The present invention relates to an underwater drone which is an unmanned mobile which can move in the water, and more particularly to a communication system for the underwater drone which performs communication between the underwater drone and a land-based controller (or maneuvering device). The present invention also relates to an airlock apparatus for the drone which transfers the drone into or from facilities or containers, or equipment sealed (or closed) against surrounding environment. The communication system for an underwater drone includes an underwater drone (1) configured to move in the water, at least one transmitting and receiving antenna (2) provided in an area where the transmitting and receiving antenna (2) can communicate with the underwater drone (1) by wireless communication, and a controller or a maneuvering device (5) connected to the at least one transmitting and receiving antenna (2) by a wired cable (4) and configured to control the underwater drone (1).

A gas turbine engine (10) comprising: a high pressure turbine (17); a low pressure turbine (19); a high pressure compressor (15) coupled to the high pressure turbine (17) by a high pressure shaft (27); a propulsor (23) and a low pressure compressor (14) coupled to the low pressure turbine (19) via a low pressure shaft (26) and a reduction gearbox (30); wherein the low pressure compressor (14) consists of four compressor stages (14) and defines a cruise pressure ratio of between 2.4:1 and 3.3:1; the high pressure compressor (15) defines a cruise pressure ratio of less than 17:1; and the high pressure compressor (15) and low pressure compressor (14) together define a cruise core overall pressure ratio of greater than 36:1.

A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and a turbine having a high pressure turbine section and a low pressure turbine section driving the fan through a speed reduction mechanism, wherein the low pressure turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

A combined circulating system of a micro gas turbine, a transportation means, and a charging system are provided. The circulating system includes the micro gas turbine, a heat exchange unit, a circulating water tank, a piston engine, and a power generating apparatus, wherein the micro gas turbine is provided with a regenerator; an exhaust port of the regenerator is connected with an air inlet of the heat exchange unit to provide a heat source to the heat exchange unit; the exhaust port of the heat exchange unit is led to atmosphere, a water inlet of the heat exchange unit is connected with a water outlet of the circulating water tank, and a steam outlet of the heat exchange unit is connected with the piston engine to enable high pressure steam to enter the piston engine to push the piston engine to produce work.

A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.

An airflow control system for a gas turbine system according to an embodiment includes: a compressor component of a gas turbine system; a mixing area for receiving an exhaust gas stream produced by the gas turbine system; an air extraction system for extracting a supply of bypass air from an excess flow of air generated by the compressor component of the gas turbine system; an enclosure surrounding the gas turbine system and forming an air passage, the bypass air flowing through the air passage and around the gas turbine system into the mixing area to reduce a temperature of the exhaust gas stream; and an exhaust processing system for processing the reduced temperature exhaust gas stream.

A system for reducing a temperature of an exhaust gas stream of a gas turbine system according to an embodiment includes: a compressor component of a gas turbine system; an airflow generation system for attachment to a rotatable shaft of the gas turbine system, the airflow generation system and the compressor component drawing in an excess flow of air through an air intake section; a mixing area for receiving an exhaust gas stream produced by the gas turbine system; an air extraction system for: extracting at least a portion of the excess flow of air generated by the airflow generation system and the compressor component to provide bypass air; and diverting the bypass air into the mixing area to reduce a temperature of the exhaust gas stream; and a fluid injection system for injecting an atomized fluid into the mixing area to reduce a temperature of the exhaust gas stream.

A power generation system includes a shroud that defines a fluid flow path. A gas turbine engine is in the fluid flow path, and the gas turbine engine includes a compressor, a combustor downstream from the compressor, and a turbine downstream from the combustor. An electric generator is in the fluid flow path upstream from the turbine, and the electric generator includes a rotor coaxially aligned with the turbine. A plurality of non-lubricated bearings rotatably support the gas turbine engine.

A gas turbine engine for an aircraft includes an engine core having a core length and comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius, wherein a ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in a range from 1.2 to 2.0; and wherein the engine core length is in a range from 150 cm to 320 cm.