Gas turbine engine

Patent metadata

Field	Value
Publication number	US-11542828-B2
Application number	US-202117200235-A
Country	US
Kind code	B2
Filing date	Mar 12, 2021
Priority date	Mar 30, 2020
Publication date	Jan 3, 2023
Grant date	Jan 3, 2023

Abstract

A gas turbine engine, includes: an engine core including a turbine, compressor, and shaft system connecting the turbine to the compressor, and forming a torque path therebetween. The shaft system is axially located by a thrust bearing located forward of the turbine, and the engine is configured, in the event of a shaft break which divides the shaft system into a front portion located by the thrust bearing and a rear portion unlocated by the thrust bearing, the rear portion is free to move axially rearwardly under a gas load. The engine further includes a shaft break detector having a forward speed sensor configured to measure a rotational speed of the front portion of the shaft system, and a rear microwave sensor configured to measure a rotational speed of the rear portion of the shaft system, wherein a shaft break can be detected based on differences in the measured speeds.

First claim

We claim:

1. A gas turbine engine comprising:
an engine core comprising a turbine, a compressor, and a shaft system connecting the turbine to the compressor, and forming a torque path therebetween wherein:
the shaft system is axially located by a thrust bearing located forward of the turbine, the gas turbine engine being configured such that, in an event of a shaft break which divides the shaft system into a front portion located by the thrust bearing and a rear portion not located by the thrust bearing, the rear portion is free to move axially rearwardly under a gas load,
the gas turbine engine further comprises a shaft break detector having a forward speed sensor configured to measure a rotational speed of the front portion of the shaft system, and a rear microwave sensor configured to measure a rotational speed and an axial movement of the rear portion of the shaft system, and
a shaft break can be detected based on differences in the rotational speed measured by the forward speed sensor and the rotational speed measured by the rear microwave sensor, and based on the axial movement measured by the rear microwave sensor.

2. The gas turbine engine of claim 1 , wherein the shaft system is supported by one or more non-thrust bearings rearwards of the thrust bearing.

3. The gas turbine engine of claim 1 , wherein the thrust bearing is positioned adjacent the compressor.

4. The gas turbine engine of claim 1 , wherein the turbine is a low pressure turbine, and the compressor is a low pressure compressor.

5. The gas turbine engine of claim 4 , wherein the low pressure compressor is a fan.

6. The gas turbine engine of claim 5 , wherein the shaft system comprises a main shaft component which extends forward from the turbine, and a fan shaft component which joins to a front end of the main shaft component and connects the fan to the shaft system, the thrust bearing axially locating the shaft system at the fan shaft component.

7. The gas turbine engine of claim 1 , further comprising one or more friction decelerators which, in the event of the shaft break, engage with the turbine to apply frictional braking thereto.

8. The gas turbine engine of claim 1 , wherein the axial movement of the rear portion of the shaft system produces tangling of rotors and stators of the turbine.

9. The gas turbine engine of claim 1 , wherein the rear microwave sensor is positioned adjacent the turbine to measure the rotational speed of the turbine.

10. The gas turbine engine of claim 9 , wherein the turbine is a multi-stage turbine, and the rear microwave sensor is positioned adjacent a stage of the turbine which is a first turbine stage on the torque path from the turbine to measure the rotational speed of the first turbine stage.

11. The gas turbine engine of claim 1 , wherein the forward speed sensor is a further microwave sensor.

12. The gas turbine engine of claim 11 , wherein the further microwave sensor is positioned adjacent a fan to measure a rotational speed of the fan.

13. The gas turbine engine of claim 1 , further comprising an electronic engine controller which receives the rotational speed measured by the forward speed sensor and the rotational speed measured by the rear microwave sensor, and is configured to detect a shaft break based on differences in the rotational speed measured by the forward speed sensor and the rotational speed measured by the rear microwave sensor.

14. The gas turbine engine of claim 13 , wherein the electronic engine controller is further configured to restrict a flow of fuel to the gas turbine engine when a shaft break is detected.

15. An aircraft, including the gas turbine engine of claim 1 .

Assignees

• Rolls Royce Plc
• Rolls Royce Deutschland Ltd & Co Kg

Inventors

• Calderon Jorge
• Brown David

Classifications

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F01D21/003

  Arrangements for testing or measuring (for measuring vibrations G01H) · CPC title

• F05D2260/82

  Forecasts · CPC title

• F05D2260/80

  Diagnostics · CPC title

• F05D2270/304

  Spool rotational speed · CPC title