Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
1999-09-07
2001-07-03
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C416S189000
Reexamination Certificate
active
06254345
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to aircraft gas turbine engine turbine blade tip shrouds and seals and, more particularly, to cooling the shroud and tip.
2. Discussion of the Background Art
Gas turbine engines frequently employ tip shrouds on individual airfoils to limit blade amplitudes when vibrating in a random manner and to guide fluid flow over the airfoils. This is particularly true in the low pressure section of a gas turbine engine. Neighboring shrouds abut in the circumferential direction to add mechanical stiffness. When a series of such assemblies are mounted together, the shrouds define in effect a continuous annular surface. Circumferentially opposite edges of the shrouds in the circumferential direction are provided with abutment faces and are designed to introduce to the assembly desired constraints.
Circumferentially extending seal teeth extend radially outwardly from the shrouds to engage seal lands to seal the gas flowpath between the shrouds and casing surrounding the rotor. The seal lands typically are in the form of a honeycomb covered stator shroud.
Gas turbine engines typically include cooling systems which provide cooling air to turbine rotor components, such as turbine blades, in order to limit the material temperatures experienced by such components. Prior art cooling systems usually acquire the air used to cool turbine components from the engine's compressor, after which it is diverted and subsequently directed to the turbine section of the engine through an axial passageway.
Low pressure turbine blades typically are not cooled. High pressure turbine blades which are typically cooled do not have deflection restraining tip shrouds. Supersonic high performance engines are being developed for long distance supersonic operation, such as for the High Speed Commercial Transport (HSCT) engine program. The low pressure turbine blades in the low pressure turbine section are exposed to high temperatures for long periods of time over most of the flight envelope with the engine operating at high power engine settings. It is also desirable to have a low engine weight and engine length.
High speed engines require better cooling techniques than those presently used. One exemplary engine for a high speed civil transport employs a low pressure turbine in close proximity to a high pressure turbine discharge. Furthermore, the engines mission requires long term exposure of the low pressure turbine to very high temperatures at high power engine settings. Aircraft gas turbine engine designers constantly strive to improve the efficiency of the gas turbine engine as well designing an engine which is low weight and short. The use of cooling air increases fuel consumption and, therefore, it is highly desirable to minimize the amount of engine work used to produce the cooling air.
SUMMARY OF INVENTION
A gas turbine engine turbine blade shrouded tip includes an airfoil tip having a cross-sectional airfoil shape, a blade tip shroud attached to the tip, and a shroud cooling circuit disposed within the blade tip shroud. The shroud cooling circuit is operable for cooling substantially all of the shroud and is in fluid communication with a hollow interior of the tip.
In one embodiment of the invention, the tip shroud has at least one circumferentially extending seal tooth on a radially outer shroud surface of the shroud extending in a radial direction away from the hollow interior. Preferably, two or more such seal teeth are employed. In a more particular embodiment of the invention, the tip shroud further includes circumferentially extending and axially spaced apart leading and trailing shroud edges, circumferentially spaced apart clockwise and counter-clockwise shroud side edges. The shroud cooling circuit includes circumferentially extending shroud cooling passages between the clockwise and counter-clockwise shroud side edges. One more particular embodiment of the invention provides forward and aft pluralities of the shroud cooling passages within the tip shroud and in fluid communication with first and second cavities respectively in the hollow interior.
In another embodiment of the invention, a blade having an airfoil with the tip shroud at a tip of the airfoil includes an airfoil cooling circuit in fluid communication with the shroud cooling circuit. In a more particular embodiment of the invention, the blade further includes forward and aft pluralities of the shroud cooling passages in fluid communication with first and second cavities respectively of the airfoil cooling circuit. The airfoil, in a more particular embodiment, has an aspect ratio of at least about 3.
A gas turbine engine turbine assembly includes a plurality of such turbine blades mounted around a periphery of a turbine rotor. The blades have airfoils extending radially outward from blade platforms to tip shrouds at airfoil tips having airfoil shapes and mounted to the rotor by roots extending radially inward from the blade platforms. The hollow interiors of the blades are in fluid communication with rotor cooling passages through the rotor. Typically, each of the hollow interiors includes one of the airfoil cooling circuits in fluid communication with the shroud cooling circuit. An annular sealing assembly is mounted to and within an engine casing and spaced radially apart from the seal teeth so as to provide a gas path seal with the seal teeth. The annular sealing assembly includes a shroud stator supporting a honeycomb material mounted to a radially inwardly facing side of the shroud stator such that the honeycomb material cooperates with the seal teeth to provide the gas path seal.
Apparatus for impingement cooling is used in one embodiment for directing impingement cooling air onto a radially outwardly facing side of the shroud stator. Such apparatus includes, in a more specific embodiment, an external teeth cooling assembly for flowing the impingement cooling air into the flowpath and around the seal teeth after it has impinged the radially outwardly facing side of the shroud stator. One external teeth cooling assembly includes a leakage path between a forward edge of the shroud stator and a support hanger which supports the shroud stator from the engine casing.
ADVANTAGES OF THE INVENTION
The internally cooled tip shroud helps the gas turbine engine to operate at a long period of time at high power engine settings with low pressure turbine blades exposed to very high temperature gas flows. The invention also allows placing the low pressure turbine blades in close proximity to the high pressure turbine discharge and, particularly, in engine designs having counter-rotating high and low pressure turbine rotors with no stators therebetween. Among the benefits of the present invention are lower engine weight and reduced engine length.
The present invention provides efficient cooling to obtain sufficient creep and oxidation component lives for the sustained high power conditions. The invention provides cooling and reduced metal temperatures of the turbine blade tip shroud to levels which allows creep and oxidation life goals to be met. The cooled tip shroud is advantageous because it allows reduction of turbine blade weight and axial length by allowing a more slender blade (higher aspect ratio) to meet vibration frequency requirements. This results from the additional support rendered by the blade to blade constraining effect of the tip shroud, which raises blade frequencies to meet design requirements.
REFERENCES:
patent: 3527544 (1970-09-01), Allen
patent: 4214851 (1980-07-01), Tuley et al.
patent: 4522557 (1985-06-01), Bouiller et al.
patent: 5064343 (1991-11-01), Mills
patent: 5941687 (1999-08-01), Tubbs
Frederick Robert A.
Harris Daniel J.
Timko Lawrence P.
General Electric Company
Hess Andrew C.
Look Edward K.
McAleenan James M.
Young Rodney M.
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