahead of the tapered shaft 224 of the high pressure compressor rotor assembly use nuts, such as the second nut 214. In some examples, these multiple compressor stages 202 can include a spool / friction joint configuration. In some examples, the joint can include a friction joint, a curved coupling, an IW, a bolted joint, or other situations that can be used with a tie bolt rotor. In addition, all subsequent multiple compressor stages 202 of the high pressure compressor rotor assembly can use the threaded joint 220 entering the forward shaft 218 and the first nut 212. Final assembly of the core can use the third nut 216 at the rear stage of the high pressure turbine rotor 222.

[0065] Figure 5ASchematically depicts a front axle comprising an integral part with tie rods Figure 3A A cross-sectional view of a turbofan engine 100 . Figure 5A You can refer to and combine the above Figure 1 、 Figures 2A-2B 、 Figures 3A-3B and Figures 4A-4B any component of the turbofan engine 100 as explained. Although depicted Figure 5A 1 is a single example of a component of turbofan engine 100, but it will be understood that any number of components may be included. Figure 5A The front axle 218 and tie rod 207 are depicted together as a one-piece, unitary component. Figure 5A The dashed box 5B in the figure corresponds to Figure 5B . The integral front shaft 218 and tie rod 207 can be constructed to engage with the blisk 304 via a threaded joint 220 without the need for tabs. The supporting thread loading surface on the tie rod 207 is oriented forward. Multiple compressor stages 202 in front of the tapered shaft 224 can be assembled using the second nut 214. In some examples, these multiple compressor stages 202 can include a spool / friction joint construction. In some examples, the joint can include a friction joint, a curved coupling, an IW, a bolted joint, or other situations that can be used with a tie bolt rotor. In addition, all subsequent multiple compressor stages 202 of the high pressure compressor rotor assembly can use the first nut 212. Final assembly of the core can use the third nut 216 at the rear stage of the high pressure turbine rotor 222.

[0066] Figure 6 Schematically depicts a structure including a tie rod 207 welded to a blisk 304. Figure 3A A cross-sectional view of a turbofan engine 100 . Figure 6 You can refer to and combine the above Figure 1 、 Figures 2A-2B 、 Figures 3A-3B 、 Figures 4A-4B and Figures 5A-5B any component of the turbofan engine 100 as explained. Although depicted Figure 6 1 is a single example of a component of turbofan engine 100, but it will be understood that any number of components may be included. Figure 6The tie rod 207 is depicted as being welded as a single-piece component at the joint with the blisk 304. The plurality of compressor stages 202 preceding the tapered shaft 224 of the high-pressure compressor rotor assembly may utilize the second nut 214. In some examples, these plurality of compressor stages 202 may include a spool / friction joint configuration. In some examples, the joint may include a friction joint, a curved coupling, an IW, a bolted joint, or other configurations that may be used with a tie-bolt rotor. Furthermore, all subsequent plurality of compressor stages 202 of the high-pressure compressor rotor assembly may utilize the first nut 212. Final assembly of the core may utilize the third nut 216 at the subsequent stage of the high-pressure turbine rotor 222.

[0067] By way of example and not limitation, the tie rod assembly can be assembled by assembling or stacking a predetermined number of multiple compressor stages 202 at one or more joints (such as friction joints). The tie rod can be assembled from the rear and threadedly engaged with one of the front shafts 218 of the predetermined number of multiple compressor stages 202. As the high-pressure compressor rotor pushes, a tensile load from the rear (such as a hydraulic tensile load) can be applied to the tie rod and the first nut can be tightened. A second predetermined number of multiple compressor stages 202 can be assembled or stacked at another joint (such as another friction joint). Another tensile load from the rear (such as another hydraulic tensile load) can be applied to the front tapered shaft and the second nut can be tightened. The high-pressure turbine module can be assembled to the high-pressure compressor rotor assembly from the rear. Another tensile load from the rear (such as a third hydraulic tensile load) can be applied to the tie rod and the third nut can be tightened.

[0068] As can be understood from the above, this document defines a configuration in which the high-pressure tie rod rotor assembly is divided into multiple clamping rings (e.g., three rings), wherein the front shaft is connected to the blade disk, which reduces the effective span of the tie rod effective span. This division helps to reduce the tie rod L / D ratio, thereby improving the vibration mode margin. In addition, the division provides a higher interface load and higher torque carrying capacity for the turbofan engine, and reduces the high pressure span, resulting in improved high pressure, low pressure dynamics. Furthermore, the specially designed front shaft is constructed to improve low cycle fatigue at the thread fillet. In addition, the blade disk, such as the high pressure compressor blade disk, can be replaced without disassembling the core rotor.

[0069] Although specific features of various embodiments may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the present disclosure, any feature of a drawing may be referenced and / or claimed in conjunction with any feature of any other drawing.

[0070] This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

[0071] Further aspects are provided by the subject matter of the following clauses:

[0072] Item 1. A turbofan engine comprising: a tie rod assembly; a plurality of coupling nuts; a front shaft; a blade disk; a threaded joint coupled to a tapered shaft of the blade disk; a high-pressure compressor rotor; and a high-pressure turbine rotor, the high-pressure turbine rotor including a tapered shaft, wherein a first coupling nut is coupled to the tapered shaft of the high-pressure compressor rotor, a second coupling nut is coupled to the front shaft, and a third coupling nut is coupled to a rear end stage of the high-pressure turbine rotor.

[0073] Clause 2. The turbofan engine of the preceding clause, wherein the threaded engagement portion is configured to engage the internal threads of the front shaft with the external threads of the tie rod.

[0074] Clause 3. The turbofan engine of any preceding clause, further comprising a plurality of compressor stages, wherein at least one compressor stage comprises a first type of disk and a second compressor stage comprises a second type of disk.

[0075] Clause 4. The turbofan engine of any preceding clause, wherein the front shaft comprises an alloy configured to improve low cycle fatigue at thread fillets.

[0076] Clause 5. The turbofan engine of any preceding clause, wherein the threaded engagement is coupled to the tie rod assembly.

[0077] Clause 6. The turbofan engine of any preceding clause, wherein the front shaft is integrally formed with the blisk as a one-piece, unitary component.

[0078] Clause 7. The turbofan engine of any preceding clause, wherein the front shaft is integrally formed with the tie rod assembly as a one-piece, unitary component.

[0079] Clause 8. The turbofan engine of any preceding clause, wherein the tie rod assembly is welded to the blisk as a single-piece component.

[0080] Clause 9. The turbofan engine of any preceding clause, wherein clamping the tie rod assembly into a plurality of loops reduces the effective span of the tie rod assembly.

[0081] Clause 10. The turbofan engine of any preceding clause, wherein the plurality of rings comprises three rings.

[0082] Item 11. A gas turbine engine comprising: a tie rod assembly; a front shaft; a blisk; a threaded joint coupled to a tapered shaft of the blisk; a first coupling nut and a second coupling nut; and a compressor rotor, wherein the first coupling nut is coupled to the tapered shaft of the compressor rotor and the second coupling nut is coupled to the front shaft.

[0083] Clause 12. The gas turbine engine of any preceding clause, further comprising a turbine rotor and a third coupling nut coupled to an aft stage of the turbine rotor.

[0084] Clause 13. The gas turbine engine of any preceding clause, wherein the front shaft comprises an alloy configured to improve low cycle fatigue at thread fillets.

[0085] Clause 14. The gas turbine engine of any preceding clause, further comprising a threaded engagement coupled to the tapered shaft of the blisk.

[0086] Clause 15. The gas turbine engine of any preceding clause, wherein the threaded engagement is coupled to the tie rod assembly.

[0087] Clause 16. The gas turbine engine of any preceding clause, wherein the front shaft is integrally formed with the blisk as a one-piece, unitary component.

[0088] Clause 17. The gas turbine engine of any preceding clause, wherein the front shaft is integrally formed with the tie rod assembly as a one-piece, unitary component.

[0089] Clause 18. The gas turbine engine of any preceding clause, wherein the tie rod assembly is welded with the blisk as a single-piece component.

[0090] Clause 19. The gas turbine engine of any preceding clause, wherein clamping the tie rod assembly into a plurality of rings reduces the effective span of the tie rod assembly.

[0091] Clause 20. The gas turbine engine of any preceding clause, wherein the plurality of rings comprises three rings.

Claims

1. A turbofan engine, characterized in that: include: tie rod assembly; a plurality of coupling nuts; front axle; leaf disk; a threaded joint coupled to the tapered shaft of the blisk; High-pressure compressor rotor; as well as A high-pressure turbine rotor comprising a tapered shaft, wherein a first coupling nut is coupled to the tapered shaft of the high-pressure compressor rotor, a second coupling nut is coupled to the front shaft, and a third coupling nut is coupled to an aft stage of the high-pressure turbine rotor.

2. The turbofan engine according to claim 1, characterized in that: in, The threaded engagement portion is configured to engage the internal threads of the front axle with the external threads of the tie rod assembly.

3. The turbofan engine according to claim 1, wherein: Further included is a plurality of compressor stages, wherein at least one compressor stage includes a first type of disk and a second compressor stage includes a second type of disk.

4. The turbofan engine according to claim 1, wherein: in, The front axle includes an alloy configured to improve low cycle fatigue at the thread fillets.

5. The turbofan engine according to claim 1, wherein: in, The threaded engagement portion is coupled to the tie rod assembly.

6. The turbofan engine according to claim 1, characterized in that: in, The front axle is integrally formed with the blisk as a one-piece, unitary component.

7. The turbofan engine according to claim 1, characterized in that: in, The front axle is integrally formed with the tie rod assembly as a one-piece, unitary component.

8. The turbofan engine according to claim 1, characterized in that: in, The tie rod assembly is welded to the blisk as a single-piece component.

9. The turbofan engine according to claim 1, characterized in that: in, Clamping the tie rod assembly into a plurality of loops reduces the effective span of the tie rod assembly.

10. The turbofan engine according to claim 9, characterized in that: in, The plurality of rings includes three rings.

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