A gas turbine includes a rotor that is rotatable about an axis, a casing configured to cover the rotor in a circumferential direction and having an annular space therein, a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing, a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas, a turbine driven by the combustion gas, and an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

A combustion tube for a gas turbine including an outlet section having a cross-section of an annular sector-shape. The outlet section includes an outer wall forming an outer peripheral boundary of the annular sector-shape, an inner wall forming an inner peripheral boundary of the annular sector-shape, and a pair of side walls forming boundaries on both sides of the annular sector-shape in a circumferential direction, respectively. The outer wall extends obliquely with respect to the inner wall such that a height of the annular sector-shape decreases toward an outlet opening of the combustion tube. A first side wall extends obliquely with respect to a second side such that a perimeter of the annular sector-shape increases toward the outlet opening of the combustion tube. An inclination angle θ.sub.1 of the first side wall with respect to the second side wall satisfies 0<θ.sub.1≤56.

A system is provided with a fuel sweep system configured to couple to a flow sleeve of a combustor along a first fuel conduit. The flow sleeve is configured to be disposed about a liner of the combustor, and the first fuel conduit is configured to extend along the flow sleeve in a compressor discharge chamber disposed about the flow sleeve. The fuel sweep system includes a first fuel sweep louver adjacent a first fuel sweep opening defined through the flow sleeve.

A combustor adapted for use in a gas turbine engine includes a combustor shell comprising metallic materials. The combustor shell is formed to define an internal space. The combustor further includes a heat shield mounted to an axially aft surface of the combustor shell within the internal space and a combustor liner arranged to extend along inner surfaces of the combustor shell within the internal space. The combustor liner cooperates with the heat shield to define a combustor chamber.

The invention relates to a turbine engine combustion assembly (20), which includes: an annular flame tube (21) including a front wall (23), a rear wall (24) and a bottom (22) arranged facing an engine shaft (30); an injection wheel (41) rotated by said engine shaft (30), partially projecting into the bottom (22) of the flame tube (21) and configured such as to spray fuel into the flame tube by centrifugation; and at least one injector (35), capable of depositing a film of fuel on said injection wheel (41), said combustion assembly being characterised in that said injector (35) is arranged so as to pass through said upstream area of the front wall (23) or the rear wall (24) of the flame tube (21), so that the injection opening thereof (37) opens into said tube (21), opposite the portion (43) of said injection wheel (41) which is located inside said flame tube (21).

The invention relates to a turbine engine combustion assembly (20), which includes: an annular flame tube (21) including a front wall (23), a rear wall (24) and a bottom (22) arranged facing an engine shaft (30); an injection wheel (41) rotated by said engine shaft (30), partially projecting into the bottom (22) of the flame tube (21) and configured such as to spray fuel into the flame tube by centrifugation; and at least one injector (35), capable of depositing a film of fuel on said injection wheel (41), said combustion assembly being characterised in that said injector (35) is arranged so as to pass through said upstream area of the front wall (23) or the rear wall (24) of the flame tube (21), so that the injection opening thereof (37) opens into said tube (21), opposite the portion (43) of said injection wheel (41) which is located inside said flame tube (21).

The present disclosure generally relates to a system with a gas turbine engine including a first combustor and a second combustor. The first combustor includes a first end cover with a first geometry and the second combustor includes a second end cover with a second geometry. The first geometry has one or more geometric differences relative to the second geometry.

In one embodiment, a turbine system includes a combustion system comprising a plurality of combustion cans, a number of sensors, each of the number of sensors coupled to a respective combustion can of the number of combustion cans, and a controller. The controller includes a memory storing one or more processor-executable routines and a processor configured to access and execute the one or more routines encoded by the memory. The one or more routines, when executed cause the processor to receive one or more signals from the number of sensors, convert the one or more signals to audio output, and output the converted audio output via one or more audio output devices.