Abstract: A combustion chamber of a combustion system has a combustion space, a support structure, a support element, and a heat shield. The heat shield has at least two segments, and each segment includes a liner element facing the combustion space and has an edge region, a gap communicating with the combustion space being formed between edge regions of adjacent segments, and a retaining device. The retaining device fixes the respective liner element on the support structure via the support element and forms a flange region that fits over the edge region of the respective liner element. The retaining device forms a first cooling passage with the support element and has at least one through-opening in the flange region. A cooling gas flows through the through.opening from the first cooling passage to the edge region.

Abstract: Apparatus for combustion of a fuel-oxidizer mixture in a combustion chamber of a turbogroup, in particular of a power plant wherein total oxidizer flow is divided into a main oxidizer flow and a secondary oxidizer flow. The main oxidizer flow is lean mixed with a main fuel flow in a premix burner, and the mixture is fully oxidized in the combustion chamber. The secondary oxidizer flow is divided into a pilot oxidizer flow and a heat-exchanging oxidizer flow. The pilot oxidizer flow is rich mixed with a pilot fuel flow, and the mixture is partially oxidized in a catalyst, with hydrogen being formed. Downstream of the catalyst, the partially oxidized pilot fuel-oxidizer mixture and the heat-exchanging oxidizer flow are together introduced into at least one zone which is suitable for stabilizing the combustion of the main fuel-oxidizer mixture.

Abstract: A pollution-free propulsion engine includes a rotating arm, a hollow axle defining a fuel delivery chamber, and hydrogen and oxygen sources. The rotating arm is formed with a detonation chamber, an opening and two tubular ducts therebetween. The axle is inserted into the opening. A pair of holes is formed in the axle to establish paths of fluid communication from the fuel delivery chamber through the ducts and into the detonation chamber as the rotating arm turns. The hydrogen source comprises a thin palladium binding layer deposited onto an aluminum sheet. Hydrogen molecules that are trapped in the binding layer are released, and the hydrogen is fed into the delivery chamber, through one duct and into the detonation chamber. At the same time, oxygen is delivered into the detonation chamber through the other duct, and the oxygen-hydrogen combination is detonated to release energy, which is converted into mechanical energy.

Abstract: A method for assembling a premixing injector is provided. The method includes providing a centerbody including a center axis and a radially outer surface, and providing an inlet flow conditioner. The inlet flow conditioner includes a radially outer wall, a radially inner wall, and an end wall coupled substantially perpendicularly between the outer wall and the inner wall. Each of the outer wall and the end wall include a plurality of openings defined therein. The outer wall, the inner wall, and the end wall define a first passage therebetween. The method also includes coupling the inlet flow conditioner to the centerbody such that the inlet flow conditioner substantially circumscribes the centerbody, such that the inner wall is substantially parallel to the centerbody outer surface, and such that a second passage is defined between the centerbody outer surface and the inner wall.

Abstract: A gas turbine engine assembly includes at least one propelling gas turbine engine and an auxiliary engine used for generating power. The propelling gas turbine engine includes a fan assembly and a core engine downstream from said fan assembly. The core engine includes a compressor, a high pressure turbine, a low pressure turbine, and a booster turbine coupled together in serial-flow arrangement such that the booster turbine is rotatably coupled between the high and low pressure turbines. The auxiliary engine includes at least one turbine and an inlet. The inlet is upstream from the high pressure turbine and is in flow communication with the propelling gas turbine engine core engine, such that a portion of airflow entering the propelling engine is extracted for use by the auxiliary engine.

Abstract: An injection nozzle for a turbomachine includes a main body having a first end portion that extends to a second end portion defining an exterior wall having an outer surface. A plurality of fluid delivery tubes extend through the main body. Each of the plurality of fluid delivery tubes includes a first fluid inlet for receiving a first fluid, a second fluid inlet for receiving a second fluid and an outlet. The injection nozzle further includes a coolant delivery system arranged within the main body. The coolant delivery system guides a coolant along at least one of a portion of the exterior wall and around the plurality of fluid delivery tubes.

Abstract: The turbine section of the turbine engine is provided with a flow of cooling air which is taken from a compressor section of the turbine engine. The air received from the compressor section is itself cooled before the air is delivered to the turbine. Heat is removed from the flow of air by a plurality of heat pipes which conduct heat away from the flow of air to lower the temperature of the air before it is provided to the turbine.

Abstract: Several methods of increasing the thrust and energy efficiency of charged particle jet engines operating in a gaseous or liquid medium have been developed. We identify the three main components of charged particle thrust generation and provide means to take maximum advantage of each. We also describe several methods to reduce the energy associated with the generation of charged particles and to minimize the number of charged particles needed to further increase energy efficiency. In addition to the methods used to increase thrust and energy efficiency, we have also developed several methods of efficiently controlling the amount and direction of thrust. Finally, we show many uses of these charged particle jet engines and ways to control them.

Abstract: A bypass turbofan gas turbine engine is started by means of electric starter motor mounted directly about a downstream end or upstream of the low pressure spool of the engine. This causes air to be driven by a fan through a bypass duct around the engine casing. Closures close to substantially seal an outlet of the bypass duct, and the air is directed into a combustion chamber of the engine and through the turbines, causing the high pressure spool to pick up speed for starting.

Abstract: Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative spool assembly for a gas turbine engine, which incorporates a compressor, a turbine and a gear assembly, includes: a shaft operative to be driven by the turbine; a first spool segment operative to couple the shaft to the compressor; and a second spool segment operative to couple the shaft to the gear assembly. The first spool segment and the second spool segment are not coupled to each other.

Abstract: A pilot assembly for a fuel injector of a gas turbine engine may include a longitudinal passageway having an outlet end. A mass flow in the longitudinal passageway may generally flow towards the outlet end during operation of the engine. The pilot assembly may also include a liquid fuel nozzle that is positioned to direct a mixture of liquid fuel and air near the outlet end, and a compressed air inlet that is configured to direct air compressed by a compressor of the engine to a compressor discharge pressure into the longitudinal passageway without a substantial loss of pressure. The longitudinal passageway may also include a flow restriction section. The flow restriction section may be a narrowed section of the longitudinal passageway, in which an upstream side of the flow restriction section may have compressed air at substantially the compressor discharge pressure and a downstream side may have air at a lower pressure and a higher velocity.

Abstract: An exhaust duct for directing a stream of combustion gases in a gas turbine engine is provided. The exhaust duct comprises an inlet for receiving the stream of combustion gases from a turbine section of the gas turbine engine; an outlet in fluid communication with the inlet; and a transition portion defining a passage between the inlet and the outlet. The passage comprises a discontinuous annular region having a first side portion and a second side portion that are separated by a flow splitter. At least one bypass channel interconnects the first side portion and the second side portion of the discontinuous annular region.

Abstract: Methods and systems for controlling a combustor for a gas turbine engine are provided. According to one example embodiment, a system includes an air control assembly associated with at least one air path of a combustor for a gas turbine engine. Additionally, the system also includes at least one sensor for sensing at least one operating parameter of the gas turbine engine. Further, the system also includes a controller operable to receive at least one operating parameter sensed by the at least one sensor, and further operable to selectively control an air control assembly based at least in part on the at least one operating parameter sensed by the at least one sensor.

Abstract: A fuel distribution apparatus has a manifold head, a waisted portion and a nozzle head. An inlet channel is formed in a mounting surface of a casing of a jet engine. An annular inlet manifold is formed in the manifold head to receive fluid from the fuel inlet channel. Six outlet fluid passageways extend substantially tangentially away from the annular inlet manifold and open out into an annular outlet manifold. Three ports are formed in the inner wall of the annular outlet manifold and the ports are connected to fuel delivery tubes which extend away from the ports in a substantially helical fashion around the waisted portion into a nozzle head. The outlet fluid passageways are arranged to direct the fluid around the outlet manifold in a first direction and the helical fuel delivery tubes extend in the opposite direction.

Abstract: An exhaust nozzle assembly provides fluidic thrust vectoring of the primary stream to enhance aircraft maneuverability. The exhaust nozzle assembly includes a cooling air passage that supplies cooling air. A vector opening injects cooling airflow into the exhaust passage to the primary stream. A cover plate is movable to partially block the vector openings and adjust the location of injection of the cooling airflow into the exhaust passage.

Abstract: A machine component with a base body made of a base material, a part of the surface of which has been equipped with cladding material having a hardness greater in comparison to the base material. The cladding material is segmented and made up of a number of cladding segments that are spaced apart on the machine component. Each cladding segment may be disposed within a respective recess on the surface of the base body forming strips of the base body and material, and each such strip is disposed between consecutive segments forming a continuous surface.

Abstract: A mid turbine frame of a gas turbine engine includes an outer case which supports a spoke casing co-axially positioned therein. The spoke casing has load transfer spokes extending radially from an inner case and secured to the outer case. A load transfer device is provided to transfer load from the spokes to the outer case in addition to load transfer through a first group of fasteners securing the spokes to the outer case, thereby forming a secondary load transfer path from the spokes. The load transfer device includes an opening of the outer case into which at least some of the spokes are inserted.

Abstract: A combustor cap is cooperable with an impingement plate in a turbine fuel nozzle and includes a plurality of cap segments independently securable to the impingement plate. The plurality of cap segments are radially and tangentially movable relative to the impingement plate.

Abstract: A method of operating a turbine engine, wherein the turbine engine includes a compressor, a combustor, a turbine, a plurality of successive axially stacked stages that include a row of circumferentially spaced stator blades and circumferentially spaced rotor blades, and a plurality of circumferentially spaced injection ports disposed upstream of a first row of stator blades in the turbine; the injection ports comprising a port through which cooling air is injected into the hot-gas path of the turbine, the method comprising: configuring the stator blades in the first row of stator blades such that the circumferential position of a leading edge of one of the stator blades is located within +/?15% pitch of the first row of stator blades of the circumferential location of the injection port midpoint of at least a plurality of the injection ports.

Abstract: A gas turbine engine having an axial flow compressor, an annular combustion chamber, a turbine, and a diffuser. The diffuser includes a flow-dividing element formed by an inner deflecting flank and an outer deflecting flank that divides a compressed gas flow into two partial flows at a branching point. The two deflecting flanks define: an angle of less than 90° along at least a portion of the deflecting flanks, and an angle between 15° and 90° between the deflecting flank angle bisector and the turbine longitudinal axis. The deflector also includes a main deflecting region arranged upstream of the branching point and directed at an acute angle from the turbine longitudinal axis toward an inner combustion chamber shell.