Front structure for a compressor device of a gas turbine engine, a method for producing a front structure and a gas turbine engine

Abstract

The disclosure relates to a front structure, in particular an intermediate housing structure (50) for a compressor (14, 15) in a gas turbine engine (10), in particular an aircraft engine (10), comprising an outer annular component (51) and an inner annular component (52), wherein the annular components (51, 52), when assembled, are arranged concentrically about a main axis of rotation (9) of the gas turbine engine (10), and static blades (40, 41, 43) for deflecting at least one air flow (A, B) entering the front structure (50) are provided, wherein the blades (40, 41, 43) are radially connected to the annular components (51, 52), characterized in that at least some of the blades (40, 41, 43) are arranged in or on a blade module (44), in particular comprising one to five blades (40, 41, 43), very particularly comprising two to three blades (40, 41, 43), and the blade module (44) is connected to one of the annular components (51, 52) by at least one mechanical connecting means (47, 48). The disclosure also relates to a method for producing the front structure (50) and to a gas turbine engine (10) comprising a front structure (50).

Description

[0001] Front structure for a compressor device of a gas turbine engine, a method for manufacturing a front structure and a gas turbine engine

[0002] Description

[0003] The present disclosure relates to a front structure, in particular an intermediate housing structure for a compressor device of a gas turbine engine with the features of claim 1, a method for manufacturing a front structure, in particular an intermediate housing structure with the features of claim 17 and a gas turbine device with the features of claim 18.

[0004] Gas turbine engines, such as aircraft engines, typically have areas with very different operating conditions. For example, in an aircraft engine, the compressor, turbine, and combustion chamber areas differ significantly in their operating parameters. Therefore, it is necessary to use components adapted to the respective operating conditions (pressure, temperature, flow conditions, etc.).

[0005] It is known, for example, from US 2017 / 0241291 A1, to designate the turbine area in a

[0006] 2024P00186 Page 2

[0007] The aircraft engine is surrounded radially on the outside by an intermediate housing. Front structures, especially intermediate housings, have mechanical structural functions to absorb and / or distribute mechanical loads within the aircraft engine. Furthermore, the front structures, especially the intermediate housings, also have flow-guiding functions.

[0008] Other areas within an aircraft engine, such as the compressor section, may also have an intermediate casing structure (see, for example, US 2018 / 0023406A1). Such an intermediate casing structure is also referred to as a front structure.

[0009] It is therefore necessary to provide appropriate mechanically stable and weight-efficient front structures, especially intermediate housing structures.

[0010] One aspect of the solution proposed here concerns the arrangement of the front structure, specifically the intermediate casing structure, for a compressor in a gas turbine engine, particularly an aircraft engine. The front structure is composed of several components. Beneath these components, at least one static blade array is provided for deflecting at least one airflow entering the front structure. This airflow can be, for example, a bypass airflow or a core airflow. Furthermore, at least one ring is provided beneath the components, on which the blade arrays are radially arranged.

[0011] The front structure comprises an outer ring-shaped component and an inner ring-shaped component, the ring-shaped components being arranged concentrically around a main axis of rotation of the gas turbine engine when assembled. The front structure also includes static blades for deflecting at least one airflow entering the front structure, the blades being radially connected to the ring-shaped components. At least one portion of the blades is arranged in or on a blade module, in particular with one to five blades, and most especially with two to three blades, and the at least one blade module is connected to

[0012] 2024P00186 Page 3 connected to at least one of the ring-shaped components by a mechanical fastener.

[0013] Platforms at the upper and lower radial ends of the static blades (guide blades) of the blade modules can form an annular space in which little pressure loss occurs.

[0014] The arrangement of the static blades in blade modules allows for particularly rapid assembly of the blades into the front structure. Furthermore, the blade modules offer structural advantages and reduce weight. The blade modules are designed to be inserted into a pre-assembled section of the front structure and then mechanically connected to it, for example, with a detachable connection or with a material-fit or form-fit connection. This simplifies manufacturing – manually or automatically – compared to blades without a modular design.

[0015] In one embodiment, the at least one mechanical connecting element is designed as a flange structure on the radially inner side of the blade module, wherein this at least one flange structure is arranged on the front or rear side of the blade module. The flange structure is, for example, integrally connected to the blade module. A flange structure can, for example, have the form of a portion angled from a platform for the blades. This angled portion can then be connected to one of the components. The at least one flange structure of the blade module can, for example, engage in a recess of the inner annular component, thus forming a positive-locking connection.Additionally or alternatively, the at least one flange structure in the assembly can be attached with a rear mounting surface on the back of the inner annular component or a front mounting surface on the front of the inner annular ring structure.

[0016] Additionally or alternatively, this can include at least one mechanical fastener such as a welded joint, riveted joint, or screwed joint.

[0017] 2024P00186 Page 4 and / or adhesive bond.

[0018] The at least one flange structure can be connected to at least one of the ring-shaped components, for example, via a riveted connection, a bolted connection, a welded connection, an adhesive connection, and / or a bonded connection. Depending on the application, detachable or permanent connections can thus be created.

[0019] The at least one mechanical fastener can also include a positive-locking connection, in particular a dovetail joint, a friction-locking connection, in particular via a cone and spring connection, or a material-locking connection, in particular via adhesive bonding and / or bonding with composite material. In principle, it is possible to combine several types of mechanical fasteners in a front structure. For example, flange structures can be combined with friction-locking fasteners.

[0020] In one embodiment, at least one blade is hollow inside to accommodate, for example, a drive shaft for an external gearbox. The cavity can also or alternatively house oil lines, other air supply lines, or cables for sensors or power supply. This utilizes the space within the blade to connect an external gearbox, such as an accessory gearbox (AGB). Simultaneously, this at least one hollow blade also serves as an aerodynamically effective element.

[0021] In another embodiment, the blades are made of metal, CFRP, or another composite material, with the blades of a blade module being made of the same material. For example, the blades of a blade module can each be made of the same material, which simplifies the manufacturing of the blade modules. In particular, it is possible for all blades to be made of CFRP or a metal.

[0022] 2024P00186 Page 5

[0023] The front structure can be constructed as a hybrid lightweight component made of metallic and composite materials. In particular, one of the ring-shaped components can be made of composite material, especially fiber-reinforced composite material, or may incorporate composite material.

[0024] In one embodiment of the front structure, between 6 and 60, and in particular between 20 and 60, blades can be arranged in 1 to 60 blade modules. In principle, it is therefore possible to arrange all blades or individual blades in and / or on a single blade module.

[0025] For particularly efficient mechanical stability, in one embodiment the blades are made of metal and are load-bearing, especially in load-intensive areas of the front structure. Accordingly, composite blades are arranged in areas subject to reduced mechanical stress. This allows the overall structure to be mechanically stable while simultaneously being as lightweight as possible.

[0026] In a structure composed of several parts, gaps may exist between the individual elements. In one embodiment, a hardening fluid material is arranged in at least one gap in an annular component for sealing purposes. The hardening fluid (also called liquid shim) primarily serves to structurally close gaps. Liquid shims are used particularly for smaller gaps, while hard shims are used for larger ones.

[0027] Sealing is necessary because air flows on the inside of the front structure, and this air should not escape from the flow path. At least one disc element, in particular compensating sheets or shims, can also be arranged in the at least one gap to compensate for tolerances in the gap. This at least one disc element can also serve to secure the position of the at least one mechanical fastener. In addition to their compensating function, the disc elements would then also serve the at least one

[0028] 2024P00186 Page 6

[0029] Spalt also has a fixing function in conjunction with the shovel modules.

[0030] A key function of the front structure is the efficient guidance of airflow by means of the blades, which are at least partially grouped into blade modules. These blades are distributed radially around the main axis of rotation within the front structure. A blade module can have a platform radially outward and / or radially inward, which forms part of an annular channel within the front structure. In particular, the platforms of the blade modules can together form an annular channel within the front structure. The shape of the platforms can be adapted to the aerodynamic conditions. The inclination of the platforms can, for example, determine the cross-section in the axial direction.

[0031] The problem is also solved by a method for manufacturing a front structure, in particular an intermediate housing structure with the features of claim 17.

[0032] First, an inner ring-shaped component of the front structure is provided. Then, at least one blade module is connected to the inner ring-shaped component using at least one mechanical fastener, and subsequently, an outer ring-shaped component is connected to the assembly after the previous assembly step. The modular design, using different materials, allows for optimization based on the respective mechanical and / or thermal loads.

[0033] The problem is also solved by a gas turbine engine with the features of claim 18.

[0034] The invention is explained in connection with the embodiments shown in the figures.

[0035] Fig. 1 shows a side sectional view of an embodiment of a

[0036] 2024P00186 Page 7

[0037] Gas turbine engine with a front structure;

[0038] Fig. 2 shows a detailed view of a gas turbine engine with a front structure in the area of ​​the low-pressure compressor;

[0039] Fig. 3 shows a perspective view of an embodiment of a front structure;

[0040] Fig. 4A shows a first embodiment of a blade module with a blade;

[0041] Fig. 4B shows a second embodiment of a blade module with two blades;

[0042] Fig. 4C shows a third embodiment of a blade module with three blades;

[0043] Fig. 5 shows a front view of the front structure according to Fig. 3, indicating views and sectional views;

[0044] Fig. 5A is a side view according to view D in Fig. 5;

[0045] Fig. 5B is a sectional view through section BB in Fig. 5;

[0046] Fig. 5C shows a cross-sectional view through the radially inner part of a blade module;

[0047] Fig. 5D shows a sectional view through section CC in Fig. 5;

[0048] Fig. 5E a sectional view through section EE- in Fig. 5;

[0049] Fig. 6 shows a perspective view of the radially inner ring-shaped component of the front structure;

[0050] Fig. 6A is a detail view F from Fig. 6;

[0051] Fig. 6B shows another detailed view of Fig. 6 according to view G in Fig. 6A.

[0052] 2024P00186 Page 8

[0053] Fig. 1 shows a gas turbine engine 10 with a main axis of rotation 9. The gas turbine engine 10 includes an air inlet 12 and a fan 23, which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11, which receives the core airflow A. The core engine 11 comprises, in axial flow order, a low-pressure compressor 14, a high-pressure compressor 15, a combustion unit 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core thrust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass channel 22 and a bypass thrust nozzle 18. The bypass airflow B flows through the bypass channel 22. In the illustrated embodiment, the fan 23 is connected to the low-pressure turbine 19 via a shaft 26 and an epicyclic planetary gear 30 and is driven by it.In principle, however, it is possible that the gas turbine engine 10 is also designed without a reducing planetary gearbox 30.

[0054] During operation, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15, where further compression takes place. The compressed air expelled from the high-pressure compressor 15 is directed into the combustion unit 16, where it is mixed with fuel and the mixture is combusted. The resulting hot combustion products then propagate through and drive the high-pressure and low-pressure turbines 17 and 19, respectively, before being expelled through the nozzle 20 to provide thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable connecting shaft 27. The fan 23 generally provides the majority of the thrust.

[0055] A front structure 50 is arranged axially at the exit of the engine nacelle 21, the function of which will be described in more detail in connection with Figures 2 to 6B. One embodiment of a front structure 50 is an intermediate housing structure.

[0056] 2024P00186 Page 9

[0057] In other embodiments, the engine nacelle 21 can extend axially over the area of ​​the low-pressure compressor 14, up to the area of ​​the high-pressure turbine 17.

[0058] It is noted that the terms “low-pressure turbine” and “low-pressure compressor,” as used herein, may be understood to mean the lowest-pressure turbine stage and the lowest-pressure compressor stage, respectively (i.e., excluding Fan 23), and / or the turbine and compressor stages connected by the lowest-rotating connecting shaft 26 in the engine (i.e., excluding the gearbox output shaft driving Fan 23). In some writings, the “low-pressure turbine” and “low-pressure compressor” referred to herein may alternatively be known as the “intermediate-pressure turbine” and “intermediate-pressure compressor.” When using such alternative nomenclature, Fan 23 may be described as a first compression stage or the lowest-pressure compression stage.

[0059] Optionally, the gearbox can drive auxiliary and / or alternative components (e.g. the intermediate pressure compressor and / or a secondary compressor).

[0060] Other gas turbine engines to which the present disclosure may apply may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of connecting shafts. As another example, the gas turbine engine shown in Fig. 1 has a split-flow nozzle 20, 22, which means that the airflow B through the bypass channel 22 has its own nozzle, separate from and radially outside the engine core nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass channel 22 and the flow through the core 11 are mixed or combined upstream of (or before) a single nozzle, which may be referred to as a mixing-flow nozzle.

[0061] 2024P00186 Page 10 or both nozzles (whether mixed or split flow) can have a fixed or variable range.

[0062] The geometry of the gas turbine engine 10 and its components is / are defined by a conventional axis system comprising an axial direction (aligned with the main axis of rotation 9), a radial direction (in the bottom-to-top direction in Fig. 1), and a circumferential direction (perpendicular to the view in Fig. 1). The axial, radial, and circumferential directions are perpendicular to each other.

[0063] The front structure 50 described below is designed as a hybrid component (HIMC hybrid intermediate casing), meaning that plastic and metal are used as materials. Figure 2 shows a detailed view of part of the front structure 50 in the area of ​​the low-pressure compressor 14.

[0064] In the hybrid front structure 50, at least one first component is made of or incorporates composite material, in particular fiber composite material, and at least one second component of the front structure 50 is made of or incorporates metal. By distributing the materials across different components, flexible options are created for a load-bearing and weight-efficient design.

[0065] Figure 3 shows an embodiment of a front structure 50 in a perspective view. This front structure 50 and its details are described below with reference to Figures 3 to 6B.

[0066] Essentially, the front structure 50 has the shape of a flat cylinder, with its radial extent around the main axis of rotation 9 being significantly larger than its axial extent. In principle, other size ratios are also possible, for example, if other components are connected to or formed integrally with the front structure 50 in the axial direction in front of and / or behind it.

[0067] 2024P00186 Page 11

[0068] On the axial side surfaces, a front flange structure 45 is arranged at the front in the direction of flight F and a rear flange structure 46 is arranged at the rear in the direction of flight F, via which the front structure 50 can be connected to other components of the gas turbine engine 10 (not shown here).

[0069] The front structure 50 has two concentric ring structures: an outer ring-shaped component 51 and an inner ring-shaped component 52. The outer ring-shaped component 51 and the blades 40, 41, 43 (with blade modules 44) are described first (Figs. 4 and 5). The inner ring-shaped component 52 and its connection to the blade modules 44 are then described (Figs. 6, 6A, 6B).

[0070] Fig. 5 shows a front view of the front structure 50 according to Fig. 3, with elevations and sectional views indicated. The side view D is shown in Fig. 5A. The section BB through the outer annular component 51 is shown in Fig. 5B. Fig. 5C shows a sectional view through the radially inner part of a blade module 44. The section CC through the outer annular component 51 is shown in Fig. 5D. The section EE through a hollow blade 43 and the outer annular component is shown in Fig. 5E.

[0071] The blades 40 and 41 are arranged radially between the annular components 51 and 52. Static blades 40 (guide blades, stators, vanes) made of composite material and static blades 41 made of metal are used, with each blade 40 and 41 being attached to the annular component 51 by a connecting element 42. The attachment can be achieved via welds, a positive-locking connection, or by a single-piece manufacturing process.

[0072] The blades 40, 41, 43, i.e., metallic and non-metallic, are connected with mechanical fasteners 42 (e.g. rivets, Hi Lock).

[0073] The outer annular component 51 is connected to the blades (see Fig. 3, 5D) by means of flat-head rivets. The distal ends of the blades 40, 41, 43 lie on the inner surface of the outer annular component 51.

[0074] 2024P00186 Page 12 flush for aerodynamic reasons. The connecting elements 42 also serve to connect the transmission connecting elements 57 and the drive shaft housing 60 to the outer annular component 51.

[0075] The blades 40, 41 serve to redirect the bypass airflow B and to transfer the forces and moments occurring in the process into support structures of the gas turbine engine 10 and / or aircraft structures not shown here.

[0076] In this embodiment, the metallic blades 41 are arranged in the areas where the mechanical loads on the blades 40, 41 are relatively high. The blades 40 can be arranged in blade modules 44, with the blades 40, 41 being fastened via the blade modules 44. In principle, all blades 40, 41, 43 can be made of CFRP or metal.

[0077] Figures 4A, 4B, and 4C show embodiments of different blade modules 44. This means that one or more blades 40, 41 are connected at their radially inner ends to a common platform 66. The connection can be made, for example, by welding or in one piece. In the illustrated embodiment, the blades 40, 41 were manufactured in one piece, with a rounded transition at the base of the blades 40, 41 to prevent a notch effect. In the illustrated embodiment, the platform 66 has blade flanges 47, 48 at the front and rear of the base, via which the blade modules 44 can be attached to the inner annular component 52 (see Figures 5C, 6, 6A, and 6B). Alternatively or additionally, the at least mechanical connecting element can be a welded connection, riveted connection, bolted connection, and / or adhesive bond.This allows the connection of platform 66 to be carried out in the same way as the connection of an outer platform 70 with the outer ring-shaped component 51.

[0078] Figure 5C shows, for example, a sectional view through the radially inner part of a blade module 44. The front blade flange 48 is shown on the left, the rear blade flange 47 on the right. The two blade flanges 47,

[0079] 2024P00186 Page 13

[0080] The 48 components are integrally connected / manufactured here with the platform 66 of the blade module 44. A fire protection device 71 in the form of an inwardly extending annular surface serves to protect the rear side of the blade module 44 for the composite blades 40, in order to protect them from hot gases from the compressor.

[0081] The type of connection with the inner ring-shaped component 52 is then shown in Fig. 6A, 6B.

[0082] Figure 4A shows a single blade 40 in this form. Figures 4B and 4C show blade modules 44 with two and three blades 40, respectively. In principle, the blade modules 44 can have blades 40 made of composite material and / or blades 41 made of metal. A blade module 44 can also have more than three blades 40, 41. A hollow blade 43, which is described below, can also be part of a blade module 44. Manufacturing is simpler if the blade module 44 is made of the same material.

[0083] The blade modules 44 particularly facilitate assembly and maintenance should blades 40, 41 ever need to be replaced. In the embodiment according to Figures 3 and 5, twenty blade modules 44 are used. The use and arrangement of blade modules 44 with composite blades 40 together with the inner annular component 52 are then explained in connection with Figures 5C, 6, 6A, 6B.

[0084] One of the metallic blades 43 (see Figs. 3, 5E) is hollow, allowing a radial drive shaft for an external gearbox (accessory gearbox AGB), which is also not shown here, to pass through the cavity (not visible in Fig. 3). However, Fig. 5A shows gearbox connecting elements 57 arranged on the outside of the outer annular component 51. The embodiment shown here has only one hollow blade 43. In principle, more than one hollow blade 43 can also be used.

[0085] 2024P00186 Page 14

[0086] Figure 5E shows a sectional view through part of the hollow metallic blade 43, with the drive shaft omitted for clarity. The hollow metallic blade 43 is connected to the outer annular component 51 shown below. In the transition area from the hollow metallic blade 43 to the external gearbox, an annular seal 58 (e.g., an O-ring) is arranged, sealing against a flange element 59 located on the outside of the outer annular component 51. The flange element 59 is part of the drive shaft housing 60 (see Figure 5A).

[0087] As shown in Fig. 5D, a gap 55 between the outer platforms 70 of the blades 40, 41, 43 (not visible in this sectional view) and the outer annular component 51 (here made of composite material) can, in one embodiment, be filled with a fluid material that hardens after assembly. Additionally or alternatively, disc elements 56, such as shims or shims, can be used to compensate for tolerances in the gap 55. This is shown schematically in Fig. 5D. The disc elements 56 can also be used to secure the position of the mechanical fasteners 47, 48 on the blade modules 44, thus minimizing or preventing air leakage from the bypass channel 22.

[0088] In principle, it is possible to produce the outer ring-shaped component 51 in a lay-up process or an additive manufacturing process, thus avoiding costly casting processes.

[0089] In the axial direction in front of and behind the areas of the blades 40, 41, 43, liners 53 are arranged on the inside of the outer annular component 51 (see Fig. 5B, 5D). Fig. 5D further shows that load distribution elements 62 are arranged radially distal to the front and rear flange structures 45, 46. These also protect against thermal loads from outside the front structure 20. Thermal protection is also provided by a [missing information - likely a specific element or component] on and / or in the [missing information - likely a specific element or component].

[0090] 2024P00186 Page 15 radially outer surface of the outer annular component 51 arranged fire protection layer 63. As can be seen in Fig. 5D, this fire protection layer 63 protects the connecting elements 42 of the blades 40, 41, 43 (not visible here), which are connected to the outer annular component 51 via the upper platform 70.

[0091] The front structure 50 can be connected to the rest of the gas turbine engine 10, with several fastening elements 54 arranged on the outer side of the annular component 51 (see Fig. 3, 5B). These fastening elements 54 serve to connect the gas turbine engine 10 to the aircraft and to connect the external gearbox (not shown) and other attachments to the front structure 50. The connection of the front structure 50 to the rest of the gas turbine engine 10 is primarily achieved via flanges on the outer annular component 51 and on the inner stroke.

[0092] These areas are subjected to particularly high mechanical loads during operation. Therefore, a sandwich structure is created consisting of the outer platform 70 (and thus the radially inner blades 40, 41, 43), the ring-shaped outer component 51, and a fastening element 54, which are held together, for example, by bolted connections. This makes these areas particularly resistant to mechanical stresses.

[0093] Furthermore, transport lugs 61 are arranged on the outside of the outer ring-shaped component 51, which are integrally connected to the fastening elements 54. The transport lugs 61 are optimized with regard to weight, load-bearing capacity, and assembly time.

[0094] Figures 6, 6A, and 6B now describe details of the inner ring-shaped component 52. Figure 6 shows a perspective view of this part.

[0095] In the embodiment shown in Fig. 6, metallic blades 41 extend radially from the inner annular component 52, which is arranged concentrically around

[0096] 2024P00186 Page 16 the main axis of rotation 9 is arranged. As mentioned above, the metallic blades 41 can be connected to the inner annular component 52 by welds, positive-locking connections, or by a one-piece manufacturing process. The composite blades 40 are not shown in Fig. 6. The blade modules 44 with the composite blades 40 are arranged in the spaces between the metallic blades 41, which will be described further in connection with Figs. 6A and 6B.

[0097] Radially below the blades 41 are support structures 64 (struts) that connect the blades 41 to the actual inner annular component 52. The support structures 64 are designed to gently transfer mechanical loads acting on the blades 41 into the front structure 50. The support structures 64 are described in more detail in Fig. 6A (highlighted as detail F in Fig. 6).

[0098] The metallic blades 41 are arranged radially inside each of an aerodynamically effective guide element 65, which covers the underlying support structure 64 of the blade 41. The bypass airflow B thus flows past the blades 41 and the guide element 65. The guide element 65 is essentially plate-shaped, with the plate inclined forward in the direction of the main axis of rotation 9, so that the flow cross-section for the bypass airflow B decreases in the flow direction.

[0099] The inner platform 66 for composite blades 40 (see also Figs. 4A, 4B, 4C, 5C) also provides an aerodynamically effective guide surface for the bypass airflow B. This platform 66 is part of the blade module 44, which is not shown in Figs. 6, 6A, 6B for clarity. The blade modules 44 are inserted into the gaps between the metal blades 41.

[0100] The rear side of the inner platform 66, facing in the direction of flight F, is connected to or integrally formed with the rear blade flange 47 (see Figs. 4A, 4B, 4C) of the blade module 44. This rear blade flange 47 can be arranged in a recess 67 of the inner annular component 52. This one

[0101] The rear blade flange 47 (not shown) of the blade modules 44 (see Fig. 4A, 4B, 4C) of the blade modules 44 (2024P00186) can then be connected to a slightly protruding rear mounting surface 69 on the rear of the inner annular component 52. This is shown in Fig. 6B, which corresponds to view G in Fig. 6A.

[0102] Accordingly, the front blade flange 48 (see Fig. 4A, 4B, 4C) of the blade modules 44 can be connected to a front mounting surface 68 on the front of the inner annular component 52 (see Fig. 6).

[0103] 2024P00186 Page 18

[0104] Reference symbol list

[0105] 9 Main axis of rotation

[0106] 10 Gas turbine engine

[0107] 11 Core engine

[0108] 12 Air intake

[0109] 14 low-pressure compressors

[0110] 15 high-pressure compressors

[0111] 16 Combustion unit

[0112] 17 High-pressure turbine

[0113] 18 Bypass thrust nozzle

[0114] 19 Low-pressure turbine

[0115] 20 Core thrust nozzle

[0116] 21 Engine nacelle

[0117] 22 Bypass channel

[0118] 23 Fan

[0119] 26 wave

[0120] 27 Connecting shaft

[0121] 30 gearboxes

[0122] 40 shovels (composite material)

[0123] 41 Shovel (metal)

[0124] 42 Connecting element between shovel and ring-shaped component

[0125] 43 hollow shovel

[0126] 44 Bucket module

[0127] 45 front flange structure

[0128] 46 rear flange structure

[0129] 47 rear shovel flange

[0130] 48 front shovel flange

[0131] 50 Front structure

[0132] 51 outer ring-shaped component of the front structure

[0133] 2024P00186 Page 19

[0134] 52 inner ring-shaped component of the front structure

[0135] 53 liners

[0136] 54 Fastening element

[0137] 55 gap

[0138] 56 disc element

[0139] 57 Gearbox connecting devices

[0140] 58 Seal

[0141] 59 Flange element

[0142] 60 drive shaft housings

[0143] 61 Transport eyelet

[0144] 62 load distribution elements

[0145] 63 Fire protection layer

[0146] 64 support structures

[0147] 65 aerodynamic guide element

[0148] 66 inner platform, especially for composite material shovels

[0149] 67 Exclusion

[0150] 68 front mounting surface

[0151] 69 rear mounting surface

[0152] 70 outer platform for shovels

[0153] 71 Fire protection device

[0154] A core airflow

[0155] B Bypass airflow

[0156] 2024P00186

Claims

Page 20 Patent claims 1. Front structure, in particular intermediate housing structure (50) for a compressor (14, 15) in a gas turbine engine (10), in particular an aircraft engine (10), comprising an outer annular component (51) and an inner annular component (52), wherein the annular components (51, 52) are arranged concentrically around a main axis of rotation (9) of the gas turbine engine (10) in the assembly, and static blades (40, 41, 43) are provided for deflecting at least one airflow (A, B) entering the front structure (50), wherein the blades (40, 41, 43) are connected radially to the annular components (51, 52), characterized in that at least a part of the blades (40, 41, 43) are in or on a blade module (44), in particular with one to five blades (40, 41, 43), in particular with two to three blades (40, 41, 43), are arranged and the blade module (44) is connected by at least one mechanical connecting means (47,48) is connected to one of the ring-shaped components (51, 52).

2. Front structure (50) according to claim 1, characterized in that the at least one mechanical connecting means is designed as a flange structure (47, 48) on the radially inner side of the blade module (44) and the at least one flange structure (47, 48) is arranged on the front or rear side of the blade module (44) and / or the at least one mechanical connecting means is designed as a welded connection, riveted connection, screwed connection and / or adhesive connection.

3. Front structure (50) according to claim 2, characterized in that the at least one flange structure (47, 48) of the blade module (44) engages in a recess (67) of the inner annular component (52). 2024P00186 Page 21 4. Front structure (50) according to claim 2 or 3, characterized in that the at least one flange structure (47, 48) is attached in assembly with a rear mounting surface (69) on the rear of the inner annular component (52) or a front mounting surface (68) on the front of the inner annular ring structure (52).

5. Front structure (50) according to at least one of claims 2 to 4, characterized in that the at least one flange structure (47, 48) is connected to at least one of the ring-shaped components (51, 52) via a rivet connection, a screw connection, a weld connection, an adhesive connection and / or a bonding connection.

6. Front structure (50) according to claim 6 according to at least one of the preceding claims, characterized in that the at least one mechanical connecting means has a positive locking connection, in particular a dovetail connection, a force locking connection, in particular via a cone and spring connection or a material locking connection, in particular via gluing and / or bonding with composite material.

7. Front structure (50) according to at least one of the preceding claims, characterized in that at least one blade (43) is hollow inside to accommodate a drive shaft for an external gearbox.

8. Front structure (50) according to at least one of the preceding claims, characterized in that the blades (40, 41, 43) are made of metal, CFRP or another composite material, wherein the blades (40, 41, 43) of a blade module (44) are in particular made of the same material.

9. Front structure (50) according to at least one of the preceding claims, characterized in that a component (51 , 52) is made of composite material, in particular fiber composite material, or comprises composite material. 2024P00186 Page 22 10. Front structure (50) according to at least one of the preceding claims, characterized in that a total of between 6 and 60, in particular between 20 and 60 blades (40, 41, 43) are arranged in and / or on 1 to 60 blade modules (44).

11. Front structure (50) according to at least one of the preceding claims, characterized in that the blades (41 ) are made of metal and are load-bearing, in particular in load-intensive areas of the front structure (50).

12. Front structure (50) according to at least one of the preceding claims, characterized in that a hardening fluid material is arranged for sealing in at least one gap (55) in an annular component (51 , 52).

13. Front structure (50) according to at least one of the preceding claims, characterized in that at least one disk element (56), in particular a disk, is arranged in the at least one gap (55) to compensate for tolerances in the gap (55).

14. Front structure (50) according to claim 11, characterized in that the at least one disc element (56) serves to secure the position of the at least one mechanical connecting means (47, 48).

15. Front structure (50) according to at least one of the preceding claims, characterized in that the blade module (44) has a platform (66, 70) radially outside and / or radially inside, which is formed as part of an annular channel in the front structure (50).

16. Front structure (50) according to claim 15, characterized in that the platforms (66, 70) of the blade modules (44) together form an annular channel. 2024P00186 Page 23 17. Method for manufacturing a front structure, in particular an intermediate housing structure (50) according to at least one of claims 1 to 16, characterized in that a) an inner annular component (52) is provided, then b) at least one blade module (44) is connected to the inner annular component (52) with at least one mechanical connecting means (47, 48), and then c) an outer annular component (51) is connected to the assembly according to step b).

18. Gas turbine engine (10) with at least one front structure (50) according to at least one of claims 1 to 16. 2024P00186

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