Solar Turbines Inc.

Caterpillar Inc.

Boris Glezer

Philip J Cederwall

Russell B Jones

Aaron R Fierstein

David M Evans

Hamid Bagheri

Narender K Bhardwaj

Gary L Boyd

Paul F Norton

Michael D Fox

Hee Koo Moon

Stuart Alexander Greenwood

Tsuhon Lin

John J Hensley

Hee-Koo Moon

Partha Dutta

Kenneth H Maden

Sheldon Smilo

Gerald D Stringham

Eli H Razinsky

Moon Hee-Koo

David W Dewis

Cooling air delivery systems for gas turbine engines are used to increase component life and increase power and efficiencies. The present system increases the component life and increases efficiencies by better utilizing the cooling air bled from the compressor section of the gas turbine engine. For example, a first portion of cooling fluid cools the leading edge of a turbine blade internally. After first contacting a predetermined area of the component, a portion of that first portion of cooling fluid is then used to film cool the component.

Gas turbine engine airfoils with improved cooling

Today, the trend is to increase the temperature of operation of gas turbine engines. To cool the components with compressor discharge air, robs air which could otherwise be used for combustion and creates a less efficient gas turbine engine. The present low thermal expansion sealing ring support system reduces the quantity of cooling air required while maintaining life and longevity of the components. Additionally, the low thermal expansion sealing ring reduces the clearance 'C','C'' demanded between the interface between the sealing surface and the tip of the plurality of turbine blades. The sealing ring is supported by a plurality of support members in a manner in which the sealing ring and the plurality of support members independently expand and contract relative to each other and to other gas turbine engine components.

Low thermal expansion seal ring support

Many government entities regulated emission from gas turbine engines including CO. CO production is generally reduced when CO reacts with excess oxygen at elevated temperatures to form CO2. Many manufactures use film cooling of a combustor liner adjacent to a combustion zone to increase durability of the combustion liner. Film cooling quenches reactions of CO with excess oxygen to form CO2. Cooling the combustor liner on a cold side (backside) away from the combustion zone reduces quenching. Furthermore, placing a plurality of concavities on the cold side enhances the cooling of the combustor liner. Concavities result in very little pressure reduction such that air used to cool the combustor liner may also be used in the combustion zone. An expandable combustor housing maintains a predetermined distance between the combustor housing and combustor liner.

Combustor for a low-emissions gas turbine engine

Past gas turbine engines have been manufactured with a variety of systems which have attempted to compensate for the varying radial clearance or interface between a tip of a turbine blade and a surrounding shroud. Such systems have engineered in large clearances, utilized special tips on blades and provided abradable structures. The present system controls a radial clearance or an interface or spaced distance between a tip of a blade and a stationary shroud. The system includes an introduction of a cool fluid or a hot fluid flow or a combination of cool fluid and hot fluid into a support case cavity. The flow is controlled by modulating a flow control apparatus and controlling the expansion of a nozzle support case and a stationary shroud relative to the position of the tip of the turbine blade. The system improves efficiency, longevity and operation of the gas turbine engine.

Turbine blade clearance control system

The invention disclosed and taught herein teaches method and apparatus for restricting the ingress of hot combustion gases into the turbine disc plenum of a gas turbine engine as the hot gases flow from a turbine nozzle assembly and into the turbine blades. A jet of engine cooling air is injected into the wake region immediately behind and down stream of the nozzle assembly's stator vanes thereby reducing the static pressure recovery otherwise occurring within the wake region. Thus, a lessor amount of buffering air is required within the turbine disc plenum to prevent the ingress of hot gases thereby resulting in an increased efficiency of the engine's cooling air delivery system and over all engine performance.

Turbine disc ingress prevention method and apparatus

Air flow delivery systems for gas turbine engines are used to increase component life and increase power and efficiencies. The present system increases the component life and increases efficiencies by directing a portion of a compressed air formed by the compressor section and located in a first fluid flow path to a second fluid flow path including a heat exchanger positioned operatively within the second flow path. The second fluid flow path results in a flow of cooled cooling air to be directed to the components of the gas turbine engine to be cooled. Thus, the components of the gas turbine engine are more efficiently cooled with a smaller quantity of air drawn from the compressor section generated air than if compressor generated air was used for cooling components of the gas turbine engine. Furthermore, more compressor generated air remains for combustion within the combustor section of the gas turbine engine.

Turbine cooling cycle

The present gas turbine engine (10) includes a disc assembly (64) defining a disc (66) having a plurality of blades (70) attached thereto. The disc (66) has a preestablished rate of thermal expansion and the plurality of blades have a preestablished rate of thermal expansion being less than the preestablished rate of thermal expansion of the disc (66). A shroud assembly (100) is attached to the gas turbine engine (10) and is spaced from the plurality of blades (70) a preestablished distance forming an interface (108) therebetween. Positioned in the interface is a seal (110) having a preestablished rate of thermal expansion being generally equal to the rate of thermal expansion of the plurality of blades (70).

Ceramic blade with tip seal

Cooling air delivery systems for gas turbine engines are used to increase component life and increase power and efficiencies. The present system increases the component life and increases efficiencies by better utilizing the cooling air bled from the compressor section of the gas turbine engine. For example, a flow of cooling air is directed to a plurality of airfoils having a leading edge and includes a cooling path therein each of the plurality of blades in which is positioned a device which causes the cooling air to swirl and more effectively absorb heat and cool the leading edge of the airfoil.

Turbine cooling system

A method and apparatus are provided for producing a temperature profile of a part operating within a machine is provided. The machine is operated with the part for a predetermined period of time. The part is placed in a light box in which a camera produces an image of the part. The image is compared with a standard set of colors and the standard color most closely corresponding to each pixel of the image is selected.

Method and apparatus for producing a surface temperature map

Cooling air delivery systems for gas turbine engines are used to increase component life and increase power and efficiencies. The present system increases the component life and increases efficiencies by better utilizing the cooling air bled from the compressor section of the gas turbine engine. For example, a flow of cooling air which is directed through internal passages of the engine and into a cavity and is used to cool the interface between a turbine disc and a plurality of turbine blades and further reduces ingestion of hot power gases into the internal portion of the gas turbine engine. The cavity has a generally tapered configuration and effectively cools the interface between the disc and the blade. Thus, increasing component life and increasing power and efficiencies of the engine.

Turbine motor and blade interface cooling system

A turbine wheel assembly includes a disk having a plurality of blades therearound. A ceramic ring is mounted to the housing of the turbine wheel assembly. A labyrinth rim seal mounted on the disk cooperates with the ceramic ring to seal the hot gases acting on the blades from the disk. The ceramic ring permits a tighter clearance between the labyrinth rim seal and the ceramic ring.

Rim seal for turbine wheel

A turbine nozzle vane assembly having a preestablished rate of thermal expansion is positioned in a gas turbine engine and being attached to conventional metallic components. The metallic components having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the turbine nozzle vane assembly. The turbine nozzle vane assembly includes an outer shroud and an inner shroud having a plurality of vanes therebetween. Each of the plurality of vanes have a device for heating and cooling a portion of each of the plurality of vanes. Furthermore, the inner shroud has a plurality of bosses attached thereto. A cylindrical member has a plurality of grooves formed therein and each of the plurality of bosses are positioned in corresponding ones of the plurality of grooves. The turbine nozzle vane assembly provides an economical, reliable and effective ceramic component having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the other component.

Low thermal stress ceramic turbine nozzle

Cooling air delivery systems for gas turbine engines are used to increase component life and increase power and efficiencies. The present system increases the component life and increases efficiencies by better utilizing the cooling air bled from the compressor section of the gas turbine engine. For example, a first flow of cooling air is directed to the nozzle and shroud assembly and another portion of the first flow is exited from the nozzle plenum into an annular reservoir to improve cooling of a rotor assembly and to at least reduce ingestion of hot power gases into the internal portion of the engine. The system further includes a second flow of cooling air which is directed through internal passages of the engine and into the annular reservoir and is used to cool the rotor assembly and to at least reduce ingestion of hot power gases into the internal portion of the gas turbine engine. To insure that the least amount of hot power gas is ingested into the internal portion of the engine is minimized a radial blade flange arrangement and a radial shroud flange arrangement axially overlap each other in a preestablished radially spaced proximity forming a buffering zone therebetween into which the mixed cooling air in the annular reservoir is directed. Thus, increasing component life and increasing power and efficiencies of the engine.

A radially-resilient seal can frequently pop out its annular groove defined in a carrier when the carrier and an overlapping housing are telescopically assembled. Such misorientation of the seal prevents proper seating of the seal against the housing and ultimately permits leakage of pressurized fluid past the seal. The present invention provides a compact sealed joint and fool-proof method of assembly which maintains proper orientation of a radially-resilient seal during telescopic assembly of a carrier and a housing. The sealed joint includes an annular member having a diametrically tapered internal surface. The annular member positively ensures that the seal will properly seat against a housing inner surface which is not only physically inaccessible to the assembler but is also hidden from the assembler's view by an outer casing during the telescopic assembly.

Sealed telescopic joint and method of assembly

Cooling air delivery systems for gas turbine engines are used to increase component life and increase efficiencies. To change the flow rate of cooling air in past engines, major disassembly of engine components was required resulting in the loss of the engine's service and excessive costs. With the changeable cooling air delivery system of the present invention, major engine disassembly is eliminated. An elongate member extends into a fluid control passage to establish a preselected flow area which can be changed. The time and expense required to change the cooling flow rate is drastically reduced because the elongate members are accessible external of the engine. The elongate member is removed from the engine and replaced by different elongate member having a different outer diameter of the stem or sleeve. The different diameter results in a new clearance or preselected flow area for cooling air to pass therethrough.

Growing community of inventors

Del Mar, CA, United States of America

Boris Glezer