Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2001-04-26
2002-11-19
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C415S191000
Reexamination Certificate
active
06481959
ABSTRACT:
BACKGROUND OF THE INVENTION
This invention relates generally to improvements in gas turbine engines, particularly with respect to improved thermal isolation of turbine components from high temperature mainstream combustor gases. More specifically, this invention relates to an inhibitor that suppresses the flow of undesired hot gases from the main flow path into internal regions in the engine that are radially inboard of the turbine section main flow path.
Gas turbine engines are generally known in the art for use in a wide range of applications such as aircraft engines and auxiliary power units for aircraft. In a typical configuration, the engine includes a turbine section having a plurality of sets or rows of stator vanes and rotor blades disposed in an alternating sequence along the axial length of a hot gas flow path of generally annular shape. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled, in turn, to a main engine shaft. Hot combustion gases are delivered from an engine combustor to the annular hot gas flow path, resulting in rotary driving of the turbine rotor disks which, in turn, drives the compressors and gearbox.
In most gas turbine engine applications, it is desirable to regulate the normal operating temperature of certain turbine components in order to prevent overheating and potential mechanical failures attributable thereto. That is, while the engine stator vanes and rotor blades are specially designed to function in the high temperature environment of the mainstream hot gas flow path, other turbine components such as the rotor disks are not designed to withstand such high temperatures. Accordingly, in many gas turbine engines, the volumetric space disposed radially inwardly or internally from the hot gas flow path comprises an internal engine cavity through which a cooling air flow is provided. The cooling air flow is normally obtained as a bleed flow from a compressor or compressor stage forming a portion of the gas turbine engine. The thus cooled internal engine cavity results in maintaining the normal steady state temperature of the rotor disks and other internal engine components at or below a selected temperature.
In the past, a relatively high amount of cooling air flow has been required to obtain satisfactory temperature control of engine components within the cooled internal engine cavity. Because use of compressor air for cooling is a performance penalty to the engine, it is highly desirable to employ methods to minimize this cooling air. The demand for cooling air has been significantly impacted by the leakage of cooling air from the internal cavity and through the space between adjacent rows of stator vanes and rotor blades, into the hot gas flow path. In addition, the demand for cooling flow has been affected by a somewhat irregular and unpredictable ingestion of mainstream hot gases from the hot gas flow path into the internal engine cavity. Various attempts to prevent flow between adjacent stator vanes and rotor blades have primarily involved the use of overlapping lip-type structures in close running clearance, often referred to as flow discouragers, but these structures have not been satisfactorily effective in preventing hot gas ingestion.
A variety of alternative baffle-type structures and techniques have been proposed, in addition to traditional flow discouragers, in efforts to minimize hot gas ingestion into the internally cooled cavity of a gas turbine engine. Such alternative approaches have included pockets of complex shape, some of which receive separate flows of cooling gas, to prevent hot gas ingestion. In the past, these techniques have been generally ineffective, or have otherwise required structures of complex shape and/or complex mounting arrangements at the time of initial engine production.
Information relevant to attempts to provide cooling air to minimize exposure of various engine components to elevated temperatures can be found in U.S. Pat. Nos. 1,819,864; 2,858,101; 3,535,873; 3,609,057; 5,466,123; 6,089,822; and 6,109,867. However, further improvements in minimizing hot gas ingestion are desirable to enhance durability and engine performance, since none of the previous inventions has been successful at eliminating hot gas ingestion into internally cooled cavities, and some suffer from one or more of the following disadvantages:
a) no work extraction is possible from the cooling air.
b) the cooling air released into the turbine disk forward cavity is not oriented specifically to the forward turbine disk surface.
c) the cooling air does not inhibit ingestion of the hot flow path gases.
d) hot flow path gases enter the turbine disk forward cavity and are then diluted by a coolant flow. The diluted, but somewhat heated air is then used for cooling purposes.
e) cooling air jets are not oriented to discharge the spent cooling air at the flow path high pressure locations associated with airfoil trailing edges.
f) cooling air flow is metered and delivered to the turbine cavity but no effort is made to directly counteract the ingestion of hot gases into the turbine cavity.
Ho et al., U.S. Pat. No. 5,545,004 assigned to the assignee of this application, disclose a recirculation pocket having a contoured shroud adapted for quick and easy installation at the time of initial engine production, wherein the contoured shroud captures ingested hot gases for effective recirculation into the main hot gas flow path of the engine. However, these recirculation pockets add weight and cost to an engine.
For the foregoing reasons, there is a need for a simple device that inhibits the flow of hot gas into cavities in the turbine sections of gas turbine engines.
SUMMARY OF THE INVENTION
The present invention fulfills the above need and specifically provides the following additional benefits:
a) because of the simplicity of the design, a gas turbine disk cavity ingestion inhibitor system can be readily incorporated on new design engines or it can be economically retrofitted on existing engines.
b) incorporation of a gas turbine disk cavity ingestion inhibitor system will minimize the volume of cooling air flow required to purge turbine disk cavities from ingested hot flow path gases.
c) component life will improve as a result of reduced degradation caused by exposure of metals to excessively hot gas flow.
d) engine performance will be enhanced by virtue of the reduced chargeable cooling core flow, resulting in an improvement in specific fuel consumption.
Turbine engine combustion air is directed through a nozzle ring, which accelerates the hot flow path gases to the proper conditions for the turbine rotor to extract work from the gas. Since the nozzle ring is a stationary part and the rotor is a rotating part, a necessary axial and radial gap exists between these parts, allowing hot gas flow to enter the disk cavity. If enough hot gas enters the disk cavity, the mixed mean temperature of the turbine disk cavity will rise to a temperature which will cause component failure. Typical measures employed in the past to minimize hot gas entry into the disk cavity include introduction of significant amounts of disk cavity cooling purge air, as well as employment of flow discouragers. The latter are comprised of overlapping lip-type structures, protruding from the stationary nozzle ring and the rotating turbine blades, and being in close running clearance to each other. Due to periodic high pressure fields that occur in the area of the turbine nozzle airfoil, these measures have been insufficient to fully overcome the flow of hot gases into the disk cavity at certain circumferential locations.
An object of the present invention is to provide a supplemental system to minimize hot gas ingestion into the circumferential locations of turbine disk cavities, where high pressure fields are experienced. The present invention achieves this object by providing a simple set of cooling air holes located on each side of the turbine nozzle stator airfoil trailing edge, and proximately placed to be below the turbine nozzle flow dis
Freiberg Douglas P.
Howe William J.
Morris Mark C.
Okpara Nnawuihe A.
Wright Carl D.
Desmond, Esq. Robert
Honeywell International , Inc.
Look Edward K.
White Dwayne J.
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