A hybrid compressed air energy storage system is provided. A method of operation thereof includes compressing air during a storage period, and extracting thermal energy therefrom to produce a cooled compressed air. The cooled compressed air may be stored in an air storage unit, the extracted thermal energy may be stored in a thermal storage device, and the stored cooled compressed air may be heated with the stored extracted thermal energy to produce a heated compressed air during a generation period. The heated compressed air may be expanded with an expander to generate power and discharge an expanded air, which may be heated with a recuperator to produce a heated expanded air. A fuel mixture including the heated expanded air may be combusted to produce an exhaust gas, which may be expanded with a second expander to generate power and discharge the expanded exhaust gas to the recuperator.

A hybrid compressed air energy storage system is provided. A method of operation thereof includes compressing air during a storage period, and extracting thermal energy therefrom to produce a cooled compressed air. The cooled compressed air may be stored in an air storage unit, the extracted thermal energy may be stored in a thermal storage device, and the stored cooled compressed air may be heated with the stored extracted thermal energy to produce a heated compressed air during a generation period. The heated compressed air may be expanded with an expander to generate power and discharge an expanded air, which may be heated with a recuperator to produce a heated expanded air. A fuel mixture including the heated expanded air may be combusted to produce an exhaust gas, which may be expanded with a second expander to generate power and discharge the expanded exhaust gas to the recuperator.

Method and computer-readable model for additively manufacturing a ducting arrangement (10) for a gas turbine engine are provided. Ducting arrangement (10) may include a duct (18) to be fluidly coupled to receive a cross-flow of combustion gases from a main combustion stage. Duct (18) includes a duct segment (23) with an expanding cross-sectional area (24) where one or more injector assemblies (26) are disposed. Injector assembly (26) includes one or more reactant-guiding structures (27) arranged to deliver a flow of reactants to be mixed with the cross-flow of combustion gases. The ducting arrangement is effective to reduce total pressure loss while providing an effective level of mixing of the injected reactants with the passing cross-flow. Respective duct components or the entire ducting arrangement may be formed as a unitized structure, such as a single piece using a rapid manufacturing technology, such as 3D Printing/Additive Manufacturing (AM) technologies.

The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper.

The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper.

A fuel nozzle for a combustor of an aircraft engine includes a nozzle body defining an a fuel passage, extending therethrough between a fuel inlet and a fuel outlet located at the outlet end that at least partially defines a nozzle tip, for directing a fuel flow into the combustor via the nozzle tip. A core air passage extends through the nozzle body for directing a core air flow into the combustor via the nozzle tip. At least two flow restrictors are disposed in series within the core air passage, the flow restrictors including an upstream flow restrictor and a downstream flow restrictor each having an orifice therein. The restricted air flow passage having a cross-sectional area smaller than that of the core air passage. The orifice in the upstream flow restrictor being at least partially offset from the orifice in the downstream flow restrictor.

A diffuser of a thermal energy machine, in particular of a gas turbine, has a diffuser inlet, a diffuser outlet, and a plurality of air-guiding elements, wherein an air mass flow enters the diffuser through the diffuser inlet, and wherein the air mass flow that has entered the diffuser exits the diffuser through the diffuser outlet and flows off as a plurality of partial air mass flows by the air-guiding elements. At least two immediately adjacent air-guiding elements of the plurality of air-guiding elements are designed in such a way that the flow-off angles thereof with respect to the circumferential surface formed by the outlet opening of the diffuser outlet extending circumferentially in the circumferential direction differ from each other.

A diffuser of a thermal energy machine, in particular of a gas turbine, has a diffuser inlet, a diffuser outlet, and a plurality of air-guiding elements, wherein an air mass flow enters the diffuser through the diffuser inlet, and wherein the air mass flow that has entered the diffuser exits the diffuser through the diffuser outlet and flows off as a plurality of partial air mass flows by the air-guiding elements. At least two immediately adjacent air-guiding elements of the plurality of air-guiding elements are designed in such a way that the flow-off angles thereof with respect to the circumferential surface formed by the outlet opening of the diffuser outlet extending circumferentially in the circumferential direction differ from each other.

A turbomachine including a turbine shaft supported by at least one bearing, at least one enclosure, housing the bearing of the turbine shaft, an oil cooling circuit of the enclosure including at least one jet configured to inject oil from the cooling circuit into the enclosure, and a regulator configured to regulate the flow of oil in the cooling circuit as a function of an oil temperature at output of the enclosure and a pressure difference at the level of the jet.

A turbomachine including a turbine shaft supported by at least one bearing, at least one enclosure, housing the bearing of the turbine shaft, an oil cooling circuit of the enclosure including at least one jet configured to inject oil from the cooling circuit into the enclosure, and a regulator configured to regulate the flow of oil in the cooling circuit as a function of an oil temperature at output of the enclosure and a pressure difference at the level of the jet.