An aspect includes a method for motoring control for multiple engines of an aircraft is provided. A controller can determine a motoring time of a first engine starting system to cool a first engine. The controller can compare the motoring time of the first engine starting system with a motoring time of one or more other engine starting systems of one or more other engines of the aircraft. The motoring time of the first engine starting system can be controlled relative to a tolerance of the motoring time of the one or more other engine starting systems by adjusting the motoring time of the first engine starting system relative to the one or more other engine starting systems in a motoring sequence based on comparing the motoring time of the first engine starting system with the motoring time of the one or more other engine starting systems.

A method for operating a gas turbine engine comprises providing fuel flow and compressed airflow to a combustor with a fuel-to-air ratio, the compressed airflow being from a compressed air source; detecting at least one parameter indicative of the fuel-to-air ratio being below a predetermined value; and bleeding compressed air from the compressed air source when the at least one parameter indicative of the fuel-to-air ratio is below the predetermined value to increase the fuel-to-air ratio to at least the predetermined value.

An object of the present invention is to provide a method and a system for implementing the method so as to alleviate the disadvantages of a reciprocating combustion engine and gas turbine when generating power. The invention is based on the idea of arranging a combustion chamber (10) outside a turbine (22) and providing compressed air from serially connected compressors to an air chamber in which the air is heated and then exhausted to the combustion chamber in order to carry out a combustion process supplemented with high pressure steam pulses.

A fuel system can include a total flow line configured to receive a total flow and a primary flow line connected to the total flow line. The primary flow line can be in fluid communication with one or more primary fuel nozzles of a nozzle assembly. The fuel system can include a secondary flow line connected to the total flow line in parallel with the primary flow line, the secondary flow line in fluid communication with a plurality of secondary flow nozzles of the nozzle assembly. The fuel system can include a flow split system configured to control a flow split between a primary flow of the primary flow line and a secondary flow of the secondary flow line.

A combustion apparatus can avoid instability in combustion by controlling a pressure ratio of fuel mixed with air. The combustion apparatus includes a casing; a pilot nozzle disposed at the center of the casing and supplied with fuel by a pilot fuel supply pipe; and a plurality of main nozzles arranged around the pilot nozzle and supplied with fuel by a main fuel supply pipe, each main nozzle including a pair of parallel fuel channels each extending to a respective fuel spray position within the main nozzle. A gas turbine adopting the combustion apparatus includes a plurality of combustors, each combustor including the casing, pilot nozzle, and plurality of main nozzles, with each main nozzle having first and second fuel channels respectively extending to a fuel spray position of the corresponding fuel channel. A pilot manifold connects the respective pilot nozzles, and a main manifold connects the respective main nozzles.

A gas turbine may include a rotatable shaft, a compressor disposed about the rotatable shaft and configured to output compressed air, and a combustor disposed about the rotatable shaft. The combustor may be configured to receive the compressed air and output high temperature compressed gas. The gas turbine may further include a power turbine disposed about the rotatable shaft and configured to receive the high temperature compressed gas, and a first liner defining a plurality of holes and disposed around the combustor. The power turbine may be configured to expand the high temperature compressed gas and rotate the rotatable shaft. The first liner may have a first end and a longitudinally opposite second end. The first end may be coupled to an inner surface of the casing at or adjacent an upstream end of the combustor and the second end may be substantially free from any connection with the casing.

A non-transitory computer readable medium with instructions stored thereon, the instructions executable by one or more processors for selecting infrequent or frequent autotuning of a combustor; and determining the health of a combustor. Also, a method of monitoring a combustor within a gas turbine engine system, comprising providing a gas turbine engine system, wherein the gas turbine engine includes an autotuning system; selecting infrequent or frequent autotuning of the combustor; and determining the health of the combustor; wherein said determining the health of a combustor comprises receiving real-time fuel gas temperature data from at least one thermocouple.

A fuel control device includes a stem fuel valve opening degree determination unit, a branch line flow rate determination unit, and a correction value determination unit. The stem fuel valve opening degree determination unit is configured to determine an opening degree of a flow rate adjustment valve of a stem fuel supply line. The branch line flow rate determination unit is configured to determine the opening degree of each flow rate adjustment valve of the branch line based on an operation situation of the gas turbine. The correction value determination unit is configured to determine a correction value of the opening degree of each flow rate adjustment valve of the branch line based on a value of a pressure difference between a fuel pressure upstream of each of nozzles connected to the branch lines respectively and a corrected fuel pressure for the fuel pressure at an outlet.

Systems and methods for indirect monitoring of combustor dynamics in a gas turbine engine include collecting vibration data acquired by a vibration sensor, which is mounted proximate to an operational component positioned relative to a combustor of a gas turbine engine. The vibration data can be transformed into a frequency domain representation at periodic intervals. The relative signal strength of the vibration data can be determined over the one or more identified frequency bands. The relative signal strength can be adjusted with hysteresis at each of the one or more identified frequency bands. Occurrence of a combustor dynamics event at one or more specific resonant frequencies can be determined based at least in part on evaluation of the adjusted relative signal strength relative to one or more event detection threshold levels.

Herein provided are methods and systems for reducing an acoustic signature of a gas turbine engine. An acceleration command for the engine is received. In response to receiving the acceleration command: a fuel flow to the engine is increased for a first predetermined time period; subsequent to the first predetermined time period, the fuel flow to the engine is reduced for a second predetermined time period; and subsequent to the second predetermined time period, the fuel flow to the engine is increased for a third time period.