Translating inlet for adjusting airflow distortion in gas turbine engine

1. A core engine for a gas turbine engine, comprising:

a compressor section, a combustion section, and a turbine section defining at least in part an engine airflow path for the core engine;

an inner flowpath surface positioned at least partially within the compressor section and defining at least in part the engine airflow path;

a core casing at least partially enclosing the compressor section and defining a forward end, the core casing comprising a moving inlet assembly at the forward end, the moving inlet assembly and inner flowpath surface together defining an inlet to the compressor section, the moving inlet assembly moveable between a first position defining a first inlet area at the inlet and a second position defining a second inlet area at the inlet, the first inlet area being greater than the second inlet area,

wherein the core engine defines a centerline axis,

wherein the moving inlet assembly is movable between the first position and the second position in a substantially uniform manner about the centerline axis of the core engine; and

an instrumented guide vane comprising one or more pressure sensor devices located at least partially in the engine airflow path for obtaining one or more measurements associated with airflow distortion,

wherein the moving inlet assembly is configured to be controlled based at least in part on signals from the one or more pressure sensor devices,

wherein the moving inlet assembly defines a leading edge of the core casing,

wherein the leading edge defines a first circumference in the first position,

wherein the leading edge defines a second circumference in the second position, and

wherein the second circumference is less than the first circumference.

2. The core engine of claim 1 , wherein the moving inlet assembly allows a first engine airflow at a first airflow speed into the engine airflow path in the first position, wherein the moving inlet assembly allows a second engine airflow at a second airflow speed into the engine airflow path in the second position, wherein at least one of the first engine airflow is greater than the second engine airflow or the first engine airflow speed is less than the second engine airflow speed.

3. The core engine of claim 1 , wherein the moving inlet assembly is configured to be controlled based on the airflow distortion in the engine airflow path, and wherein the airflow distortion is an inlet airflow distortion.

4. The core engine of claim 1 , wherein the compressor section includes a first compressor and a second compressor, wherein the moving inlet assembly is configured to be controlled based on the airflow distortion in the engine airflow path, and wherein the airflow distortion is an airflow mismatch between the first and second compressors.

5. The core engine of claim 1 , wherein the moving inlet assembly is additionally movable to an intermediate position, wherein the intermediate position defines an intermediate inlet area, wherein the intermediate inlet area is less than the first inlet area and greater than the second inlet area.

6. The core engine of claim 5 , wherein the moving inlet assembly is movable between the first, second, and intermediate positions based on signals from a controller to adjust the airflow distortion in the engine airflow path.

7. A method for adjusting airflow distortion in a gas turbine engine on an aircraft, the gas turbine engine comprising a compressor section, a combustion section, and a turbine section in series flow, the compressor section, combustion section, and turbine section defining at least in part an engine airflow path, the gas turbine engine further comprising an inner flow path surface positioned at least partially within the compressor section and defining at least in part the engine airflow path, the gas turbine engine further comprising a core casing at least partially enclosing the compressor section and defining a forward end, the method comprising:

determining, by one or more control devices, an airflow distortion condition associated with the engine airflow path, the airflow distortion condition determined at least in part from one or more measurements obtained by an instrumented guide vane; and

controlling, by the one or more control devices, a moving inlet assembly to adjust the airflow distortion condition of the gas turbine engine having been determined at least in part from the one or more measurements obtained by the instrumented guide vane, wherein the core casing comprises the moving inlet assembly at the forward end, the moving inlet assembly and inner flow path surface together defining an inlet to the compressor section, the moving inlet assembly moveable between a first position defining a first inlet area and a second position defining a second inlet area, the first inlet area being greater than the second inlet area,

wherein the core engine defines a centerline axis,

wherein the moving inlet assembly is movable between the first position and the second position in a substantially uniform manner about the centerline axis of the core engine,

wherein the moving inlet assembly defines a leading edge of the core casing,

wherein the leading edge defines a first circumference in the first position,

wherein the leading edge defines a second circumference in the second position, and

wherein the second circumference is less than the first circumference.

8. The method of claim 7 , wherein the moving inlet assembly is additionally movable to an intermediate position, wherein the intermediate position defines an intermediate inlet area, wherein the intermediate inlet area is less than the first inlet area and greater than the second inlet area.

9. The method of claim 8 , wherein the moving inlet assembly allows a first engine airflow into the engine airflow path in the first position, wherein the moving inlet assembly allows a second engine airflow into the engine airflow path in the second position, wherein the moving inlet assembly allows an intermediate engine airflow into the engine airflow path in the intermediate position, wherein the intermediate engine airflow is less than the first engine airflow and greater than the second engine airflow.

10. The method of claim 8 , wherein controlling the moving inlet assembly comprises moving the moving inlet assembly between the first, second, and intermediate positions.

11. The method of claim 9 , wherein controlling the moving inlet assembly to adjust the airflow distortion condition comprises controlling the moving inlet assembly to reduce the airflow distortion condition.

12. The method of claim 7 , wherein determining the airflow distortion condition associated with the engine airflow path comprises obtaining the one or more measurements using one or more pressure sensor devices, wherein controlling the moving inlet assembly to adjust the airflow distortion condition comprises controlling the moving inlet assembly based at least in part on the one or more measurements obtained using the one or more pressure sensor devices.

13. A gas turbine engine system for an aircraft comprising:

a compressor section, a combustion section, and a turbine section defining at least in part an engine airflow path for a core engine;

an inner flowpath surface positioned at least partially within the compressor section and defining at least in part the engine airflow path;

a core casing at least partially enclosing the compressor section and defining a forward end, the core casing comprising a moving inlet assembly at the forward end, the moving inlet assembly and inner flowpath surface together defining an inlet to the compressor section, the moving inlet assembly moveable between a first position defining a first inlet area at the inlet and a second position defining a second inlet area at the inlet, the first inlet area being greater than the second inlet area, wherein the core engine defines a centerline axis, wherein the moving inlet assembly is movable between the first position and the second position in a substantially uniform manner about the centerline axis of the core engine; and

a controller operably connected to the moving inlet assembly, the controller comprising one or more processors and one or more memory devices located on the aircraft, the one or more memory devices storing instructions that when executed by the one or more processors cause the one or more processors to perform operations, the operations comprising:

determine an airflow distortion condition within the engine airflow path; and

control the moving inlet assembly to adjust an airflow through the engine airflow path to adjust the airflow distortion condition,

wherein the moving inlet assembly defines a leading edge of the core casing,

wherein the leading edge defines a first circumference in the first position,

wherein the leading edge defines a second circumference in the second position, and

wherein the second circumference is less than the first circumference.

14. The gas turbine engine system of claim 13 , wherein the core engine comprises one or more pressure sensor devices located at least partially in the engine airflow path configured to obtain measurements to determine the airflow distortion condition, wherein the moving inlet assembly is controlled by the controller based at least in part on the measurements obtained by the one or more pressure sensor devices.

15. The gas turbine engine system of claim 13 , wherein the moving inlet assembly is additionally movable to an intermediate position, wherein the intermediate position defines an intermediate inlet area, wherein the intermediate inlet area is less than the first inlet area and greater than the second inlet area.

16. The gas turbine engine system of claim 13 , wherein controlling the moving inlet assembly comprises moving the moving inlet assembly between the first position, the second position, and intermediate positions.