Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2000-06-16
2002-05-07
Kwon, John (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
Reexamination Certificate
active
06382906
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor and ignited for generating hot combustion gases which flow downstream through several turbine stages. A high pressure turbine powers the compressor, and a low pressure turbine powers a fan in a typical turbofan aircraft engine application for powering an aircraft in flight.
A turbine stage includes a stationary turbine nozzle having a plurality of hollow vanes extending radially between outer and inner bands. The vanes have airfoil configurations for guiding the combustion gases between corresponding turbine rotor blades disposed downstream therefrom. The blades are mounted to the perimeter of a rotor disk which rotates during operation for providing power to turn the compressor or fan during operation.
Since the turbine nozzle vanes are heated during operation by the hot combustion gases which flow thereover, cooling air bled from the compressor is channeled inside the vanes for cooling thereof. Vane cooling is quite sophisticated and complex and many patents are found which disclose various cooling configurations.
Nozzle vanes typically include impingement baffles therein through which the cooling air is directed in impingement against the inner surface of the vane for providing internal cooling thereof. The vanes typically also include film cooling holes through the walls thereof for discharging the spent cooling air in a film over the outer surface of the vane for providing additional protection against the hot combustion gases.
The turbine nozzle is typically supported at its outer bands to a surrounding annular casing. The cooling air bled from the compressor is channeled through the casing and into the individual vanes through corresponding flowpaths. Since the turbine nozzle is directly subject to the hot combustion gases, it is heated to greater temperatures than that of the surrounding casing through which the cooling air is provided. Differential thermal expansion and contraction between the turbine nozzle and supporting casing must be accommodated for preventing unacceptably large thermal stresses during operation which would adversely affect the useful life of the nozzle.
Accordingly, a typical turbine nozzle is formed in a plurality of arcuate nozzle segments, with two or more vanes being integrally formed with corresponding outer and inner band segments, typically in unitary castings. The nozzle segments are fixedly joined to the supporting casing and include suitable seals between the adjoining band segments thereof. Segmenting the turbine nozzle interrupts the circumferential continuity thereof and permits unrestrained differential thermal movement between the nozzle and surrounding casing.
In an exemplary second stage turbine nozzle found in commercial use in this country for many years, a perforate impingement ring surrounds the nozzle outer bands for providing impingement cooling thereof. A separate .flowpath into the individual nozzle vanes is provided in this design by a plurality of inlet tubes, commonly referred to as spoolies, which extend radially between the impingement ring and the corresponding impingement baffles of the several vanes.
Each impingement baffle includes an integral inner spoolie receptacle or cup in the form of a tube fixedly joined to the top of the baffle which receives one end of the spoolie. The other end of the spoolie is received in a corresponding outer spoolie cup fixedly joined through the impingement ring.
The respective pairs of outer and inner spoolie cups are generally radially aligned with each other for receiving therebetween respective ones of the spoolies which bridge the cups and provide flow communication for channeling the cooling air into the individual vanes independently of the airflow through the impingement holes of the impingement ring.
A particular problem of this fixed-geometry spoolie design is the differential thermal expansion and contraction between the impingement ring supporting the outer ends of the spoolies and the nozzle segments from which extend the corresponding inner spoolie cups.
The circumferential positions of the inner and outer spoolie cups necessarily change from their initial cold position during assembly of the nozzle components, to their hot positions during normal operation of the engine at elevated temperature.
Accordingly, the differential thermal movement between the outer and inner spoolie cups is conventionally accommodated by introducing relatively large annular bands around both ends of the individual spoolies. The spoolie bands provide frictional seals inside the respective spoolie cups, and permit a limited amount of tilting or cocking of the individual spoolies without binding or spoolie deformation.
In this way, the outer and inner spoolie cups may be initially radially aligned during cold assembly, with the spoolies inserted therein. And at operating temperature of the turbine, the differential movement between the respective pairs of outer and inner spoolie cups is accommodated by tilting or cocking of the individual spoolies about the sealing bands thereof.
Since the spoolie bands extend radially outwardly from the outer surface of the individual spoolies, the corresponding inner diameters of the individual spoolies is inherently smaller. Since each vane requires a predetermined amount of cooling airflow therethrough for effective cooling during operation, the inner diameters of the spoolies must be suitably sized for meeting that airflow requirement. This in turn requires a correspondingly larger diameter for the spoolie bands which increases the overall size of the spoolie and supporting spoolie cups, and corresponding increases engine weight. Increased engine weight decreases overall efficiency of the engine and is a primary design objective for aircraft engines.
Since adjacent vanes in each nozzle segment are fixedly joined together by the integral band segments, they expand and contract differently than adjacent vanes in adjoining nozzle segments which are not integrally connected by the adjoining bands, but are instead joined together by the supporting casing. Differential thermal movement between the spoolie cup pairs may be otherwise accommodated by providing different positions of the inner spoolie cups for each nozzle vane. This is undesirable since it decreases nozzle segment uniformity, and increases the need for a larger inventory of parts.
Accordingly, it is desired to provide a turbine nozzle having an improved configuration of spoolies and supporting cups for accommodating differential thermal movement during operation.
BRIEF SUMMARY OF THE INVENTION
An impingement baffle for a turbine nozzle vane includes a perforate sleeve for discharging impingement air therethrough. A cap closes the sleeve at a top end, and includes an inlet hole. A spoolie cup includes a flange at the bottom thereof which is disposed atop the cap around the inlet hole for channeling air therein. A retainer is joined to the cap to loosely trap the cup flange thereatop for permitting limited lateral sliding movement thereof.
REFERENCES:
patent: 3540810 (1970-11-01), Kercher
patent: 4288201 (1981-09-01), Wilson
patent: 5795128 (1998-08-01), Eichstadt
patent: 6158955 (2000-12-01), Caddell, Jr. et al.
Brassfield Steven Robert
Tressler Judd Dodge
Webb Alan Lionel
General Electric Company
Hess Andrew C.
Kwon John
Young Rodney M.
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