A combined heat and power system comprises a shaft (4), a compressor (6) coupled to the shaft to compress intake gas to form compressed gas; a recuperator (10) to heat the compressed gas to form heated compressed gas; a combustor (12) to combust a fuel and the heated compressed gas to form combustion gas; a turbine (8) coupled to the shaft to expand the combustion gas to form exhaust gas; a load (24) coupled to the shaft; an exhaust outlet (18) to expel the exhaust gas to a heater for heating a fluid based on heat from the exhaust gas; a recuperator channel (28) providing a path for the exhaust gas to flow from the turbine to the exhaust outlet through the recuperator; and a bypass channel (22) providing a path for the exhaust gas to flow from the turbine to the exhaust outlet bypassing the recuperator.

There is described a method and system for in-flight start of an engine. The method comprises rotating a propeller; generating electrical power at an electric generator embedded inside a propeller hub from rotation of the propeller; transmitting the electrical power from the electric generator to an engine starter mounted on a core of the engine via an electric power link; and driving the engine with the engine starter to a sufficient speed while providing fuel to a combustor to light the engine to achieve self-sustaining operation of the engine.

A double walled stator housing includes a first stator housing wall, a second stator housing wall located radially outward from the first stator housing wall, and an air gap located between the first and the second stator housing walls. The housing also includes at least one support structure attached to the first stator housing wall and the second stator housing wall, spanning the air gap and configured to minimize heat transfer between the first wall and the second wall.

A turbofan engine (10), including a ducted fan (22) and an engine (16) having a nozzle (24) for expelling burnt gases (N1) by the engine in the upstream direction of the turbofan engine, the free end (26) of the nozzle being located upstream of the fan (22).

An aircraft gas turbine engine (110) comprises first and second non-coaxial propulsors (113a, 113b), each propulsor (113a, 113b) being driven by a common gas turbine engine core (176) comprising a propulsor drive turbine (143) arranged to drive the first and second propulsors (113a, 113b) via a propulsor drive coupling (127). The core (176) further comprises a first core module (190) comprising a first compressor (129) and a first turbine (131) interconnected by a first shaft (177), and a second core module (191) comprising a second compressor (128) and the propulsor drive turbine (143) interconnected by a second shaft (127), the first and second core modules (190, 191) being axially spaced.

A gas turbine engine (10) comprises: a low pressure compressor (18) comprising a centrifugal compressor stage (24); a low pressure turbine (44) configured to drive a load (12), the low pressure turbine (44) being provided rearwardly of the low pressure compressor (18); a high pressure turbine (42) provided rearwardly of the low pressure turbine (44); a combustor (40) provided rearwardly of the high pressure turbine (42); a high pressure compressor (32) provided rearwardly of the combustor (40); and an intercooler heat exchanger (28) configured to exchange heat between core airflow (B) exiting the low pressure compressor (18) and fan airflow (A), wherein the high pressure turbine (42) and high pressure compressor (32) are coupled by a high pressure shaft (46), and the low pressure turbine (44), low pressure compressor (18) and load (12) are coupled by a low pressure shaft (22).

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 nozzle formed of one piece for a jet engine includes a mixing tube, a fuel conduit integrally formed with the mixing tube, and an opening through the fuel conduit and directed radially into the mixing tube.

A gas turbine engine for an aircraft includes a compressor, a combustion chamber, and a turbine having at least one stator, and at least one rotor. Each stator and rotor is formed by a plurality of blades, a fluid channel is formed between two consecutive blades, and each blade has two opposing surfaces. The compressor is in fluid communication with a first group of stator channels, and the combustion chamber is in fluid communication with a second group of stator channels, such that heat exchange can be performed through two opposing surfaces of at least one stator blade. The outer and the inner walls define a duct for the passage of the heated fluid through the rotor blades, and the outer wall is also arranged for directing the compressed air towards the combustion chamber.

A gas turbine engine includes a rotatable first shaft, a first disk connected to the first shaft, a second disk connected to the first shaft, a combustor radially outward from the first disk and the second disk, and a heat exchanger connected to the combustor aft of the second disk. The first disk includes a row of low pressure compressor blades and a row of high pressure turbine blades connected to a radially outer end of the row of low pressure compressor blades. The second disk includes a row of high pressure compressor blades and a row of low pressure turbine blades connected to a radially outer end of the row of high pressure compressor blades.