This disclosure claims the benefit of UK Patent Application No. GB 1901514.8, filed on 4 Feb. 2019, which is hereby incorporated herein in its entirety.

The disclosure relates to a gas turbine engine, and particularly a gas turbine engine for use in an aircraft.

Modern gas turbine engines generally have up to three compressor-turbine groups, also referred to as spools, connected by respective concentric shafts. These compressor-turbine groups are responsible for the compression and expansion of air passing through the engine.

For example, the Rolls-Royce Trent 1000 aerospace gas turbine engine has three compressor-turbine groups: the low pressure compressor-turbine group, the intermediate pressure compressor-turbine group, and the high pressure compressor-turbine group. Each of these has a corresponding compressor and turbine. In the case of the low pressure compressor-turbine group, the low pressure compressor is generally referred to as the fan.

It is possible, during operation of the gas turbine engine, for a shaft connecting a compressor and turbine pair to mechanically fail. The breakage in the shaft leads to an instantaneous decoupling between the turbine and its respective compressor. As a result the compressor may decelerate rapidly, as it is no longer driven by the turbine, and the turbine may accelerate rapidly, as it is no longer driving the compressor.

The rapid acceleration of the turbine is particularly concerning, as over-speed events can lead to disintegration of the turbine, including possibly bursting of the turbine disc, and further damage to the gas turbine engine.

It is important then to rapidly take the appropriate action when a shaft break occurs.

Accordingly, in a first aspect, the present disclosure provides a method of controlling a gas turbine engine having a shaft connecting a compressor to a turbine and further having a reheat system, the method including the steps of: operating the engine using the reheat system to provide a mass flow rate of reheat fuel into a gas flow of the gas turbine engine downstream of an exit of the turbine; detecting a shaft break event in the shaft and in response to this detection, maintaining the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit, whereby the maintained mass flow rate of reheat fuel raises a back pressure downstream of the turbine and thereby reduces a rotational speed of the turbine.

Advantageously, the method according to the first aspect can decrease a terminal speed of the turbine after a shaft break event by reducing the amount of energy available to drive the turbine, meaning that, in the event of a disc burst, the debris would have a lower kinetic energy. Alternatively, the reduced energy can be parlayed into a reduced size and/or weight of the disc. This can also lower the debris' kinetic energy in the event of a disc burst.

The maintenance of the mass flow rate of the reheat fuel can be implemented in practice by (A) providing the same mass flow rate into the gas flow after the shaft break event as before the shaft break event, or (B) actually increasing the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit after the shaft break event. That is, the increased mass flow rate of option (B) can be regarded as being made up of two portions: a first portion is the maintained mass flow rate, and the second portion is an additional mass flow rate. The additional mass flow rate can further raise the back pressure downstream of the turbine and thereby further reduce the rotational speed of the turbine.

The method of the first aspect may have any one, or any combination, of the following optional features.

The gas flow into which the mass flow rate of reheat fuel is provided is typically an exhaust of the turbine.

The step of maintaining the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit may be performed by maintaining a fuel schedule for the reheat system in a pre-shaft break event configuration. Advantageously, this facilitates implementation of the method of the first aspect by a computer-based controller, such as an electronic engine controller, rather than by a mechanical device which might incur a weight penalty on the resulting gas turbine engine.

The raised back pressure acts to oppose the gas flow in its path from compressor to turbine, and it is this which reduces a rotational speed of the turbine. Preferably, raising the back pressure downstream of the turbine moves the working line of the compressor towards a surge region. This can encourage the compressor to surge or stall, resulting in a change in the direction of airflow through the gas turbine engine which can further reduce the rotational speed of the turbine.

In a second aspect, the present disclosure provides a gas turbine engine, comprising: a shaft connecting a compressor to a turbine; a reheat system; and an electronic engine controller wherein the electronic engine controller is configured to: operate the engine using the reheat system to provide a mass flow rate of reheat fuel into a gas flow of the gas turbine engine downstream of an exit of the turbine; detect a shaft break event in the shaft; and in response to this detection, maintain the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit, whereby the maintained mass flow rate of reheat fuel raises a back pressure downstream of the turbine and thereby reduces a rotation speed of the turbine.

The engine of the second aspect may have any one, or any combination, of the following optional features.

The gas flow into which the mass flow rate of reheat fuel is provided is typically an exhaust of the turbine, or may be a bypass flow.

Maintaining the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit may be performed by the electronic engine controller maintaining a fuel schedule for the reheat system in a pre-shaft break event configuration.

Raising the back pressure downstream of the turbine may move the working line of the compressor towards a surge region. This can encourage the compressor to surge or stall, resulting in a change in the direction of airflow through the gas turbine engine which can further reduce the rotational speed of the turbine.

In a third aspect, the present disclosure provides an aircraft including the gas turbine engine of the second aspect.

Further aspects of the present invention provide: a computer program comprising code which, when run on a computer, causes the computer to perform the method of the first aspect; a computer readable medium storing a computer program comprising code which, when run on a computer, causes the computer to perform the method of the first aspect; and a computer system programmed to perform the method of the first aspect.

