Rotary kinetic fluid motors or pumps – Bearing – seal – or liner between runner portion and static part – Between blade edge and static part
Reexamination Certificate
1999-06-08
2001-10-30
Verdier, Christopher (Department: 3745)
Rotary kinetic fluid motors or pumps
Bearing, seal, or liner between runner portion and static part
Between blade edge and static part
C415S173100, C060S039300, C029S464000, C029S889100, C029S889200, C029S889210
Reexamination Certificate
active
06309177
ABSTRACT:
TECHNICAL FIELD
The invention is directed to a method of assembling a gas turbine engine using a concentricity ring between mating components to compensate for cumulative assembly tolerances, and in particular to achieve an extremely close tolerance which is essential in assembling high pressure gas turbines within a mating turbine shroud to minimize the blade tip gap and optimize engine performance.
BACKGROUND OF THE ART
In the assembling of a gas turbine engine several interconnecting components are assembled together, all of which are machined with high precision although within a specified dimensional tolerance. The cumulative effect of these tolerances in assemblies of multiple parts must be accommodated especially when extremely high precision in assembly is required.
One example of such precision is in the assembly of high-pressure turbines and their associated turbine shrouds immediately downstream of a combuster in the gas turbine engine. It has been estimated that the tip clearance, between the tips of turbine blades and the turbine shroud surrounding the rotating blades, is so critical to engine performance that each 0.001-inch of excess tip gap results in approximately a 0.25% decrease in engine performance. As a result, extreme care is taken in ensuring that high-pressure turbine blades are accurately assembled together with their associated turbine shrouds.
In the prior art the method of assembling high-pressure turbines and their shrouds involves assembling the engine including the combuster and high-pressure turbine shaft within acceptable assembly tolerances. The high-pressure turbines are then mounted to the high-pressure turbine shaft and invariably the rotational axis of the turbine is somewhat eccentric of the longitudinal axis of the combuster due to accumulation of machining tolerances.
In order to accurately fit the turbine shroud about the blades of the turbine with minimal tip gap, it is common practice to custom grind the internal surface of each shroud to precisely match the eccentricity and outside diameter of the turbine blades in operation.
As will be appreciated, the withdrawal of the turbine shroud from assembly operations and precision grinding involve significant delay in the assembly operation as well as high costs as a result of the skilled labour involved in the process. In effect, while the custom grinding operation is being carried out on the turbine shrouds, the assembly operation is halted imposing significant manufacturing difficulties due to space and scheduling commitments.
In general, all machinery assembly operations can benefit from the use of standardized components rather than custom fitting each component as required. Standardization of components and simplification of the assembly process, inevitably will reduce costs and increase the speed of production.
It is an object of the present invention, therefore, to eliminate the custom grinding of turbine shrouds and permit the standardization of turbine shroud manufacture regardless of the eccentricity produced in assembly of turbines used in gas turbine engines.
It is a further object of the invention to provide a simple means by which eccentricity of turbines can be accommodated without custom grinding or removal of components from the flow of assembly and production.
It is a further object of the invention to enable eccentricity of the assembled turbines to be accommodated without significant alteration to the prior art structure of gas turbine engine components. Ideally, only minimal modification is desirable.
DISCLOSURE OF THE INVENTION
The invention provides a method of assembling a gas turbine engine using an eccentic concentricity ring to compensate for assembly tolerances.
A typical gas turbine engine has a longitudinal engine axis with at least one turbine disposed on a shaft assembly at a predetermined axial position for rotation about a longitudinal shaft axis. The turbine includes a hub with a circumferentially spaced apart array of turbine blades each having a radially outward blade tip. The blade tips during turbine rotation define a tip surface of rotation concentric the shaft axis.
Due to cumulative tolerances in part manufacture and assembly, the shaft axis is invariably radially eccentric the engine axis at the axial position of the turbine an assembly eccentricity within the range between zero and a predetermined allowable assembly tolerance. The engine is assembled progressively to the stage where the static outer engine housing assembly has an engine housing flange with a cylindrical internal surface concentric the engine axis, and a static turbine shroud having a shroud flange with an external cylindrical surface is to be removably mounted to the engine housing flange with an internal turbine shroud surface of rotation matching the blade tip surface of rotation.
The invention relates to the step of positioning a concentricity ring with: an outer cylindrical surface engaging the internal cylindrical surface of the engine housing flange; and an inner cylindrical surface engaging the external cylindrical surface of the turbine shroud flange, the outer and inner cylindrical surfaces of the concentricity ring being eccentric a ring eccentricity within the range between zero and said predetermined allowable assembly tolerance to rectify the assembly eccentricity in a simple expeditious manner.
Further details of the invention and its advantages will be apparent from the detailed description and drawings included below.
REFERENCES:
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patent: 4222708 (1980-09-01), Davison
patent: 4343592 (1982-08-01), May
patent: 4548546 (1985-10-01), Lardellier
patent: 4723075 (1988-02-01), German
patent: 4786232 (1988-11-01), Davis et al.
patent: 5193377 (1993-03-01), Sendzimir et al.
patent: 5400133 (1995-03-01), Hinton et al.
patent: 5409656 (1995-04-01), Naruse et al.
patent: 5562408 (1996-10-01), Proctor et al.
patent: 5713671 (1998-02-01), Egeter
Moase William Donald
Swiderski Joseph John
Astle Jeffrey W.
Pratt & Whitney Canada Corp.
Verdier Christopher
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