A combustion chamber of an aircraft turbomachine having a main axis includes: a body of revolution coaxial with the main axis having a plurality of combustion tubes extending mainly in the direction of the main axis and being distributed in a ring about the main axis; a first perforated rotary disc mounted at a first axial end of the body and rotatable about the main axis to selectively open or close a first end of each of the combustion tubes; and a second perforated rotary disc mounted at a second axial end of the body and rotatable about the main axis to selectively open or close a second end of each of the combustion tubes. The body includes a plurality of cooling segments extending mainly in the direction of the main axis, which are distributed in a ring about the main axis and around the combustion tubes.

A combustion chamber of an aircraft turbomachine having a main axis includes: a body of revolution coaxial with the main axis having a plurality of combustion tubes extending mainly in the direction of the main axis and being distributed in a ring about the main axis; a first perforated rotary disc mounted at a first axial end of the body and rotatable about the main axis to selectively open or close a first end of each of the combustion tubes; and a second perforated rotary disc mounted at a second axial end of the body and rotatable about the main axis to selectively open or close a second end of each of the combustion tubes. The body includes a plurality of cooling segments extending mainly in the direction of the main axis, which are distributed in a ring about the main axis and around the combustion tubes.

The method and system are for implementing the method so as to alleviate the disadvantages of a reciprocating combustion engine and gas turbine when generating power. A combustion chamber is arranged outside a turbine and provides compressed air from a turbocharger powered with a heat source in order to carry out a combustion process supplemented with high pressure steam pulses.

The method and system are for implementing the method so as to alleviate the disadvantages of a reciprocating combustion engine and gas turbine when generating power. A combustion chamber is arranged outside a turbine and provides compressed air from a turbocharger powered with a heat source in order to carry out a combustion process supplemented with high pressure steam pulses.

A pulse detonation combustion system includes: an inlet pipe; an intake cone disposed in the inlet pipe, having an end provided with a pneumatic valve, and including an atomizing air transfer tube, a fuel transfer tube, and a conical swirl nozzle connected to the atomizing air transfer tube and the fuel transfer tube; an atomizing air intake tube connected with the atomizing air transfer tube; a fuel supply tube connected with the fuel transfer tube; a pulse detonation combustion chamber located downstream of and communicated to the inlet pipe, and provided with a spark plug mounting seat for mounting a spark plug; a gas energy distribution adjustment device located downstream of and communicated to the pulse detonation combustion chamber; and a transition section located downstream of and communicated to the gas energy distribution adjustment device.

The method is for controlling a deflagration combustion process in a pistonless combustor. The method includes scavenging combustion products of the previous cycle, introducing air into the combustor thereby initiating a flow pattern having a first flow component within the combustor. Air is introduced into the pistonless combustor in a nonparallel angle in relation to the previous air input and thereby creating a second flow component to the flow pattern for increasing speed of combustion propagation. Fuel mixed into the air is introduced for creating a fuel-air mixture flowing within the flow pattern, and igniting the fuelair mixture within the pistonless combustor thereby increasing pressure within the pistonless combustor.

The method is for controlling a deflagration combustion process in a pistonless combustor. The method includes scavenging combustion products of the previous cycle, introducing air into the combustor thereby initiating a flow pattern having a first flow component within the combustor. Air is introduced into the pistonless combustor in a nonparallel angle in relation to the previous air input and thereby creating a second flow component to the flow pattern for increasing speed of combustion propagation. Fuel mixed into the air is introduced for creating a fuel-air mixture flowing within the flow pattern, and igniting the fuelair mixture within the pistonless combustor thereby increasing pressure within the pistonless combustor.

A module (4) for an aircraft turbomachine comprises an assembly of constant-volume combustion chambers, and including a first sub-assembly of first chambers succeeding each other along a given sense (76) and forming series of chambers (S1), and within each series (S1), a first ignition chamber (C1.1) located at one of both circumferential ends of the series is defined, the first ignition chamber (C1.1) being connected to the first directly consecutive chamber (C1.2) along the given sense (76) so as to supply the same with exhaust gases, and so forth up to the first chamber (C1.3) located at the other circumferential end of the series. In addition, a control device (46) is configured such that for all the first ignition chambers (C1.1), diametrically opposite two by two, the combustion cycles are simultaneously initiated. Finally, a second sub-assembly comprising second combustion chambers (C2.1-C2.3) is also provided.

A module (4) for an aircraft turbomachine comprises an assembly of constant-volume combustion chambers, and including a first sub-assembly of first chambers succeeding each other along a given sense (76) and forming series of chambers (S1), and within each series (S1), a first ignition chamber (C1.1) located at one of both circumferential ends of the series is defined, the first ignition chamber (C1.1) being connected to the first directly consecutive chamber (C1.2) along the given sense (76) so as to supply the same with exhaust gases, and so forth up to the first chamber (C1.3) located at the other circumferential end of the series. In addition, a control device (46) is configured such that for all the first ignition chambers (C1.1), diametrically opposite two by two, the combustion cycles are simultaneously initiated. Finally, a second sub-assembly comprising second combustion chambers (C2.1-C2.3) is also provided.

A turbine engine system utilizes one or more augmented combustion modules to produce an exhaust that is fed into the turbine portion of the engine and wherein power is produced by the augmented combustion module for use to drive the main shaft and/or for auxiliary purposes. An augmented combustion module is configured between the compressor and the turbine of the engine and receives compressed air from the compressor and ignites an air/fuel-mixture to turn a shaft that can be used to produce power. The shaft may be coupled with an electrical power generator, a pump, a hydraulic or pneumatic power generator and/or power conversion or transmission devices and/or coupled with the main shaft of the turbine engine. The power from a power generator may be stored in a battery, hydraulic accumulator or pneumatic accumulator and may be used to power auxiliary electrical, hydraulic or pneumatic devices.