METHOD OF FORMING HYBRID METAL CERAMIC COMPONENTS

Abstract

A monolithic composite turbine component includes at least one first region of a first material and one second region of a second material formed by solid freeform fabrication (SFF). The first material may be a metal and the second material may be a ceramic or a ceramic matrix composite. Transition regions between the metal region and ceramic region are functionally graded regions to minimize internal stress during temperature fluctuations.

Description

FIELD OF THE INVENTION

This invention relates to turbine engine components.

BACKGROUND

In the pursuit of ever higher efficiencies, gas turbine manufacturers have long relied on continually increasing turbine inlet temperatures to provide boosts to overall engine performance. In typical modern engine applications, the gas path temperatures within the turbine exceed the melting point of some component materials. As a result, dedicated cooling air extracted from the compressor is used to cool the gas path components in the engine incurring significant cycle penalties especially when cooling is utilized in the low pressure turbine (sometimes also referred to as the power turbine).

To reduce these cycle penalties, much research has gone into implementing high temperature materials into the construction of these components. A popular area of research material is the development of ceramic matrix composites (CMC's). Turbine applicable CMC's are typically composed of silicon-carbon fibers in silicon or boron doped silicon infiltration matrices. These CMC's are laid up in alternating plies whose orientation is tailored to the intended direction of maximum tensile loads. Other materials of use are monolithic ceramics made of, for example, silicon or silicon compounds. While having significantly higher temperature capabilities, strengths compared to metallic parts are orders of magnitude lower limiting the capability of these materials in implementation. Further, silicon based components are susceptible to water degradation, and exhibit silicon migration into the surrounding metallic parts that they may come in contact with at elevated temperatures.

Ceramics offer benefits of higher temperature capability and lower cooling requirements which correlate to improved efficiency and reduced emissions. One of the main challenges of implementing ceramics for high pressure turbine application, is their low thermal shock resistance and low overall strength. In an effort to overcome these limitations, development of numerous structural hybrid designs and joining techniques have been made available that combine metal and ceramic elements that minimize thermal stresses and thermal expansion rates. These mechanical interfaces, however, have been shown to be unreliable for implementation into highly stressed applications, such as a turbine blade or vane. The interfaces must be specially coated for contact with metallic parts due to abrasive wear and silicon migration. Further, mechanical interfaces necessitate a sealing mechanism which incurs a cooling leakage penalty that increases the need to flow cooling air into the part.

Stress free ceramic to metal joins in these components would offer improved performance.

SUMMARY

A monolithic composite turbine component includes at least one region of a first material attached to a second region of a second material by solid freeform fabrication (SFF). The first region is preferably metal and the second region is preferably ceramic or ceramic matrix composite. The transition regions between the first and second regions are functionally graded to minimize internal stress during operation.

In an embodiment, a method of forming a monolithic composite turbine component includes forming a first region attached to a second region by solid freeform fabrication (SFF) additive manufacturing. The transition regions between the first and second regions are functionally graded to minimize internal stress during operation.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic representation of a two component powder bed additive manufacturing process.

FIG. 2 is a perspective view of a monolithic composite turbine nozzle with functionally graded material transitions.

FIG. 3 is a schematic representation of a two component selective laser melting powder deposition additive manufacturing process.

FIG. 4 is a schematic representation of a monolithic composite component of the invention.

DETAILED DESCRIPTION

The thermodynamic efficiency of a gas turbine engine scales directly as the absolute temperature difference between the inlet and the exhaust temperatures of the working fluid in the gas path of the engine. The drive to use materials with increased temperature capability is continually hindered by the intrinsic structural properties of turbine materials at elevated temperatures. This difficulty is being addressed by the use of composite materials with widely differing thermal mechanical response properties. Ceramics exhibit high temperature resistance to deformation, but low resistance to fracture even at high temperatures. Metals, in contrast, exhibit high resistance to deformation and fracture, but are limited to applications in the vicinity of their melting temperatures. Combinations of metal and ceramic structures in composite materials such as ceramic matrix composites (CMC's) wherein the filler (metal or ceramic) is a fiber or otherwise elongated form of material and the matrix is a ceramic formed by, for instance, molten or gaseous infiltration, can exhibit both high temperature fracture resistance due to fiber reinforcement and environmental degradation due to erosion.

Situations exist, however, in engine design wherein the beneficial and contrasting, properties of ceramics, metals and CMC's are all required in a single composite component. In this case, the joints between separate parts of a component may be a weak link due to thermal and thermal mechanical property disparities across each joint in a cyclic thermal gradient. Thermal expansion discontinuities across, for instance, a metal and ceramic or CMC interface may result in measureable stress concentrations leading to environmental degradation at the joint and possible failure.

With the advent of solid freeform fabrication (SFF) technology using, for instance, additive manufacturing techniques, turbine components exposed to varying severe environments in a hot gas path can be formed wherein different materials can be used with positive results in areas of a part exposed to different environments.

Solid freeform fabrication is a process wherein three-dimensional (3D) objects are produced by a layer by layer technique or the deposition of individual solid, liquid or semi-solid portions of a material on specific regions of an object according to a digital model of the object stored in the memory of a computer controlled deposition apparatus. Common energy sources used in SFF are laser and electron beams. Metals, ceramics, and polymers can be used as build materials. Direct SFF fabrication of useful components are now known in the art.

Powder based layer by layer additive SFF manufacturing processes suitable for the present invention include selective laser sintering (SLS), direct laser sintering (DLS), selective laser melting (SLM), direct laser melting (DLM) and others known in the art. A preferred technique for the present invention is direct laser melting.

Powder based direct material deposition SFF manufacturing processes include direct laser melting (DLM), direct laser deposition (DMD), laser engineered net shaping (LENS), shape deposit manufacturing (SDM), directed light fabrication (DLF), three dimensional cladding and others known in the art. A preferred technique for the present invention is direct laser melting.

An embodiment of the present invention is the fabrication of a monolithic composite turbine component containing regions of at least two different materials, preferably ceramic and metal by solid freeform fabrication (SFF). This embodiment requires at a minimum, at least two powder sources. A layer by layer fabrication embodiment of the invention is shown in FIG. 1. Powder bed SFF additive manufacturing process 10 includes manufacturing chamber 12 containing devices that produce two component SFF objects by powder bed additive manufacturing. An example of process 10 is selective laser melting (SLM). SFF SLM process 10 includes powder storage chambers 14 and 16. In a preferred embodiment, storage chamber 14 contains metal powder 24 and storage chamber 16 contains ceramic powder 26. SFF SLM process 10 further includes build chamber 18, laser 20, scanning mirror 22, piston 28 in storage chamber 14, piston 30 in build chamber 18 and piston 32 in storage chamber 16.

During operation of SFF SLM process 10, metal powder 24 is fed upward by piston 28 and spread over build platform 30 by roller 34. After powder 24 is spread on build platform 30 by roller 34, roller 34 retracts to position 42 shown in phantom lines and laser 20 and scanning mirror 22 are activated by a control system (not shown) to melt selective areas of surface S of powder 24 in build chamber 18 according to a computer model of solid free form object 36 to form a single solidified layer of SFF object 36.

In the next step, piston 28 advances to expose another layer of metal powder 24 in chamber 14 and piston 30 recedes in build chamber 18 to accept another layer of powder. Roller 34 then advances to spread another layer of powder 24 on build chamber 18 and then retracts to position 42 while laser 20 and scanning mirror 22 are activated to fuse a selected area of surface S to form another layer of SFF object 36. The process continues until, for instance, feature 38 is formed from metal powder 24.

In an embodiment, if, at this point, it is desired to form ceramic component 44 on metal feature 38, from ceramic powder 26 in storage chamber 16, the process changes. Roller 34 advances to position 44, shown in phantom lines. Piston 32, indexes up one layer thickness to expose ceramic powder 26 to roller 34. Build platform 30 indexes down one layer thickness and roller 34 proceeds to spread one layer of ceramic powder 26 on build chamber 18 creating new ceramic surface S. Roller 34 then retracts to position 44. Laser 20 and scanning mirror 22 are activated to fuse selected areas of ceramic powder 26 at surface S. Piston 32 indexes upward one layer of thickness and build platform indexes downward one layer thickness. Roller 34 then spreads another layer of ceramic powder 26 on the solidified surface of ceramic powder 26. Laser 20 and scanning mirror 22 are then activated to fuse selected areas of ceramic powder 26 at surface S according to a computer model and control system (not shown) of SLM process 10 to form subcomponent 40 of component 36.

If another subcomponent, such as metal subcomponent 42 is to be added to component 36 on subcomponent 40, the process is changed to a process forming metal addition 42 attached to subcomponent 40, that is similar to the process forming metal subcomponent 38 as discussed above.

In an embodiment, SFF component 36 may be a turbine nozzle comprising ceramic vane 40 attached to metal platforms 38 and 42. Platforms 38 and 42 may be formed of nickel base, iron base or cobalt base superalloys, titanium or titanium alloys. Vane 40 may be formed of a ceramic, silicon, silicon compound or a CMC. The ceramic may be silicon carbide, silicon nitride, silicon oxynitride, aluminum oxide and others known in the art. A CMC may be C/SiC, SiC/SiC, SiC/C, or mixtures thereof.

In an embodiment, SFF component 36 may be a rotating airfoil, bucket or blade. In another embodiment, SFF component 36 may be a combustor panel, combustor heat shield or combustor fuel nozzle.

As schematically shown in FIG. 1, interfaces 44 and 46 represent sharp transitions in monolithic composite component 36 from metal to ceramic interface 44, and ceramic to metal interface 46. If transition regions 44 and 46 are one layer thick, differences in coefficient of thermal expansion (CTE) across such a thin interface will cause failure of the interfaces in severe or even moderate thermal cycling during operation. It is the purpose of this invention to increase the structural integrity of interfaces such as those represented by metal to ceramic interfaces 44, and 46 in composite metal ceramic turbine components by forming functionally graded transition regions at the interfaces.

Functionally graded materials are a class of materials in which the material properties of a component vary with position throughout the component in a predetermined manner usually to minimize design constraints put on the component. In the example of a turbine nozzle illustrated schematically as component 36 in FIG. 1, it is advantageous to decrease the severity of the coefficient thermal expansion (CTE) mismatch across metal to ceramic interfaces 44 and 46. This may be achieved by creating functionally graded interface regions at interfaces 44 and 46 during the SFF formation of component 36.

In an embodiment, the functionally graded regions may be regions in which the compositions change from, for instance, 100 percent ceramic to 100 percent metal across a certain distance usually in a stepwise or continuous manner. According to the rule of mixtures the property P(X) of a composite material composed of a mixture of two types of materials such as, A and B with properties P_A and P_B and volume fractions V_A and V_B at a position X across a graded interface is: P_TOTAL(X)=P_AV_A(X)+P_BV_B(X)

As the transition region expands in size, the severity of mismatch decreases accordingly. Research suggests that a smooth transition in CTE across functionally graded zones 44 and 46 of the present invention requires at least 20 layers with each layer thickness ranging from 100 to 1000 microns to guarantee structural integrity under service conditions. Each intermediate zone or layer may have its own geometrical configuration defined in the build file and in the control system defining the specific mixture of metal to ceramic ratio. For example, the compositional transition of a functionally graded interface between a metal platform and a turbine vane of the invention may be:

0-30% span	100% platform metal
30-31% span	90% platform metal	10% ceramic
31-32% span	50% platform metal	50% ceramic
32-34% span	25% platform metal	75% ceramic
35-100% span	100% ceramic

Densification of each layer depends on the materials, laser power, scanning speed preheating temperatures and other input parameters necessary for consolidation. For instance, ceramic fusing issues may require sintering aids that would reduce the input power density for fusion for the composite powder

A perspective view of schematic monolithic composite metal ceramic turbine nozzle 36 of the invention shown in FIG. 1, with functionally graded material transitions is shown in FIG. 2 wherein like components are numbered accordingly. Nozzle 36 comprises metal platforms 38 and 42, ceramic vanes 40 and functionally graded transitions 45 and 47 replacing abrupt transitions 44 and 46.

An alternate method of producing SFF monolithic composite turbine components with functionally graded zones and interfaces is by powder based direct material deposition. Direct material deposition offers an additional degree of freedom over layer by layer SFF processes in that all three spacial dimensions (x,y,z) can be addressed simultaneously during a build. In direct material deposition, small amounts of material in solid, semi-molten or molten form are individually deposited to a SFF body according to a CAD model stored in memory in a direct material deposition system. Examples are laser engineered net shaping (LENS), shape deposit manufacturing (SDM), directed light manufacturing (DLF), direct laser melting (DLM), laser based additive manufacturing (LBAM), and others known in the art. A preferred technique for the present invention is direct laser melting.

Since an inert atmosphere is not generally required in some direct material deposition systems because of the direct material delivery process, larger parts may be produced and higher dimensional deposition paths and build structures may be achieved.

A schematic of SFF DLM direct material deposition process 50 is shown in FIG. 3. Process 50 includes manufacturing chamber 52, first powder 54, second powder 56, build platform 58, laser 60, scanning mirror 62, SFF deposited build structure 64 and molten region 66. During SFF direct material deposition process 50, first powder 54 is deposited at a controlled rate onto build platform 58 toward region 66 as indicated by arrows a. Second powder 56 is also deposited at a controlled rate onto build platform 58 toward region 66 as indicated by arrows b. Although two powder sources are indicated in FIG. 3, other powder sources may be employed. Material deposition in SFF direct material deposition process 50 is preferably coaxial with laser 60.

Laser 60 is activated according to a process schedule in a control system (not shown), of process 50 to fuse powders 54 and 56 in region 66 on solidified SFF structure 64. Solidification of direct deposited region 66 results in controlled incremental addition of first and second powder materials 54 and 56 to SFF structure 64. Build platform 58 is capable of three dimensional (x,y,z) translation allowing three dimensional SFF features to be formed region by region as well as layer by layer. A translation direction of platform 58 is indicated by arrow t.

A benefit of SFF direct material deposition is that free standing components can be formed with internal and external functionally graded structures to resist complex thermal mechanical induced internal stress fields during operation. Functional grading may be utilized to counter internal property gradients in single material types such as metal/metal, ceramic/ceramic, and CMC/CMC components in addition to metal/ceramic, metal/CMC, etc. composites. An example is shown in FIG. 4 in which schematic turbine nozzle 70, comprises metal platforms 72 and 74 and composite ceramic vane 76. In an embodiment, platforms 72 and 74 are nickel based superalloys and vane 76 is a composite ceramic structure with leading edge 76A and trailing edge 76B formed of two different ceramic materials. Leading edge 76A is a first ceramic with a lower CTE than trailing edge 76B which is a second ceramic. During operation, the thinner trailing edge heats up and cools down faster than the thicker leading edge. As a result, trailing edge 76B is subjected to higher longitudinal transient tensile and compressive stresses than leading edge 76A as taught by commonly owned US2009/0028697 to Shi et al. and incorporated herein in entirety as reference. Functionally graded regions 78 and 80 between metal platforms 72 and 74 and ceramic vane components 76A and 76B act to minimize internal stress due to CTE mismatch as discussed earlier. Functionally graded transition region 82 between ceramic trailing edge 76A and ceramic leading edge 76B acts to minimize internal stress due to longitudinal strain mismatch resulting from CTE mismatch between the trailing and leading edge components of composite vane 76.

It is to be understood that, while CTE is one property utilized in the invention to minimize or eliminate internal stress concentrations, other physical, mechanical, optical, chemical and other properties known in the art may be invoked in the SFF additive manufacturing fabrication of functionally graded features of turbine components of the present invention. In particular, in addition to CTE, heat capacity, thermal conductivity and Young's modulus are properties contributing to internal stress generation and turbine components during service.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments of the present invention.

A monolithic composite turbine component may include: at least one region of a first material; a second region of a second material; and at least one functionally graded transition region of the first material and the second material between a first region and a second region.

The component of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:

A first material may be a metal;

A second material may be a ceramic or a ceramic matrix composite;

The component may be formed by a solid freeform (SFF) additive manufacturing process;

The solid freeform (SFF) additive manufacturing process, may be selective laser melting (SLM);

The component may be a nozzle or a vane;

The component may be a rotating airfoil, bucket or blade;

The component may be a combustor panel, combustor heat shield, or combustor fuel nozzle;

The metal may be selected from the group consisting of nickel base, iron base, cobalt base superalloy, titanium and titanium alloy;

The ceramic may be selected from the group consisting of silicon carbide, silicon nitride, silicon oxynitride and aluminum oxide;

The ceramic matrix composite may be selected from the group consisting of SiC/SiC, SiC and SiC/Si.

A method of forming a monolithic composite turbine component containing separate regions of different materials may comprise: forming at least one first region of a first material; forming a graded transition region on the first region with composition of the transition region changing from a first material to a second material as the forming of the transition region progresses; and forming a second region of a second material on the transition region.

The method of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:

The first material may be a metal;

The second material may be a ceramic or ceramic matrix composite;

Forming may comprise solid free form (SFF) additive manufacturing;

Additive manufacturing may comprise direct laser melting;

The metal may be selected from a group consisting of nickel base, iron base, and cobalt base superalloy, titanium, and titanium alloys;

The ceramic may be selected from a group consisting of silicon carbide, silicon nitride, silicon oxynitride, and aluminum oxide;

The ceramic matrix composite may be selected from a group consisting of SiC/SiC, C/SiC, and SiC/Si;

The component may be a nozzle, vane, rotating airfoil, bucket, blade, combustor panel, combustor heat shield or combustor fuel nozzle.

While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.

Claims

1. A monolithic composite turbine component comprising:

at least one first region of a first material;

at least one second region of a second material; and

at least one functionally graded transition region of the first material and the second material between a first region and a second region of the component

2. The component of claim 1 wherein the first material is a metal.

3. The component of claim 1 wherein the second material is a ceramic or a ceramic matrix composite.

4. The component of claim 1 wherein the component is formed by a solid freeform (SFF) additive manufacturing process.

5. The component of claim 4 wherein the solid freeform (SFF) additive manufacturing process is selective laser melting (SLM).

6. The component of claim 1 wherein the component is a nozzle or a vane

7. The component of claim 1 wherein the component is a rotating airfoil, bucket, or blade.

8. The component of claim 1 wherein the component is a combustor panel, combustor heat shield, or combustor fuel nozzle.

9. The component of claim 2 wherein the metal is selected from the group consisting of nickel base, iron base, cobalt base superalloy, titanium, and titanium alloy.

10. The component of claim 3 wherein the ceramic is selected from the group consisting of silicon carbide, silicon nitride, silicon oxynitride, and aluminum oxide.

11. The component of claim 3 wherein the ceramic matrix composite is selected from the group consisting of SiC/SiC, C/SiC, and SiC/Si.

12. A method of forming a monolithic composite turbine component containing separate regions of different materials comprising:

forming at least one first region of a first material;

forming a graded transition region on the first region with composition of the transition region changing from a first material to a second material as the forming of the transition region progresses; and

forming a second region of a second material on the transition region.

13. The method of claim 12 wherein the first material is a metal.

14. The method of claim 12 wherein the second material is a ceramic or ceramic matrix composite.

15. The method of claim 12 wherein forming comprises solid freeform (SFF) additive manufacturing.

16. The method of claim 15 wherein additive manufacturing comprises direct laser melting.

17. The method of claim 13 where the metal is selected from a group consisting of nickel base, iron base and cobalt base superalloy, titanium and titanium alloy.

18. The method of claim 14 wherein the ceramic is selected from the group consisting of silicon carbide, silicon nitride, silicon oxynitride, and aluminum oxide.

19. The method of claim 14 wherein the ceramic matrix composite is selected from the group consisting of SiC/SiC, C/SiC, and SiC/Si.

20. The method of claim 12 wherein the component is a nozzle, vane, rotating airfoil, bucket, blade, combustor panel, combustor heat shield, or combustor fuel nozzle.