A method for inhibiting particulate ingress in a combustion system includes acquiring particulate ingress data, determining a probable particulate ingress location on an air inlet of the combustion system in dependence upon the particulate ingress data, and deploying a filter screen at the determined probable particulate ingress location to inhibit particulate ingress at the determined probable particulate ingress location.

A control system for turbine systems configured to utilize an intelligent model of particulate presence and accumulation within turbine systems to address engine maintenance, erosion, corrosion, and parts failure mitigation is disclosed. The control system may build an intelligent model of fluid flow based on the data value measured by at least one sensor and based on a database of known data values to provide an estimation of amount of ingress of air intake particles into the turbine system, fouling within the turbine system, erosion of at least a portion of the turbine system, and performance degradation rate of the turbine system.

Systems, assemblies, and methods to enhance the efficiency of operation of a gas turbine engine may include a turbine housing positioned to at least partially enclose the gas turbine engine, and a filtration assembly connected to the turbine housing to supply at least partially filtered intake air to an inlet assembly associated with the gas turbine engine. The filtration assembly may include a pre-cleaner including one or more inertial separators configured to separate a first portion of particles and/or liquid from ambient air supplied to the gas turbine engine, thereby to provide at least partially filtered intake air, and one or more filters positioned downstream of the pre-cleaner to separate a second portion of the particles and/or liquid from the at least partially filtered intake air.

Systems, assemblies, and methods to enhance the efficiency of operation of a gas turbine engine may include a turbine housing positioned to at least partially enclose the gas turbine engine, and a filtration assembly connected to the turbine housing to supply at least partially filtered intake air to an inlet assembly associated with the gas turbine engine. The filtration assembly may include a pre-cleaner including one or more inertial separators configured to separate a first portion of particles and/or liquid from ambient air supplied to the gas turbine engine, thereby to provide at least partially filtered intake air, and one or more filters positioned downstream of the pre-cleaner to separate a second portion of the particles and/or liquid from the at least partially filtered intake air.

An inertial particle separator (IPS) for a gas turbine engine, has: inner and outer walls extending about a central axis, an inlet defined between the inner and outer walls and oriented axially; swirling vanes extending at least radially between the inner and outer walls and circumferentially distributed around the central axis, the swirling vanes configured for inducing a circumferential component in an airflow flowing between the swirling vanes; a plenum between the inner and outer walls downstream of the swirling vanes, the plenum circumferentially extending about the central axis, the outer wall converging toward the central axis in a direction of the airflow; and a splitter radially between the inner and outer walls downstream of the plenum and circumferentially extending around the central axis, a particle outlet radially between the splitter and the outer wall, an air outlet radially between the inner wall and the splitter.

An inertial particle separator (IPS) for a gas turbine engine, has: inner and outer walls extending about a central axis, an inlet defined between the inner and outer walls and oriented axially; swirling vanes extending at least radially between the inner and outer walls and circumferentially distributed around the central axis, the swirling vanes configured for inducing a circumferential component in an airflow flowing between the swirling vanes; a plenum between the inner and outer walls downstream of the swirling vanes, the plenum circumferentially extending about the central axis, the outer wall converging toward the central axis in a direction of the airflow; and a splitter radially between the inner and outer walls downstream of the plenum and circumferentially extending around the central axis, a particle outlet radially between the splitter and the outer wall, an air outlet radially between the inner wall and the splitter.

A combustion chamber (15) comprising a wall at least partially defining a combustion zone and having a first surface (41) facing away from the combustion zone and a second surface (43) facing the combustion zone, the wall having at least one effusion cooling aperture (69, 73) extending there-through from the first surface to the second surface, the effusion cooling aperture having an inlet in the first surface and an outlet in the second surface, the first surface having a particle separator (84) at least partially located upstream of the inlet of the effusion cooling aperture, the particle separator projecting away from the first surface and away from the combustion zone.

A combustion chamber (15) comprising a wall at least partially defining a combustion zone and having a first surface (41) facing away from the combustion zone and a second surface (43) facing the combustion zone, the wall having at least one effusion cooling aperture (69, 73) extending there-through from the first surface to the second surface, the effusion cooling aperture having an inlet in the first surface and an outlet in the second surface, the first surface having a particle separator (84) at least partially located upstream of the inlet of the effusion cooling aperture, the particle separator projecting away from the first surface and away from the combustion zone.

A turboprop gas turbine engine mountable to an aircraft has an engine core and a gearbox driving a propeller, the engine core and the gearbox being enclosed within a nacelle. The propeller is located rearward of the gearbox and the engine core relative to a direction of travel of the aircraft. An air intake is disposed within the nacelle and formed to direct ambient air into the engine core. The air intake includes an air inlet duct, having a forward-facing intake inlet receiving the ambient air, with an upstream section and a downstream section. The upstream section is in fluid communication with the intake inlet and extends downstream from the intake inlet. The downstream section fluidly connects to and directs air from the upstream section into the engine air inlet. A second air outlet duct is located within the nacelle and directs air into an air-cooled-oil-cooler (ACOC).

A turboprop gas turbine engine mountable to an aircraft has an engine core and a gearbox driving a propeller, the engine core and the gearbox being enclosed within a nacelle. The propeller is located rearward of the gearbox and the engine core relative to a direction of travel of the aircraft. An air intake is disposed within the nacelle and formed to direct ambient air into the engine core. The air intake includes an air inlet duct, having a forward-facing intake inlet receiving the ambient air, with an upstream section and a downstream section. The upstream section is in fluid communication with the intake inlet and extends downstream from the intake inlet. The downstream section fluidly connects to and directs air from the upstream section into the engine air inlet. A second air outlet duct is located within the nacelle and directs air into an air-cooled-oil-cooler (ACOC).