Patent Application
20190093488
The present disclosure concerns an aerofoil component of a turbomachine and a method of manufacturing an aerofoil component for a turbomachine.
Turbomachines, such as gas turbine engines, use rotors which comprise a plurality of aerofoil components, typically referred to as blades. Such rotors may be used, for example, in the fan, compressors and turbines. The blades are often welded to a central disk or ring to form a monolithic component referred to as a blisk (bladed disk) or bling (bladed ring).
The blades are typically manufactured from Titanium, such as Titanium 6AI-4V (Ti6-4), and are welded to the disk or ring using a solid state welding process, such as linear friction welding. The disk or ring is also typically formed of Titanium and so the resulting component is formed of a single material.
Titanium can be limited in its high cycle fatigue capability and this produces limitations in the design. These limitations result in additional thickness in the aerofoil form, reducing fan efficiency and adding additional weight in the component. Compressor aerofoils are also affected by phenomenon such as aerodynamic flutter which, in order to protect against such events, presents further such design limitations.
It is therefore desired to provide an aerofoil component which addresses these issues.
According to an aspect there is provided an aerofoil component for a turbomachine, the aerofoil component comprising: a central core formed from a metal matrix composite; and an external layer comprising a pressure surface, a suction surface, a leading edge, a trailing edge and a root, the external layer being formed by a metal which covers the metal matrix composite of the central core.
The external layer may further comprise a tip such that the external layer entirely encapsulates the metal matrix composite of the central core.
The metal matrix composite may comprise a reinforcing material in a metal matrix.
The metal matrix may be formed from the same metal as the external layer. In some examples, the metal matrix may be formed from the same base metal as the external layer, but may be formed from a different alloy. However, in other examples, the same alloy may be used for both the metal matrix and the external layer.
The reinforcing material may be a particulate material.
The reinforcing material may be titanium boride or titanium carbide.
The external layer may be formed from titanium (including alloys of titanium).
A plurality of aerofoil components may be used to form a rotor. The aerofoil components may be joined to a hub via the root. For example, the aerofoil components may be joined to the hub using solid state welding or diffusion bonding.
The central core of one or more of the aerofoil components may be spaced a radial distance from its root which is different to that of one or more of the other aerofoil components.
According to another aspect there is provided a method of manufacturing an aerofoil component for a turbomachine, the method comprising: covering a central core formed from a metal matrix composite within an external layer formed by a metal to form a blank; consolidating the blank to form an intermediate form; and forging the intermediate form to form the aerofoil component with the external layer surrounding the central core of metal matrix composite and forming a pressure surface, a suction surface, a leading edge, a trailing edge and a root.
The external layer may additionally form a tip such that the external layer entirely encapsulates the metal matrix composite of the central core.
The blank may be consolidated by extrusion.
The blank may be consolidated by rolling.
The blank may be consolidated by hot isostatic pressing.
The intermediate form may be cut prior to forging so as to determine a radial distance of the central core from the root in the forged aerofoil component.
The method may further comprise connecting a plurality of said aerofoil components to a hub to form a rotor, wherein the central core of one or more of the aerofoil components is spaced a radial distance from its root which is different to that of one or more of the other aerofoil components.
A plurality of central cores may be covered by the external layer, and the method may further comprise: cutting the intermediate form into a plurality of sections each comprising a central core covered by the external layer and then forging the sections to form a plurality of aerofoil components.
The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein.
Embodiments will now be described by way of example only, with reference to the Figures, in which:
With reference to
The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by suitable interconnecting shaft.
Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
The aerofoil portion 28 comprises a pressure surface 30 and an opposing suction surface (not visible). A leading edge 32 and a trailing edge 34 are defined between the opposing pressure and suction surfaces along the lateral sides of the aerofoil portion 28. The aerofoil portion 28 extends the root portion 26 to a tip 36 at its distal, free end.
The blade 24 is fabricated from a composite material. Specifically, the root portion 26 and the external surfaces of the aerofoil portion 28 (i.e. the pressure and suction surfaces, the leading and trailing edges, and the tip) are formed from a first material. A central core 38 formed from a second material is provided within the aerofoil portion 28 and surrounded by the first material. The central core 38 extends in a span wise direction between the root portion 26 and the tip 36, and in a chord wise direction between the leading and trailing edges 32, 34.
The first material is a metal or metal alloy. Specifically, in this example, the first material is Ti6-4. The second material is a metal matrix composite consisting of a metal matrix and a reinforcing material. In this example, the reinforcing material is a particulate material which may be formed from, for example, a ceramic, such as TiC or TiB. The metal matrix is formed from the same material as the first material and so is also Ti6-4 in this example. In other examples, the metal matrix may be formed from the same base metal, but a different alloy.
In step 1, a blank is formed which comprises the first material (i.e. the metal or metal alloy) and the second material (i.e. the metal matrix composite). This may be achieved as shown in
In
In
In step 2, the blank of
In step 3, the consolidated blank is then extruded (using conventional procedures) to form a bar, as shown in
The extruded bar is then forged to form the blade 24 in step 4. The blades 24 may be forged close to the final required aerodynamic size and shape. However, the blades 24 may undergo some final finishing, post forging, such as machining, welding, heat treating, polishing and inspection.
In another example, the tube 40 of
The blade 24 is selectively reinforced, comprising a reinforced core, but with all outer surfaces being unreinforced, including the root and tip. The central core 38 reinforces the blade 24, thereby increasing its stiffness. Consequently, the fatigue loading and the susceptibility to flutter is reduced compared to a blade formed entirely from the first material. However, the first material is used for all external surfaces of the blade 24, particularly the root portion 26 and so allows existing linear friction welding parameters to be used to attach the blade 24 to the hub.
The increased stiffness due to the reinforcement of the aerofoil portion 28 reduces the fatigue stress at the peak limiting location for a given engine load, resulting in an increased component life.
Alternatively, the blade 24 could be redesigned to reduce the thickness of the aerofoil portion 28 (while retaining the same stiffness) to improve the aerodynamic efficiency of the fan 13.
The blade 24 utilises a material with good damage tolerance properties (e.g. Ti6-4) on the leading edge where the blade 24 is susceptible to foreign object damage (FOD), while the overall blade 24 benefits from the strength and stiffness of the core 38.
Further, having a leading edge formed from a single material (e.g. Ti) allows the use of existing material addition repair techniques, thereby reducing the life cycle cost of the component.
The specific method described above provides a blade having a reinforced core, whilst utilising existing extrusion and forging techniques.
The second material is chosen to provide the required increase in stiffness, but with a flow stress that is well matched to the first material during the extrusion step.
The method also allows the radial position of the reinforcing core 38 to be varied simply by selecting the appropriate cutting position during preparation of the extruded bar for forging. This presents the opportunity to produce a set of blades 24 that are deliberately “mis-tuned” (i.e. their individual dynamic response is different). This may reduce the risk of flutter and enable a lighter or more efficient design capable of meeting the required design criteria for flutter.
Although the first material has been described as being a titanium alloy, it will be appreciated that other materials could be used. Similarly, other materials with increased stiffness which are capable of being bonded (either directly or indirectly) to the base material during the HIP stage (or via any other consolidation process) and capable of being extruded during the extrusion stage may be used for the core.
Although the blade has been described with reference to a fan rotor, it will be appreciated that it may be used in other aerofoil components, particularly for blades found elsewhere in a gas turbine engine, such as in compressors and turbines. It may also be used in other types of turbomachines, such as steam turbines.
Although it has been described that the core is entirely encapsulated within the first material, in other examples the core may only be partially covered by the external layer of first material. In particular, the core may be exposed at its tip.
It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.
This specification is based upon and claims the benefit of priority from U.S. Provisional Patent Application No. 62/561729 filed on 22 Sep. 2017, the entire contents of which are incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| 62561729 | Sep 2017 | US |
