Abstract: A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system (10) for the downstream turbine assembly.

Abstract: A cooling system configured to cool aspects of the turbine engine between a compressor and a turbine assembly is disclosed. In at least one embodiment, the cooling system may include one or more mid-frame cooling channels extending from an inlet through one or more mid-frame torque discs positioned downstream of the compressor and upstream of the turbine assembly. The inlet may be positioned to receive compressor bleed air. The mid-frame cooling channel may be positioned in a radially outer portion of the mid-frame torque disc to provide cooling to outer aspects of the mid-frame torque disc such that conventional, low cost materials may be used to form the mid-frame torque disc rather than high cost materials with capacity to withstand higher temperatures. The cooling fluid routed through the mid-frame cooling channel in the mid-frame torque disc may be exhausted into a cooling system (10) for the downstream turbine assembly.

Abstract: An active bypass flow control system for controlling bypass compressed air based upon leakage flow of compressed air flowing past an outer balance seal between a stator and rotor of a first stage of a gas turbine in a gas turbine engine is disclosed. The active bypass flow control system is an adjustable system in which one or more metering devices may be used to control the flow of bypass compressed air as the flow of compressed air past the outer balance seal changes over time as the outer balance seal between the rim cavity and the cooling cavity wears In at least one embodiment, the metering device may include an annular ring having at least one metering orifice extending therethrough, whereby alignment of the metering orifice with the outlet may be adjustable to change a cross-sectional area of an opening of aligned portions of the outlet and the metering orifice.

Abstract: An active bypass flow control system for controlling bypass compressed air based upon leakage flow of compressed air flowing past an outer balance seal between a stator and rotor of a first stage of a gas turbine in a gas turbine engine is disclosed. The active bypass flow control system is an adjustable system in which one or more metering devices may be used to control the flow of bypass compressed air as the flow of compressed air past the outer balance seal changes over time as the outer balance seal between the rim cavity and the cooling cavity wears. In at least one embodiment, the metering device may include a valve formed from one or more pins movable between open and closed positions in which the one pin at least partially bisects the bypass channel to regulate flow.

Abstract: A floating brush seal in a rim cavity of a turbine in a gas turbine engine, where the floating brush seal includes a seal holder in which the floating brush seal floats, and a expandable seal that fits within two radial extending seal slots that maintains a seal with radial displacement of the floating brush seal and the seal holder.

Abstract: A floating air riding seal for a gas turbine engine with a rotor and a stator, an annular piston chamber with an axial moveable annular piston assembly within the annular piston chamber formed in the stator, an annular cavity formed on the annular piston assembly that faces a seal surface on the rotor, where the axial moveable annular piston includes an inlet scoop on a side opposite to the annular cavity that scoops up the swirling cooling air and directs the cooling air to the annular cavity to form an air cushion with the seal surface of the rotor.

Abstract: The present invention is an industrial gas turbine engine with a circulating fluidized bed cooling system to provide cooling to certain parts of the engine such as a transition duct or a stator vane. The circulating fluidized bed cooling system can be used to cool a stator vane, a casing, or any stationary part of the engine. A circulating fluidized bed cooling system uses very fine particles that pass along with a cooling fluid such as cooling air and provide for a much higher heat transfer coefficient than does turbulent flow cooling air because of the particles. The fine particles produce conduction cooling from the hot surface to the particles that are then carried along in the circulating fluid flow to another location where the heat picked up by the particles is then transferred out from the particles.

Abstract: A gas turbine engine having a rotor with blades and a stationary vane, a platform seal is formed between the blade and vane for inhibiting ingestion of hot gas from a hot gas flow through the turbine into turbine wheel spaces, the platform seal including axial extending platforms on the blade and vane, and radial extending fingers extending from the platforms and forming restrictions between the fingers and the platforms, and a buffer cavity formed between the restrictions, where the fingers are so arranged in a generally radial direction that the vane can be removed from the turbine engine in a radial direction without having to remove the blades first. In additional embodiments, the platform seal assembly can have two or three buffer cavities formed between additional restrictions.

Abstract: An active bypass flow control system for controlling bypass compressed air based upon leakage flow of compressed air flowing past an outer balance seal between a stator and rotor of a first stage of a gas turbine in a gas turbine engine is disclosed. The active bypass flow control system is an adjustable system in which one or more metering devices may be used to control the flow of bypass compressed air as the flow of compressed air past the outer balance seal changes over time as the outer balance seal between the rim cavity and the cooling cavity wears. In at least one embodiment, the metering device may include a valve formed from one or more pins movable between open and closed positions in which the one pin at least partially bisects the bypass channel to regulate flow.

Abstract: An active bypass flow control system for controlling bypass compressed air based upon leakage flow of compressed air flowing past an outer balance seal between a stator and rotor of a first stage of a gas turbine in a gas turbine engine is disclosed.

Abstract: A turbine inter-stage seal with active clearance control. The inner diameter of the stator vane includes a seal support structure with a plurality of segmented seal supports each movable by an actuator located on the vane outer diameter with a plunger or push rod extending through the hollow portion of the vanes. A segmented seal arrangement, such as segmented brush seals, are secured to the underside of the seal support segments and form a complete annular seal for the inter-stage between adjacent rotor disks. A proximity probe or microwave sensor detects the brush seal clearance, and a controller regulates the brush seal clearance.

Abstract: A triple acting radial seal used as an interstage seal assembly in a gas turbine engine, where the seal assembly includes an interstage seal support extending from a stationary inner shroud of a vane ring, the interstage seal support includes a larger annular radial inward facing groove in which an outer annular floating seal assembly is secured for radial displacement, and the outer annular floating seal assembly includes a smaller annular radial inward facing groove in which an inner annular floating seal assembly is secured also for radial displacement. A compliant seal is secured to the inner annular floating seal assembly. The outer annular floating seal assembly encapsulates the inner annular floating seal assembly which is made from a very low alpha material in order to reduce thermal stress.

Abstract: A turbine disc used in a turbo-pump that is exposed to a large thermal gradient across the sides of the rotor disc. The warm side of the turbine disc is covered by a cover plate that forms a sealed space between the turbine disc and the cover plate, and this space is evacuated or insulated to form a complete vacuum or as near as a complete vacuum as possible. The surfaces of the turbine disc and cover plate that form the vacuum space are also coated with a reflective coating such as with a highly polished surface of low emissivity in order to minimize the radiation heat loss through the vacuum space. The turbine disc can have a cover plate forming a vacuum space or insulated cavity on both sides of the turbine disc. Ribs are also used within the vacuum space as supports for the cover plate due to the difference in pressure from the ambient side and the vacuum.

Abstract: A test rig that reproduces high temperature and high pressure conditions found in a gas turbine engine for testing materials under these conditions. The test rig is formed of a plurality of modules connected in series to form the test rig. A specimen rotation module is connected to a mechanical load module, which is connected to a combustor module, which is connected to a test section module on which one or more test specimens are placed, which connects to an exhaust section module, which is connected to a thrust bearing & air inlet section module. The modules are cylindrical in shape and connected in series to form a compact test rig.

Abstract: A gas turbine engine having a rotor with blades and a stationary vane, a platform seal is formed between the blade and vane for inhibiting ingestion of hot gas from a hot gas flow through the turbine into turbine wheel spaces, the platform seal including axial extending platforms on the blade and vane, and radial extending fingers extending from the platforms and forming restrictions between the fingers and the platforms, and a buffer cavity formed between the restrictions, where the fingers are so arranged in a generally radial direction that the vane can be removed from the turbine engine in a radial direction without having to remove the blades first. In additional embodiments, the platform seal assembly can have two or three buffer cavities formed between additional restrictions.

Abstract: A test facility is provided for testing materials under high temperature, pressure, and mechanical loads. The facility provides a physically sealed system that simulates conditions in hot sections or gas turbine engines. A test article is coated with a test material and exposed to a hot combusting flow in a test section housing. The article may be a pipe or conduit member oriented at any direction to the flow. A second cooler flow of fluid is channeled through the test article to create a sharp temperature gradient in the test article and through the test material. A liquid-cooled sleeve is disposed about the test article to create an annular channel of combusting flow over the test article. The downstream end of the second cooler flow is connected to the upstream end of the main hot flow at the combustion chamber.

Abstract: In a gas turbine engine having a rim cavity formed between a rotor disk and a stator, and a brush seal or other seal member providing an airflow seal between the stator and a side plate of the rotor disk, a thermally responsive valve device is located in the bypass passage of the stator to form a parallel air flow path with the brush seal. As the brush seal wears and the leakage airflow around the worn brush seal increases, the air flow temperature in the rim cavity would decrease. The thermally responsive valve device would detect the decreasing airflow temperature, and reduce the airflow passing through the bypass passage in order to prevent the temperature of the rim cavity from decreasing. The temperature responsive valve device makes use of a bimetallic valve member that regulates airflow based upon temperature.

Abstract: A test facility is provided for testing materials under high temperature, pressure, and mechanical loads. The facility provides a physically scaled system that simulates conditions in hot sections of gas turbine engines. A test article is coated with a test material and exposed to a hot combusting flow in a test section housing. The article may be a pipe or conduit member oriented at any direction to the flow. A second cooler flow of fluid is channeled through the test article to create a sharp temperature gradient in the test article and through the test material. A liquid-cooled sleeve is disposed about the test article to create an annular channel of combusting flow over the test article. The downstream end of the second cooler flow is connected to the upstream end of the main hot flow at the combustion chamber.

Abstract: A test facility is provided for testing materials under high temperature, pressure, and mechanical loads. The facility provides a physically scaled system that simulates conditions in hot sections of gas turbine engines. A test article is coated with a test material and exposed to a hot combusting flow in a test section housing. The article may be a pipe or conduit member oriented at any direction to the flow. A second cooler flow of fluid is channeled through the test article to create a sharp temperature gradient in the test article and through the test material. A liquid-cooled sleeve is disposed about the test article to create an annular channel of combusting flow over the test article. The downstream end of the second cooler flow is connected to the upstream end of the main hot flow at the combustion chamber.

Abstract: Featured are a device, system and method for dissipating at least some heat energy generated by one or more heat generating components of a flywheel energy storage system (FESS). The method includes providing a heat pipe member, having first and second ends, and a heat dissipating member thermally engaged with the heat pipe member second end and configured to transfer heat energy therefrom. The method also includes thermally engaging the heat pipe member first end to the FESS so that at least some heat energy generated by the FESS heat generating component is communicated to the first end and thence through the heat pipe member to the heat dissipating member. Further, the method includes locating the heat dissipating member in a heat sink remote from the FESS.