compact compressor

Abstract

Methods, apparatus, systems, and articles of manufacture for compact compressors are disclosed, including a gas turbine engine defining axial and radial directions. The gas turbine engine includes an axial compressor and a radial compressor, wherein the axial compressor is located axially forward of the radial compressor, and a blade assembly includes a separator shroud to separate incoming air into an axial airflow for the axial compressor and a radial airflow for the radial compressor. The blade assembly rotates relative to the axial compressor and in the opposite direction to the radial compressor, and wherein the blade assembly is located axially rearward of the radial compressor.

Description

Technical Field

[0001] This disclosure generally relates to gas turbines, and more specifically to compact compressors. Background Technology

[0002] Gas turbines (also known as gas turbines and turbine engines) are used in a variety of applications, such as industrial power generation and aircraft propulsion systems. These gas turbines consist of one or more compressor stages, a combustor, and one or more turbine stages. Aircraft include gas turbines (e.g., gas turbines, turbine engines, etc.) in their propulsion systems to generate thrust. Summary of the Invention

[0003] This article discloses methods, systems, and products corresponding to compact compressors.

[0004] Some examples provide a gas turbine engine defining axial and radial directions, the gas turbine engine including an axial-flow compressor. The example gas turbine engine also includes a radial-flow compressor, wherein the axial-flow compressor is located axially forward of the radial-flow compressor. The example gas turbine engine also includes a blade assembly including a separator shroud to separate incoming air into an axial airflow for the axial-flow compressor and a radial airflow for the radial-flow compressor, the blade assembly rotating relative to the axial-flow compressor and rotating in the opposite direction relative to the radial-flow compressor, and wherein the blade assembly is located axially rearward of the radial-flow compressor.

[0005] Some examples provide a turbofan engine defining an axial direction, the turbofan engine including an axial compressor stage, the axial compressor stage including a casing and multiple airfoils. The example turbofan engine also includes a centrifugal compressor stage, the centrifugal compressor stage including an impeller and multiple diffuser channels, the centrifugal compressor stage being axially rearward of the axial compressor stage. The example turbofan engine also includes a fan disposed axially rearward of the centrifugal compressor stage.

[0006] Some examples provide a compressor assembly for an aircraft engine, the aircraft engine defining axial and radial directions, the compressor assembly including an axial compressor stage. The example compressor assembly also includes a radial compressor stage axially rearward of the axial compressor stage. The example compressor assembly also includes a blade assembly including an inner airfoil and an outer airfoil separated by a separator shroud to divide incoming air into an axial airflow for the axial compressor and a radial airflow for the radial compressor, and is positioned axially rearward of the radial compressor stage, wherein the axial compressor stage, the radial compressor stage, and the inner airfoil define a flow passage. Attached Figure Description

[0007] Figure 1It is an exemplary conventional turbofan engine used to propel aircraft.

[0008] Figure 2 The image shows a front view of a turbofan engine that can be implemented in conjunction with the teachings of this disclosure, including an AA cut line, the turbofan engine including an example compact compressor and an example blade assembly.

[0009] Figure 3 It includes Figure 2 Example of a compact compressor turbofan engine along Figure 2 The front right top 3D view of the AA cutting line.

[0010] Figure 4 yes Figure 2 A partial perspective view of an example compact compressor, showing an example axial compressor.

[0011] Figure 5 yes Figure 2 Another partial perspective view of an example compact compressor, showing an example axial compressor.

[0012] Figure 6 yes Figure 2 A partial perspective view of an example compact compressor, showing an example radial flow compressor.

[0013] Figure 7 yes Figure 2 Another partial perspective view of an example compact compressor, showing an example radial compressor.

[0014] Figure 8 It is used for Figure 2 An example rotor of an example blade assembly, which can be combined with... Figure 2 The turbofan engine was implemented.

[0015] Figure 9 It is installed in a way that allows for combination Figure 2 The turbofan engine implemented Figure 8 Example rotor Figure 2 A partial 3D view of an example blade assembly.

[0016] Figure 10 It is for Figure 2 The turbofan engine provides an example gear assembly that rotates in the opposite direction.

[0017] The accompanying drawings are not drawn to scale. Instead, the thickness of layers or regions may be enlarged in the drawings. Although these figures show layers and regions with sharp lines and boundaries, some or all of these lines and / or boundaries may be idealized. In reality, boundaries and / or lines may be unobservable, mixed, and / or irregular. Generally, the same reference numerals will be used throughout the drawings and the accompanying written description to refer to the same or similar parts. As used herein, unless otherwise stated, the term "above" describes the relationship of two parts relative to the ground. If the second part has at least one portion between the ground and the first part, then the first part is above the second part. Similarly, as used herein, the first part is "below" the second part when the first part is closer to the ground than the second part. As stated above, the first part may be above or below the second part for one or more of the following: when there are other parts between them, when there are no other parts between them, when the first part and the second part are in contact, or when the first part and the second part are not in direct contact with each other. As used herein, a statement that any part (e.g., layer, film, region, area, or plate) is located on another part in any manner (e.g., positioned, situated, set, or formed, etc.) indicates that the referenced part is in contact with or above the other part, with one or more intermediate parts located between them. As used herein, unless otherwise stated, a connection reference (e.g., attachment, coupling, connection, and joining) may include intermediate members between elements referenced by the connection reference and / or relative movement between those elements. Therefore, a connection reference does not necessarily imply that two elements are directly connected and / or have a fixed relationship with each other. As used herein, a statement that any part is in contact with another part is defined to mean that there is no intermediate part between the two parts.

[0018] Unless otherwise specifically stated, descriptors such as “first,” “second,” “third,” etc., used herein are not intended to in any way assign or otherwise indicate any meaning of priority, physical order, arrangement in a list, and / or sorting, but are merely used as labels and / or arbitrary names to distinguish elements for ease of understanding of the disclosed examples. In some examples, the descriptor “first” may be used to refer to an element in the detailed description, while different descriptors (e.g., “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 may, for example, otherwise share the same name. As used herein, “about” and “approximately” refer to dimensions that may be inaccurate due to manufacturing tolerances and / or other real-world defects. Detailed Implementation

[0019] "Comprising" and "including" (and all their forms and tenses) are used herein as open-ended terms. Therefore, whenever a claim uses "comprising" or "including" (e.g., including, comprising, having, etc.) in any form as a preamble or in any type of claim statement, it should be understood that additional elements, terms, etc., may be present without exceeding the scope of the corresponding claim or statement. As used herein, when the phrase "at least" is used as a transitional term, for example, in the preamble of a claim, it is open-ended in the same way that the terms "comprising" and "including" are open-ended. The term "and / or," when used, for example, in the form of A, B, and / or C, refers to any combination or subset of A, B, and C, such as (1) A alone, (2) B alone, (3) C alone, (4) A and B, (5) A and C, (6) B and C, and (7) A and B and C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A and B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A or B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. As used herein in the context of describing the conduct or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A and B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing the conduct or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A or B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, and (3) at least one A and at least one B.

[0020] As used herein, singular references (e.g., “a,” “an,” “first,” “second,” etc.) do not exclude plurals. As used herein, the term “a” or “an” refers to one or more of that entity. The terms “a” (or “an”), “one or more,” and “at least one” are used interchangeably herein. Furthermore, although listed separately, multiple means, elements, or method actions can be implemented, for example, by a single unit or processor. Moreover, while individual features may be included in different examples or claims, these may be combined, and inclusion in different examples or claims does not imply that the combination of features is infeasible and / or disadvantageous.

[0021] Aircraft include engines used as propulsion systems to generate mechanical power and forces such as thrust. A gas turbine (also called a gas engine or turbine engine) is an internal combustion engine that can be implemented in the propulsion system of an aircraft. For example, a gas turbine can be implemented in combination with a turbofan engine or a turbojet engine. Gas turbines also have important applications in fields such as industrial power generation.

[0022] As used herein, the terms “axial” and “longitudinal” refer to directions parallel to the centerline axis of the gas turbine (e.g., turbofan engine, core gas turbine engine, etc.), while “radial” refers to a direction perpendicular to the axial direction, and “tangential” or “circumferential” refers to directions perpendicular to both the axial and radial directions. Therefore, as used herein, “radially inward” refers to a radial direction from the outer periphery of the gas turbine toward the centerline axis of the gas turbine, while “radially outward” refers to a radial direction from the centerline axis of the gas turbine toward the outer periphery of the gas turbine. As used herein, the terms “forward,” “front,” and “ahead” refer to a relatively upstream position in the airflow passing through or around the component, while the terms “rear” and “rear” refer to a relatively downstream position in the airflow passing through or around the component.

[0023] The basic operation of a gas turbine implemented by a turbofan engine in conjunction with an aircraft's propulsion system involves drawing in a stream of fresh atmospheric air through the front of the turbofan engine. In turbofan engine operation, a first portion of the intake air bypasses the core gas turbine engine of the turbofan engine to directly generate thrust. A second portion of the intake air travels through a conventional booster compressor (e.g., a first compressor) located between a fan in the core gas turbine engine (e.g., the gas turbine) and a high-pressure compressor (e.g., a second compressor). The booster compressor is used to increase or boost the pressure of the second portion of the intake air before it enters the high-pressure compressor. The air stream can then travel through the high-pressure compressor, which further pressurizes the air stream. Both the booster compressor and the high-pressure compressor include a set of blades attached to a rotor and / or shaft. The blades rotate at high speeds (e.g., absolute speed, relative to stationary blades, etc.), and each blade subsequently compresses the air stream. The high-pressure compressor then supplies the pressurized air stream to a combustion chamber (e.g., a combustor). In some examples, the high-pressure compressor supplies the pressurized air stream at speeds of hundreds of miles per hour. In some cases, the combustion chamber includes one or more fuel injector rings that inject a steady stream of fuel into the combustion chamber, where the fuel mixes with the pressurized air stream. A second use of the compressor (especially a high-pressure compressor) is to exhaust air for use in other aircraft systems (e.g., cabin pressure, heating, and air conditioning).

[0024] In the combustion chamber of a core gas turbine engine, fuel is ignited by an electric spark provided by an igniter, where in some examples the fuel burns at temperatures exceeding 2,000 degrees Fahrenheit. The resulting combustion produces a high-temperature, high-pressure gas stream (e.g., hot combustion gas) that passes through another set of blades called turbines. For example, turbines may include a low-pressure turbine and a high-pressure turbine. Each of the low-pressure and high-pressure turbines comprises an intricate array of alternating rotating blades and stationary airfoil-section blades (e.g., wheel blades). The high-pressure turbine is located axially downstream of the combustor and axially upstream of the low-pressure turbine. As the hot combustion gas passes through the turbine, it expands through the blades and / or wheel blades, causing the rotating blades to engage with the rotors of both the high-pressure and low-pressure turbines to rotate.

[0025] The rotating blades of the high-pressure and low-pressure turbines serve at least two purposes. The first purpose of the rotating blades is to drive a fan, high-pressure compressor, and / or booster compressor to draw more pressurized air into the combustion chamber. For example, in a twin-shaft design of a turbofan engine, the low-pressure turbine (e.g., the first turbine) can be attached to and power-connected to the booster compressor (e.g., the first compressor) and the fan via a first shaft, collectively referred to as the first shaft of the gas turbine, such that the rotation of the low-pressure turbine rotor drives the booster compressor rotor and the fan. Similarly, the high-pressure turbine (e.g., the second turbine) can be attached to and power-connected to the high-pressure compressor (e.g., the second compressor) via a second shaft coaxial with the first shaft, collectively referred to as the second shaft of the gas turbine, such that the rotation of the high-pressure turbine rotor drives the high-pressure compressor rotor. The second purpose of the rotating blades is to rotate a generator operatively coupled to the turbine section to generate electricity. For example, the turbine can generate electricity for use in aircraft, power plants, etc.

[0026] Considering the static, dynamic, centrifugal, and / or thermal stress limitations and weight of core gas turbine engines and / or turbofan engines, the design goal of aircraft engines (e.g., turbofan engines) is typically to compress as much air as possible within the compressor of a core gas turbine engine. A metric defining the compression effect of a compressor is its compression ratio (e.g., pressure ratio). The compression ratio of a turbofan engine compressor is the ratio of the pressure at the compressor outlet (e.g., the high-pressure compressor outlet at the combustion chamber of a gas turbine) to the pressure at the fan inlet. A higher compression ratio increases the thermal efficiency of the turbine engine and reduces its specific fuel consumption (e.g., the ratio of fuel used to thrust generated by the jet engine). Therefore, increasing the compression ratio of a gas turbine compressor can increase the thrust generated by a jet engine (e.g., a turbofan engine, etc.) and / or increase the fuel efficiency of the jet engine. In turn, one objective of gas turbine design is to increase the compression ratio. Although the examples disclosed herein are discussed in conjunction with turbofan jet engines, it should be understood that the examples disclosed herein can be implemented in conjunction with turbojet engines, turboprop jet engines, gas turbines and / or gas turbines for power generation, or any other suitable application where increasing the compression ratio across one or more compressors is desired.

[0027] In some examples disclosed herein, the booster compressor and / or low-pressure compressor includes an axial compressor, which includes one or more stages, each stage including a rotating airfoil ring. The airfoils of the stage increase the kinetic energy of the inlet air and compress the air. In some of these examples, the axial compressor includes only rotating airfoils (e.g., blades) and does not include a stator portion. In other examples, air diffuses over the stator to the rear of the airfoils for further compression. In some examples disclosed herein, the booster compressor also includes a centrifugal compressor. A centrifugal compressor, unlike an axial compressor, includes a rotating centrifugal impeller comprising annularly arranged blades. The impeller includes a radially larger rotor and / or a central portion that moves rearward (e.g., away from the inlet) of the turbine engine. The centrifugal compressor draws in inlet air incident on the impeller, which increases the kinetic energy of the air and compresses it as it moves radially outward through the impeller. The impeller is surrounded by one or more stationary diffusers for further compression. In some examples, the diffuser includes an annular array of stationary blades (e.g., a stator) and / or one or more discrete channels. A mixed-flow compressor combines aspects of an axial compressor and a centrifugal compressor, thereby discharging compressed air at an angle (e.g., diagonally) between the radial and axial directions. Additionally or alternatively, the example turbofan engines disclosed herein include one or more mixed-flow compressors. More generally, as used herein, a radial compressor includes both (a) a centrifugal compressor and (b) a mixed-flow compressor, and directs compressed air in at least a partial radial direction.

[0028] The example compact compressors disclosed herein (e.g., example compact compressors for turbofan engines) include axial-flow and radial-flow compressors, collectively referred to as compact compressors, located at the position of a conventional nose cone (e.g., a rotating cone) of a turbofan engine. The example compact compressors do not include a stator assembly. Compared to a conventional turbofan engine, the compact compressors are located at the front of the turbofan engine, in front of the fan, and are implemented as part of an example blade assembly. The compact compressors include an axial-flow compressor in series with the radial-flow compressor. In some examples, the radial-flow compressor may be a mixed-flow compressor that discharges compressed air radially outward and axially rearward of the turbofan engine. In some examples, the radial-flow compressor may be a centrifugal compressor and can therefore discharge a large amount of air in the radial direction. In some examples, only the radial-flow compressor is located at the position of a conventional nose cone (e.g., a rotating cone) of the turbofan engine. In such examples, the radial-flow compressor, implemented as part of the example blade assembly in front of the fan, discharges compressed air through the turbofan engine without the assistance of an axial-flow compressor.

[0029] Advantageously, the examples disclosed herein can increase the compression ratio of a turbofan engine and reduce the length of a turbofan engine by positioning a booster compressor (e.g., a compact compressor) at the nose cone position, thereby reducing the material used to construct the turbofan engine. In the example blade assembly, an outer airfoil is separated from an inner airfoil by an annular separator shroud (e.g., a splitter) behind the compact compressor. The outer airfoil (e.g., a fan) leads to a bypass duct behind the blade assembly. The inner airfoil receives compressed air from the compact compressor and leads to the example high-pressure compressor. The blade assembly can be mounted via axial slots (e.g., dovetail slots, fir slots, etc.) on the rotor. The rotor is coupled to a spool (e.g., a shaft) and a turbine (e.g., a low-pressure turbine) via a gearbox, such that the rotor and blade assembly rotate in opposite directions relative to the compact compressor. In other examples, the blade assembly and rotor can be implemented as a bladed disk (blade disk).

[0030] In some examples, the blades of the axial compressor implemented with the compact compressor disclosed herein include a wide chord length relative to the chord length of the blades of a conventional booster compressor. In some examples, the implementation of the blade assembly and the compact compressor reduces the length and weight of the turbocharger. In some examples, the compact compressor and / or blade assembly have a higher pressure ratio and require fewer stages than a conventional booster compressor due to the high tip speeds of the blades of the axial compressor and the impeller of the compact compressor, as well as due to the counter-rotation between the compact compressor and the blade assembly. In some examples, the geometry of the compact compressor eliminates the need for inlet guide vanes (IGV) and variable stator vanes (VSV), thereby reducing the complexity of the turbocharger including the compact compressor. In some examples, the absence of the stator portion of the compact compressor reduces icing on components associated with the compact compressor compared to components associated with a conventional booster compressor. In some examples, arranging the compact compressor at the front of the turbocharger simplifies the assembly of the turbocharger including the compact compressor and allows for maintenance of the compact compressor without disassembling the turbocharger. In some examples, engines that include compact compressors burn less fuel than turbocharged engines that include conventional supercharged compressors.

[0031] Figure 1 This is a schematic diagram of an example conventional turbofan gas turbine engine 102. The example conventional turbofan engine 102 includes an example core gas turbine engine 106, an example fan section 108, an example housing 110, an example annular inlet 112, an example conventional booster compressor 114, an example high-pressure compressor 116 (e.g., a high-pressure multistage axial compressor), an example combustor 118, a first example turbine 120, a first example drive shaft 122, a second example turbine 124, a second example drive shaft 126, an example exhaust nozzle 128, an example axial fan rotor assembly 130, an example annular fan housing 132, an example guide vane 134, an example fan rotor blade 136, an example downstream section 138, an example airflow duct 140, an example speed reduction device 142, an example inlet 150, and example combustion products 158.

[0032] For illustrative purposes, a conventional turbofan engine 102 is shown having a longitudinal or axial centerline axis 104 extending through the conventional turbofan engine 102. The direction of the flow is determined by... Figure 1 Arrow 148 is shown in the diagram. These directional terms are for convenience only and do not require a specific orientation of the structure being described. Figure 1It also includes annotated direction diagrams for the axial direction A, radial direction R, and circumferential direction C. Generally, as used herein, the axial direction A is a direction extending approximately parallel to the centerline axis 104, the radial direction R is a direction extending perpendicularly outward from the centerline axis 104, and the circumferential direction C is a direction extending concentrically around the centerline axis 104.

[0033] Figure 1 The conventional turbofan engine 102 includes a core gas turbine engine 106 and a fan section 108 positioned upstream therefrom. The core gas turbine engine 106 typically includes a generally tubular housing 110 defining an annular inlet 112. Furthermore, the housing 110 may further enclose and support a conventional booster compressor 114, which is used to increase the pressure of air entering the core gas turbine engine 106 to a first pressure level. A high-pressure compressor 116 may then receive pressurized air from the conventional booster compressor 114 and further increase the pressure of this air to a second pressure level.

[0034] exist Figure 1 In the example shown, the pressurized air exiting the high-pressure compressor 116 can then flow to the combustor 118, where fuel is injected into the pressurized air stream, and the resulting mixture is combusted within the combustor 118. The high-energy combustion products are guided from the combustor 118 to a first (high-pressure) turbine 120 via a first (high-pressure) drive shaft 122 to drive the high-pressure compressor 116, and then to a second (low-pressure) turbine 124 via a second (low-pressure) drive shaft 126, which is substantially coaxial with the first drive shaft 122, to drive the conventional supercharger 114 and the fan section 108. After each of the drive turbines 120 and 124, the combustion products can be discharged from the core gas turbine engine 106 via exhaust nozzles 128 to provide propulsive thrust.

[0035] In some examples, each of the conventional booster compressor 114 and the high-pressure compressor 116 may include multiple compressor stages, each stage comprising an annular array of stationary compressor blades and an annular array of rotating compressor blades positioned immediately downstream of the compressor blades. Similarly, each of the turbines 120, 124 may include multiple turbine stages, each stage comprising an annular array of stationary nozzle blades and an annular array of rotating turbine blades positioned immediately downstream of the nozzle blades.

[0036] In addition, such as Figure 1As shown, the fan section 108 of a conventional turbofan engine 102 typically includes a rotatable axial fan rotor assembly 130 configured to be surrounded by an annular fan housing 132. The fan housing 132 may be configured to be supported relative to the core gas turbine engine 106 by a plurality of substantially radially extending, circumferentially spaced outlet guide vanes 134. Thus, the fan housing 132 may surround the fan rotor assembly 130 and its corresponding fan rotor blades 136. Furthermore, a downstream section 138 of the fan housing 132 may extend over the outer portion of the core gas turbine engine 106 to define a secondary or bypass airflow duct 140 providing additional propulsive jet thrust.

[0037] In some examples, the second (low-pressure) drive shaft 126 is directly coupled to the fan rotor assembly 130 to provide a direct drive configuration. Alternatively, the second drive shaft 126 may be coupled to the fan rotor assembly 130 via a reduction gear 142 (e.g., a reduction gear or gearbox) to provide an indirect drive or geared configuration. Such a reduction gear may also be provided between any other suitable shafts and / or spools within the conventional turbofan engine 102 as needed or required.

[0038] During operation of the conventional turbofan engine 102, an initial airflow (indicated by arrow 148) enters the engine 102 through the relevant inlet 150 of the fan housing 132. The airflow 148 then passes through the fan blades 136 and splits into a first compressed airflow (indicated by arrow 152) moving through a duct 140 and a second compressed airflow (indicated by arrow 154) entering the conventional supercharger compressor 114. The pressure of the second compressed airflow 154 then increases and enters the high-pressure compressor 116 (as indicated by arrow 156). After mixing with fuel and burning within the combustor 118, the combustion products 158 exit the combustor 118 and flow through the first turbine 120. Thereafter, the combustion products 158 flow through the second turbine 124 and exit the exhaust nozzle 128, providing thrust to the conventional turbofan engine 102.

[0039] exist Figure 1 In the example shown, the conventional turbofan engine 102 includes a rotating cone 160 (e.g., a nose cone, etc.) at the front of the axial fan rotor assembly 130. The rotating cone 160 is used to direct air incident on the axial fan rotor assembly 130 to the first and / or second compressed air flow (e.g., an intake) of the conventional turbofan engine 102, indicated by arrows 152, 154, and to increase the aerodynamics of the conventional turbofan engine 102. However, the rotating cone 160 includes a large portion of the conventional turbofan engine 102 (e.g., a large portion of the axial and radial length of the conventional turbofan 102, the large volume portion of the conventional turbofan 102, etc.).

[0040] Figure 2 This is a front view of an example turbofan engine 200 (e.g., a turbofan engine, a gas turbine engine, etc.) that can be implemented in conjunction with the teachings of this disclosure. Figure 1 Compared to a conventional turbofan engine 102, the turbofan engine 200 includes an example compact compressor 202 and an example blade assembly 204. Figure 2 In the example shown, the blade assembly 204 includes a separator shroud and is positioned axially rearward of the compact compressor 202. During operation of the turbofan engine 200, the first portion of the inlet airflow (e.g., similar to...) Figure 1 The first part of arrow 152 enters the turbofan engine 200 at the front of the blade assembly 204. Similarly, the second part of the inlet airflow (e.g., similar to...) Figure 1 The second part of arrow 154, for the working fluid of the core gas turbine engine including the compact compressor 202, enters the turbofan engine 200 at the front of the compact compressor 202. Figure 2 In the example shown, the turbofan engine 200 includes an example fan housing 206 surrounding a compact compressor 202 and a blade assembly 204. Figure 2 The diagram shows the AA cutting line 208. For example, aspects of the turbofan engine 200 (e.g., the compact compressor 202 and blade assembly 204) can be manufactured using subtractive manufacturing techniques (such as computer numerical control (CNC) milling, electrochemical machining (ECM), etc.) and can be formed from materials such as titanium alloys (e.g., titanium-aluminum alloys and / or titanium-chromium alloys), steel alloys (e.g., steel-chromium alloys), nickel alloys (e.g., nickel-copper alloys, nickel-iron alloys, nickel-chromium alloys, nickel-niobium alloys and / or nickel-carbon alloys), and / or ceramic matrix composites (CMC). Additionally or alternatively, aspects of the turbofan engine 200 can be manufactured using additive manufacturing techniques and / or formed from other materials.

[0041] Figure 3 yes Figure 2 Example turbofan engine 200 including compact compressor 202 along Figure 2 The front right top 3D view of the AA cutting line 208 cut. Figure 3 In the example shown, the compact compressor 202 ( Figure 2 This includes an example axial compressor 302 (e.g., an example axial compressor stage, an example bladeless compressor, etc.) and an example radial compressor 304 (e.g., a centrifugal compressor stage, a mixed-flow compressor stage, etc.) axially rear of the axial compressor 302. For example, the axial compressor 302 is axially forward of the radial compressor 304 relative to the turbofan engine 200. Blade assembly 204 ( Figure 2This includes an example inner airfoil 306 separated from the example outer airfoil 308 by an exemplary separator shroud 310. In some examples, the outer airfoil 308 of the blade assembly 204 defines a bypass flow passage disposed in the gas turbine engine (e.g., turbofan engine 200) (e.g., defined in the separator shroud 310 and fan housing 206). Figure 2 The fan is located in the bypass flow channel between the axial compressor 302 and the radial compressor 304, and the inner airfoil 306 of the blade assembly 204 is disposed in the main flow channel of the gas turbine engine shared by the inner airfoil 306, the axial compressor 302, and the radial compressor 304. Although in Figure 3 Only a portion of the inner airfoil 306 and outer airfoil 308 are visible and / or marked in the view, but the blade assembly 204 includes other airfoils 306, 308 arranged in a ring around and mounted on the example first rotor 312.

[0042] The turbofan engine 200 also includes an example low-pressure spool 314 (e.g., similar to...). Figure 1 The second (low-pressure) drive shaft 126). Figure 3 In the example shown, the low-pressure spool 314 is force-transmittingly connected to the axial compressor 302 and the radial compressor 304. The low-pressure spool 314 is also force-transmittingly connected to the first rotor 312 and the blade assembly 204 via a gearbox and / or gear assembly. Figure 10 The gear assembly, discussed in more detail, provides counter-rotation between (a) the combination of the axial compressor 302 and the first rotor 312 and (b) the radial compressor 304.

[0043] exist Figure 3 In the example shown, the first part of the inlet airflow (e.g., similar to...) Figure 1 The first portion of arrow 152 enters the turbofan engine 200 at the front of the blade assembly 204. More specifically, the first portion of the inlet airflow drives the outer airfoil 308 of the blade assembly 204. The first portion of the airflow then bypasses the rest of the turbofan engine 200 through a bypass duct and exits at the rear of the turbofan engine 200. Figure 3 (Not shown in the view). The second part of the inlet airflow (e.g., similar to...) Figure 1 The second part of arrow 154 enters at the front of the axial compressor 302 of the turbofan engine 200. The axial compressor 302 compresses the second part of the airflow, which then moves through the radial compressor 304. The radial compressor 304 further compresses the second part of the airflow, which then moves through the inner airfoil 306. The second part of the airflow then moves to the high-pressure compressor (not shown), for example... Figure 1 High-pressure compressor 116.

[0044] exist Figure 3 In the example shown, the separator shroud 310 prevents the first and second portions of the airflow from mixing. Therefore, the outer airfoil 308 functions as the fan of the turbofan engine 200. Because the inner airfoil 306 and the outer airfoil 308 are in a force-transmitting connection, when the first rotor 312 is driven (e.g., by a low-pressure turbine (e.g., turbine 124)), the inner airfoil 306 compresses the second portion of the airflow, while the outer airfoil 308 compresses the first portion of the airflow.

[0045] Figure 4 It is along Figure 2 AA cutting line 208 cut Figure 2 A partial perspective view of an example compact compressor 202, showing... Figure 3 Example axial compressor 302 and example radial compressor 304. In Figure 4 In the example shown, blade assembly 204 and fan housing 206 are removed. The axial compressor 302 includes example blades 402 mounted on an example second rotor 404. Although in Figure 4 In the view, only a portion of the blades 402 are visible and / or marked, but the axial compressor 302 includes other blades 402 arranged annularly around and mounted on the second rotor 404. Figure 4 In the example shown, the second rotor 404 is connected to the low-voltage spool 314 ( Figure 2 The blades 402 are integrally formed around and integrally formed with the example annular disk 406, forming an impeller disk (e.g., a bladed disk). The impeller disk can be manufactured using subtractive manufacturing techniques (e.g., computer numerical control (CNC) milling, electrochemical machining (ECM), etc.) and can be mounted on the second rotor 404. The axial compressor 302 also includes a front-bent housing 408 surrounding the impeller disk. The front-bent housing 408 is used to radially outwardly guide the inlet air to the outer airfoil 308 (as shown in the diagram). Figure 3 (As shown in the diagram) or radially inwardly directed towards the axial compressor 302. After the air is compressed by the blades 402, it moves to the radial compressor 304, combining... Figure 6 To be discussed in more detail. Alternatively, the axial compressor 302 may be provided with an annular blade array extending from the housing 408 toward the rear of the axial compressor 302 to further compress the air.

[0046] In some examples, blades 402 have a larger chord length compared to the blades of the supercharger 114 of a conventional turbofan engine 102. As used herein, "chord length" is the distance from the leading edge to the trailing edge of the blade. For example, blades 402 of an axial compressor 302 can have a larger chord length, such that the blade's chord-to-height ratio is between 2 and 6, while the standard chord-to-height ratio is approximately 1. The increased chord length can improve fuel efficiency.

[0047] Figure 5 It is along Figure 2 AA cutting line 208 cut Figure 2 Another partial perspective view of the example compact compressor 202 shows Figure 3 Example axial compressor 302. In Figure 5 In the middle, the forward-bent housing 408 is visible.

[0048] Figure 6 It is along Figure 2 AA cutting line 208 cutting includes Figure 3 Example of a radial flow compressor 304 Figure 2 A partial perspective view of an example compact compressor 202. Figure 6 In the example shown, for visual clarity, blade 402 and annular disk 406 have been removed (in... Figure 4 The bladed disk is shown in the image. Figure 6 In the example, the radial flow compressor 304 includes an impeller 602. Figure 6 In the example, impeller 602 includes example blade 604 positioned between example central portion 606 and example curved portion 608 of housing 408. Although in Figure 6 Only a portion of the impeller 604 is visible and / or marked in the view, but the radial compressor 304 includes other impellers 604 annularly surrounding the central portion 606. Moving towards the rear of the compact compressor 202, the radius of the central portion 606 increases, and the distance between the central portion 606 and the curved portion 608 of the housing 408 also increases. Therefore, a second portion of the air is compressed and moved radially outward through the impeller 602. Figure 6 In the middle, impeller 602 and Figure 4 The second rotor 404 is connected and / or integrated.

[0049] Figure 7 yes Figure 2 Example compact compressor 202 along Figure 2 Another partial perspective view of the AA-cut line 208 cut shows the example radial compressor 304. (Relative to...) Figure 6 ,exist Figure 7In the middle, the curved portion 608 of the housing 408 is removed for visual clarity. The impeller 602 is surrounded by one or more stationary diffusers 702. Figure 7 In the example shown, diffuser 702 is a discrete-channel diffuser comprising a ring array of discrete channels 704 extending through example manifold 706. Figure 7 In this configuration, diffuser 702 is fluidly connected to example cavity 708 included in manifold 706. The reduced velocity of the air as it flows through discrete channel 704 further compresses it. A second portion of the air then flows to the inner airfoil 306 of counter-rotating blade assembly 204 (both within...). Figure 3 (as shown in the image), and flows to the high-pressure compressor (not shown).

[0050] Figure 8 It can be combined Figure 2 The turbofan engine 200 implemented Figure 3 Example of the first rotor 312. In Figure 8 In the example shown, the first rotor 312 is provided with an axial groove 802 (e.g., a dovetail groove, a fir groove, etc.) to mount the blade assembly 204. Figure 2 (As shown in the image). Figure 8 In this design, the first rotor 312 also includes a first protrusion 804 and a second protrusion 806 to retain the reciprocal protrusions of the blade assembly 204. One or more pins can connect the blade assembly 204 to the first rotor 312. The first rotor 312 is connected via a gearbox (not shown) to a low-pressure spool 314 (not shown) and a turbine (e.g., a low-pressure turbine), such that the first rotor 312 and the blade assembly 204 rotate in opposite directions relative to the compact compressor 202. The first rotor 312 therefore includes a central opening 808 through which the low-pressure spool 314 extends. Figure 8 In the middle, the first rotor 312 is coaxial with the low-voltage shaft 314.

[0051] Figure 9 It is installed in a way that allows for combination Figure 2 The turbofan engine 200 implemented Figure 3 On the first rotor 312 Figure 2 A partial perspective view of an example blade assembly 204, where a portion of the blade assembly 204 has been removed. Figure 9 In the example shown, a discrete cavity 902 is defined between every two adjacent annular inner airfoils 306, through which a second portion of the air is compressed and moves through the blade assembly 904.

[0052] Figure 10 It is along Figure 2 AA cutting line 208 cut Figure 2A turbofan engine 200 includes an example gear assembly 1002 (e.g., a gearbox) to provide counter-rotation between components. The example gear assembly 1002 is used in axial compressor 302 and radial compressor 304. Figure 3 The example gear assembly 1002 also provides counter-rotation between the blade assembly 204 and the first rotor 312. Figure 10 In the example shown, gear assembly 1002 is configured in a star shape, including gears of various sizes (e.g., smaller radius gears, larger radius gears, etc.). In the example star shape, the bracket of gear assembly 1002 is fixed or attached to a frame rather than rotating about the engine axis. The smaller radius gear of example gear assembly 1002 is coupled below the low-pressure line spool 314 at a first position 1004, while the larger radius gear is coupled to the first rotor 312 at a second position 1006. For example, gear assembly 1002 can drive (a) axial compressor 302 and vane assembly 204 in a counterclockwise and / or positive direction, and can drive (b) first rotor 312 and radial compressor 304 in a clockwise and / or negative direction. Axial compressor 302 and vane assembly 204 are both on a first shaft (e.g., low-pressure line spool 314) and rotate in the same first direction. Radial compressor 304 is on a second shaft (e.g., first rotor 312) and rotates in a second direction. Clockwise and counterclockwise rotation can occur at different speeds via gears of different sizes in gear assembly 1002.

[0053] For example, air enters the engine in axial compressor 302, travels to radial compressor 304, moves along hub flow path 1008, and flows to high-pressure compressor 120. An example gearbox (e.g., a star gearbox) 1002 achieves low-speed rotation in one direction (e.g., clockwise) and high-speed rotation in the opposite direction (e.g., counterclockwise), where the speed (e.g., the ratio of low speed to high speed) is based on the gear ratio between the first rotor 312 and the shaft 314 (e.g., the low-pressure line shaft 314) (e.g., the gear radius in gearbox 1002 defines the low-speed to high-speed ratio between the first rotor 312 and the shaft 314, etc.). Therefore, for example, gearbox 1002 can be used to drive axial compressor 302 and the associated shaft 314 (e.g., the low-pressure line shaft 314) at a first speed in a first direction, while simultaneously driving radial compressor 304 at a second speed higher than the first speed in a second opposite direction. Therefore, the axial compressor 302 and the radial compressor 304 can be arranged together and positioned to replace the conventional nose cone (e.g., rotary cone) of the turbofan engine 200.

[0054] Based on the above, it should be understood that example methods, apparatus, and articles have been disclosed that reduce material usage and improve the efficiency of turbofan engines by placing axial-flow and radial-flow compressors (e.g., collectively referred to as compact compressors) in spaces traditionally dedicated to rotating cones.

[0055] Although certain example methods, apparatuses, and articles have been disclosed herein, the scope of this patent is not limited thereto. Rather, this patent covers all methods, apparatuses, and articles that fall fully within the scope of the claims of this patent.

[0056] Further aspects of the invention are provided by the subject matter of the following clauses:

[0057] 1. A gas turbine engine device defining an axial direction and a radial direction, the gas turbine engine comprising: an axial compressor; a radial compressor, wherein the axial compressor is located axially forward of the radial compressor; and a blade assembly including a separator shroud for separating incoming air into an axial airflow for the axial compressor and a radial airflow for the radial compressor, the blade assembly rotating relative to the axial compressor and rotating in the opposite direction relative to the radial compressor, and wherein the blade assembly is located axially rearward of the radial compressor.

[0058] 2. The device according to any of the preceding clauses, wherein the outer airfoil of the blade assembly defines a fan disposed in a bypass flow passage of the gas turbine engine, and wherein the inner airfoil of the blade assembly is disposed in the main flow passage of the gas turbine engine shared by the inner airfoil, the axial compressor, and the radial compressor.

[0059] 3. The device according to any of the preceding clauses, wherein the radial compressor is a centrifugal compressor comprising an impeller and at least one diffuser passage.

[0060] 4. The device according to any of the preceding clauses, wherein the impeller includes a solid central portion, the radial diameter of which increases toward the rear of the gas turbine engine in the axial direction.

[0061] 5. The device according to any of the preceding clauses, wherein the gas turbine engine is a turbofan engine.

[0062] 6. The apparatus according to any of the preceding clauses, wherein the radial flow compressor is a mixed flow compressor.

[0063] 7. The device according to any of the preceding clauses further includes a gearbox located axially rearward of the blade assembly to provide counter-rotation between (a) the blade assembly and the axial compressor and (b) the radial compressor.

[0064] 8. The device according to any of the preceding clauses further includes a gearbox located axially rearward of the blade assembly to provide counter-rotation between the axial compressor and the radial compressor.

[0065] 9. A turbofan engine device defining an axial direction, comprising: an axial compressor including a housing and a plurality of airfoils; a centrifugal compressor including an impeller and a plurality of diffuser channels, the centrifugal compressor being axially rear of the axial compressor; and a fan disposed axially rear of the centrifugal compressor.

[0066] 10. The device according to any of the preceding clauses, wherein the ratio of the chord length to the height of said plurality of airfoils is between 2 and 6.

[0067] 11. The device according to any of the preceding clauses, wherein at least one of (a) the axial compressor or (b) the centrifugal compressor is manufactured using a subtractive manufacturing process.

[0068] 12. The apparatus according to any of the preceding clauses, wherein the axial compressor and the centrifugal compressor at least partially define a booster compressor, and the apparatus further includes a high-pressure compressor disposed axially rearward of the fan.

[0069] 13. The device according to any of the preceding clauses, wherein the fan partially defines a blade assembly including a separator shroud to separate incoming air into an axial airflow for the axial compressor and a radial airflow for the centrifugal compressor, the fan being disposed in a bypass flow passage.

[0070] 14. The device according to any of the preceding clauses, wherein the blade assembly further includes an inner airfoil disposed in a flow passage shared with the axial compressor stage and the centrifugal compressor stage.

[0071] 15. The apparatus according to any of the preceding clauses, wherein the axial compressor and the centrifugal compressor are connected to the low-pressure turbine via a low-pressure shaft.

[0072] 16. A compressor assembly for an aircraft engine, the aircraft engine defining an axial direction and a radial direction, the compressor assembly comprising: an axial compressor; a radial compressor located axially rear of the axial compressor; and a blade assembly including an inner airfoil and an outer airfoil separated by a separator shroud to divide incoming air into an axial airflow for the axial compressor and a radial airflow for the radial compressor, and disposed axially rear of the radial compressor, wherein the axial compressor, the radial compressor, and the inner airfoil define a flow channel.

[0073] 17. The apparatus according to any of the preceding clauses, wherein the radial compressor is a mixed-flow compressor, further comprising a plurality of diffuser channels.

[0074] 18. The apparatus according to any of the preceding clauses, wherein the radial compressor is a centrifugal compressor, further comprising a plurality of diffuser channels.

[0075] 19. The device according to any of the preceding clauses, wherein the blade assembly is mounted to the central rotor.

[0076] 20. The device according to any of the preceding clauses, wherein the blade assembly rotates in the opposite direction relative to the radial compressor.

[0077] 21. The apparatus according to any of the preceding clauses, wherein at least one of (a) the axial flow compressor or (b) the radial flow compressor includes one or more milling components.

[0078] 22. A gas turbine engine apparatus defining an axial and radial direction, the gas turbine engine comprising: a radial compressor located in a nose cone; and a blade assembly including a separator shroud to separate incoming air into an axial airflow for the axial compressor and a radial airflow for the radial compressor, the blade assembly rotating counter-rotating relative to the radial compressor, and wherein the blade assembly is located axially rearward of the radial compressor.

[0079] 23. The device according to any of the preceding clauses, wherein the outer airfoil of the blade assembly defines a fan disposed in a bypass flow passage of the gas turbine engine, and wherein the inner airfoil of the blade assembly is disposed in the main flow passage of the gas turbine engine shared by the inner airfoil and the radial compressor.

[0080] 24. The apparatus according to any of the preceding clauses, wherein the radial compressor is a centrifugal compressor comprising an impeller and at least one diffuser passage.

[0081] 25. The device according to any of the preceding clauses, wherein the impeller includes a solid central portion, the radial diameter of which increases toward the rear of the gas turbine engine in the axial direction.

[0082] 26. The device according to any of the preceding clauses, wherein the gas turbine engine is a turbofan engine.

[0083] 27. The apparatus according to any of the preceding clauses, wherein the radial flow compressor is a mixed flow compressor.

[0084] 28. The device according to any of the preceding clauses further includes a gearbox located axially rearward of the blade assembly to provide counter-rotation between (a) the blade assembly and (b) the radial compressor.

[0085] 29. A turbofan engine device defining an axial direction, comprising: an axial compressor including a housing and a plurality of airfoils; a centrifugal compressor including an impeller and a plurality of diffuser channels, the centrifugal compressor being axially rear of the axial compressor; and a fan disposed axially rear of the centrifugal compressor.

[0086] 30. The device according to any of the preceding clauses, wherein the ratio of the chord length to the height of the plurality of airfoils is between Examples 2 and 6.

[0087] 31. The device according to any of the preceding clauses, wherein the centrifugal compressor is manufactured using a subtractive manufacturing process.

[0088] 32. The apparatus according to any of the preceding clauses, wherein the centrifugal compressor at least partially defines a booster compressor, and the apparatus further includes a high-pressure compressor disposed axially rearward of the fan.

[0089] 33. The device according to any of the preceding clauses, wherein the fan partially defines a blade assembly including a separator shroud to separate incoming air into an axial airflow for the axial compressor and a radial airflow for the centrifugal compressor, the fan being disposed in a bypass flow passage.

[0090] 34. The device according to any of the preceding clauses, wherein the blade assembly further includes an inner airfoil disposed in a flow passage shared with the centrifugal compressor stage.

[0091] 35. The apparatus according to any of the preceding clauses, wherein the centrifugal compressor is connected to the low-pressure turbine via a low-pressure shaft.

[0092] 36. A compressor assembly for an aircraft engine, the aircraft engine defining an axial direction and a radial direction, the compressor assembly comprising a radial compressor and a blade assembly, the blade assembly comprising an inner airfoil and an outer airfoil separated by a separator shroud to divide incoming air into an axial airflow for the axial compressor and a radial airflow for the radial compressor, and disposed axially rearward of the radial compressor, wherein the radial compressor and the inner airfoil define a flow passage.

[0093] 37. The apparatus according to any of the preceding clauses, wherein the radial compressor is a mixed-flow compressor, further comprising a plurality of diffuser channels.

[0094] 38. The apparatus according to any of the preceding clauses, wherein the radial compressor is a centrifugal compressor, further comprising a plurality of diffuser channels.

[0095] 39. The device according to any of the preceding clauses, wherein the blade assembly is mounted to the central rotor.

[0096] 40. The device according to any of the preceding clauses, wherein the blade assembly rotates in the opposite direction relative to the radial compressor.

[0097] 41. The apparatus according to any of the preceding clauses, wherein the radial flow compressor includes one or more milled components.

[0098] The following claims are incorporated herein by reference, each of which exists independently as a separate embodiment of this disclosure.

Claims

1. A gas turbine engine, the gas turbine engine defining an axial direction and a radial direction, characterized in that, The gas turbine engine includes: Axial flow compressor; A radial compressor, wherein the axial compressor is located axially ahead of the radial compressor, and wherein the axial compressor and the radial compressor at least partially define a booster compressor; A blade assembly including a separator shroud to provide splitting between a first airflow and a second airflow, the blade assembly rotating relative to the axial compressor and rotating counter-rotating relative to the radial compressor, wherein the blade assembly is located axially rearward of the radial compressor; and A high-pressure compressor is located axially rearward of the blade assembly.

2. The gas turbine engine according to claim 1, characterized in that, The outer airfoil of the blade assembly defines a fan disposed in a bypass flow passage of the gas turbine engine, and the inner airfoil of the blade assembly is disposed in the main flow passage of the gas turbine engine shared by the inner airfoil, the axial compressor, and the radial compressor.

3. The gas turbine engine according to claim 1, characterized in that, The radial compressor is a centrifugal compressor that includes an impeller and at least one diffuser channel.

4. The gas turbine engine according to claim 3, characterized in that, The impeller includes a solid central portion, the radial diameter of which increases toward the rear of the gas turbine engine in the axial direction.

5. The gas turbine engine according to claim 1, characterized in that, The gas turbine engine mentioned above is a turbofan engine.

6. The gas turbine engine according to claim 1, characterized in that, The radial flow compressor mentioned above is a mixed flow compressor.

7. The gas turbine engine according to claim 1, characterized in that, It further includes a gearbox located axially rearward of the blade assembly to provide counter-rotation between the axial compressor and the radial compressor.

8. The gas turbine engine according to claim 1, characterized in that, It further includes a gearbox located axially rearward of the blade assembly to provide counter-rotation between (a) the blade assembly and the axial compressor and (b) the radial compressor.

9. A turbofan engine, wherein the turbofan engine defines an axial direction, characterized in that, The turbofan engine includes: An axial compressor, the axial compressor comprising a housing and a plurality of airfoil components; A centrifugal compressor, the centrifugal compressor including an impeller and a plurality of diffuser channels, the centrifugal compressor being disposed axially rearward of the axial compressor, wherein the axial compressor and the centrifugal compressor at least partially define a booster compressor; A fan, the fan being disposed axially rearward of the centrifugal compressor; and A high-pressure compressor is located axially rear of the fan.

10. The turbofan engine according to claim 9, characterized in that, The ratio of the chord length to the height of the plurality of airfoils is between 2 and 6.

11. The turbofan engine according to claim 9, characterized in that, At least one of the axial compressor (a) or the centrifugal compressor (b) is manufactured using a subtractive manufacturing process.

12. The turbofan engine according to claim 9, characterized in that, The fan partially defines a blade assembly including a separator shroud to split (a) a first airflow corresponding to an outer airfoil of a bypass flow passage and (b) a second airflow from the centrifugal compressor, the fan being disposed in the bypass flow passage.

13. The turbofan engine according to claim 12, characterized in that, The blade assembly further includes an inner airfoil disposed in a flow channel shared with the axial compressor stage and the centrifugal compressor stage.

14. The turbofan engine according to claim 9, characterized in that, The axial compressor and the centrifugal compressor are connected to the low-pressure turbine via a low-pressure shaft.

15. A compressor assembly for an aircraft engine, the aircraft engine defining an axial direction and a radial direction, characterized in that, The compressor assembly includes: Axial flow compressor; A radial compressor, the radial compressor being located axially rearward of the axial compressor, wherein the axial compressor and the radial compressor at least partially define a booster compressor; A blade assembly including an inner airfoil and an outer airfoil separated by a separator shroud to provide flow splitting between a first airflow and a second airflow, and disposed axially rearward of the radial compressor, wherein the axial compressor, the radial compressor, and the inner airfoil define a flow passage; and A high-pressure compressor is located axially rearward of the blade assembly.

16. The compressor assembly for an aircraft engine according to claim 15, characterized in that, The radial compressor is a mixed-flow compressor and further includes multiple diffuser channels.

17. The compressor assembly for an aircraft engine according to claim 15, characterized in that, The radial compressor is a centrifugal compressor and further includes multiple diffuser channels.

18. The compressor assembly for an aircraft engine according to claim 15, characterized in that, The blade assembly is mounted to the central rotor.

19. The compressor assembly for an aircraft engine according to claim 15, characterized in that, The blade assembly rotates in the opposite direction to the radial compressor.

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