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Flow path assemblies and methods for assembling a flow path assembly of a gas turbine engine are provided. For example, a flow path assembly comprises a unitary outer wall including combustor and turbine portions that are integrally formed as a single unitary structure; a single piece, generally ann

Flow path assemblies and methods for assembling a flow path assembly of a gas turbine engine are provided. For example, a flow path assembly comprises a unitary outer wall including combustor and turbine portions that are integrally formed as a single unitary structure; a single piece, generally annular inner band; and a plurality of nozzle airfoils extending from the unitary outer wall to the inner band. Each nozzle airfoil interfaces with the inner band to position the inner band within the assembly. An exemplary assembly method comprises inserting an inner band into a flow path having a unitary outer wall as its outer boundary; inserting a plurality of nozzle airfoils into the flow path; and securing the nozzle airfoils with respect to the unitary outer wall. The inner band interfaces with an inner end of each nozzle airfoil to radially locate the inner band within the flow path.

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1. A flow path assembly of a gas turbine engine, the flow path assembly comprising: a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage and a second turbine stage of

1. A flow path assembly of a gas turbine engine, the flow path assembly comprising: a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage and a second turbine stage of a turbine section of the gas turbine engine, the combustor portion and the turbine portion being integrally formed as a single unitary structure;an inner band formed as a single piece, annular structure; anda plurality of nozzle airfoils extending from the unitary outer wall to the inner band,wherein the turbine portion comprises an outer band of a nozzle portion of the first turbine stage,a shroud of a blade portion of the first turbine stage,an outer band of a nozzle portion of the second turbine stage, anda shroud of a blade portion of the second turbine stage, andwherein each of the plurality of nozzle airfoils interfaces with the inner band to position the inner band within the flow path assembly. 2. The flow path assembly of claim 1, wherein the plurality of nozzle airfoils radially locate the inner band within the flow path assembly. 3. The flow path assembly of claim 1, wherein the plurality of nozzle airfoils form spokes to center the inner band within the unitary outer wall in a hub and spoke configuration. 4. The flow path assembly of claim 1, wherein each of the plurality of nozzle airfoils has an inner end and an outer end, and wherein the inner end of each of the plurality of nozzle airfoils is received in one of a plurality of pockets defined in the inner band. 5. The flow path assembly of claim 1, further comprising a backing ring radially inward of the inner band, wherein the backing ring interfaces with the inner band to position the backing ring within the flow path assembly. 6. The flow path assembly of claim 1, wherein each of the unitary outer wall, the inner band, and the plurality of nozzle airfoils are formed from a ceramic matrix composite material. 7. A flow path assembly of a gas turbine engine, the flow path assembly comprising: a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage and a second turbine stage of a turbine section of the gas turbine engine, the combustor portion and the turbine portion being integrally formed as a single unitary structure;an inner band formed as a single piece, annular structure;a backing ring formed as a single piece, annular structure; anda plurality of nozzle airfoils extending from the unitary outer wall to the inner band,wherein the turbine portion comprises an outer band of a nozzle portion of the first turbine stage,a shroud of a blade portion of the first turbine stage,an outer band of a nozzle portion of the second turbine stage, anda shroud of a blade portion of the second turbine stage, andwherein each of the plurality of nozzle airfoils attaches to the inner band to radially position and restrain the inner band within the flow path assembly, andwherein the backing ring attaches to the inner band adjacent a radially inner side of the inner band to position the backing ring within the flow path assembly. 8. The flow path assembly of claim 7, wherein the inner band is made from a different material than the backing ring. 9. The flow path assembly of claim 8, wherein the unitary outer wall, the inner band, and the plurality of nozzle airfoils are formed from a ceramic matrix composite material, and wherein the backing ring is formed from a metallic material. 10. The flow path assembly of claim 7, wherein the plurality of nozzle airfoils form spokes to center the inner band within the unitary outer wall in a hub and spoke configuration. 11. The flow path assembly of claim 7, wherein the backing ring attaches to the inner band using a straight spline joint. 12. The flow path assembly of claim 7, wherein the radially inner side of the inner band defines an opening and the backing ring defines an aperture, and wherein a pin is received in the opening and the aperture to attach the backing ring to the inner band. 13. The flow path assembly of claim 7, wherein the combustor portion of the unitary outer wall comprises an outer liner of the combustor, and wherein the turbine portion of the unitary outer wall comprises an outer band of a nozzle portion of the first turbine stage,a shroud of a blade portion of the first turbine stage,an outer band of a nozzle portion of a second turbine stage, anda shroud of a blade portion of the second turbine stage. 14. A method for assembling a flow path assembly of a gas turbine engine, the flow path assembly defining a flow path through a combustion section and at least a portion of a turbine section of the gas turbine engine, the flow path assembly comprising a unitary outer wall that defines an outer boundary of the flow path, the unitary outer wall including a combustor portion extending through the combustion section and a turbine portion extending through at least a first turbine stage and a section turbine stage of the turbine section, the turbine portion comprising an outer band of a nozzle portion of the first turbine stage, a shroud of a blade portion of the first turbine stage, an outer band of a nozzle portion of the second turbine stage, and a shroud of a blade portion of the second turbine stage, the combustor portion and the turbine portion being integrally formed as a single unitary structure, the method comprising: inserting an inner band into the flow path;inserting a plurality of nozzle airfoils into the flow path; andsecuring the plurality of nozzle airfoils with respect to the unitary outer wall,wherein the inner band interfaces with an inner end of each nozzle airfoil of the plurality of nozzle airfoils to radially locate the inner band within the flow path. 15. The method of claim 14, wherein the inner band is formed as a single piece, annular structure. 16. The method of claim 15, further comprising securing a backing ring adjacent an inner side of the inner band, wherein securing the backing ring to the inner band positions the backing ring within the flow path assembly, and wherein the backing ring is formed as a single piece, annular structure. 17. The method of claim 16, wherein the backing ring is secured to the inner band using a straight spline joint. 18. The method of claim 16, further comprising inserting a pin into an aperture in the backing ring and an opening in the inner band to secure the backing ring adjacent the inner side of the inner band.