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A stacked laminate component for a turbine engine that may be used as a replacement for one or more metal components is provided. The stacked laminate component can have a body formed by a process of stacking and laminating layers to define a radially inner surface along the hot gas path. The layers

A stacked laminate component for a turbine engine that may be used as a replacement for one or more metal components is provided. The stacked laminate component can have a body formed by a process of stacking and laminating layers to define a radially inner surface along the hot gas path. The layers can be substantially orthogonal to the radially inner surface. The layers can be at least a first layer of a first material and a second layer of a second material. At least the first material is a ceramic matrix composite. The second material can have a higher thermal conductivity or a higher creep strength than the first material.

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1. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers with a radially inner surface along the radially outer boundary of the hot gas path and a radially outer surface, the body being formed at lea

1. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers with a radially inner surface along the radially outer boundary of the hot gas path and a radially outer surface, the body being formed at least in part by a plurality of layers;wherein each layer has a height dimension extending orthogonal to the radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to the radially inner surface; andwherein the plurality of layers being at least a first layer formed from a first material and a second layer formed from a second material, wherein the first material is a ceramic matrix composite;wherein the second material has a higher resistance to creep deformation than the first material. 2. The component of claim 1, wherein the second material has a higher thermal conductivity than the first material. 3. The component of claim 1, wherein the second layer is recessed from the first layer along the radially inner surface. 4. The component of claim 1, wherein the first layer is recessed from the second layer along the radially outer surface. 5. The component of claim 1, further comprising a coating on the radially inner surface, wherein the first layer is recessed from the second layer along the radially inner surface and wherein the second layer extends into the coating. 6. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers with a radially inner surface along the radially outer boundary of the hot gas path and a radially outer surface, the body being formed at least in part by a plurality of layers;wherein each layer has a height dimension extending orthogonal to the radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to the radially inner surface; andwherein the plurality of layers being at least a first layer formed from a first material and a second layer formed from a second material, wherein the first material is a ceramic matrix composite;wherein the second material is a ceramic matrix composite. 7. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers with a radially inner surface along the radially outer boundary of the hot gas path and a radially outer surface, the body being formed at least in part by a plurality of layers;wherein each layer has a height dimension extending orthogonal to the radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to the radially inner surface; andwherein the plurality of layers being at least a first layer formed from a first material and a second layer formed from a second material, wherein the first material is a ceramic matrix composite;wherein the first and second layers each comprise a plurality of first and second layers that are positioned in an alternating pattern along the body. 8. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers with a radially inner surface along the hot gas path and a radially outer surface, the body being formed at least in part by a plurality of layers;wherein each layer has a height dimension extending orthogonal to the radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to the radially inner surface;wherein the plurality of layers being at least a first layer formed from a first material and a second layer formed from a second material, wherein the first material is a ceramic matrix composite; anda coating on the radially inner surface, wherein the first layer extends into the coating. 9. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers with a radially inner surface along the hot gas path and a radially outer surface, the body being formed at least in part by a plurality of layers;wherein each layer has a height dimension extending orthogonal to the radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to the radially inner surface;wherein the plurality of layers being at least a first layer formed from a first material and a second layer formed from a second material, wherein the first material is a ceramic matrix composite; andan overwrap that provides a compressive preload on the body. 10. The component of claim 9, wherein the overwrap is a ceramic matrix composite. 11. The component of claim 9, wherein the overwrap is formed from one of a first overwrap material having a higher coefficient of thermal expansion and a higher secondary processing shrinkage than the plurality of layers, a second overwrap material having a lower coefficient of thermal expansion and a higher secondary processing shrinkage than the plurality of layers or a third overwrap material having a substantially similar coefficient of thermal expansion and a higher secondary processing shrinkage than the plurality of layers. 12. A gas turbine component exposed to a hot gas path of a gas turbine, the component comprising: a body formed by a process of stacking and laminating layers to define a radially inner surface along the radially outer boundary of the hot gas path;wherein each layer has a height dimension extending orthogonal to the radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to the radially inner surface;wherein the layers are formed from at least a first layer of a first material and a second layer of a second material;wherein at least the first material is a ceramic matrix composite; andwherein the second material has at least one of a higher thermal conductivity or a higher creep strength than the first material. 13. The component of claim 12, wherein the second material is a ceramic matrix composite, a sapphire fiber felt or a mullite whisker felt. 14. The component of claim 12, wherein the first and second layers each comprise a plurality of first and second layers that are positioned in an alternating pattern along at least a portion of the body. 15. The component of claim 12, further comprising a coating on the radially inner surface, wherein the second material has a higher thermal conductivity than the first material, wherein the body has a radially outer surface, wherein the second layer is recessed from the first layer along the radially inner surface, wherein the first layer extends into the coating, and wherein the first layer is recessed from the second layer along the radially outer surface. 16. The component of claim 12, further comprising a coating on the radially inner surface, wherein the second material has a higher creep strength than the first material, wherein the first layer is recessed from the second layer along the radially inner surface and wherein the second layer extends into the coating. 17. The component of claim 12, further comprising a ceramic matrix composite overwrap that provides a compressive preload on the body. 18. A method of manufacturing a gas turbine component comprising: providing at least a first material and a second material, the first material being a ceramic matrix composite, the second material having at least one of a higher thermal conductivity or a higher creep strength than the first material;stacking and laminating the first and second materials to define a body comprising layers, the first and second materials being arranged in alternating layers along at least a portion of the body;cutting the body, wherein each layer has a height dimension extending orthogonal to a radially inner surface that is greater than a width dimension extending parallel to the radially inner surface, thereby causing each of the layers to be positioned substantially orthogonal to a radially inner surface of the body; andapplying an overwrap that provides a compressive preload on the body. 19. The method of claim 18, wherein the component is a ring seal segment or a combustor heat shield.