Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
1999-11-30
2001-05-08
Ryznic, John E. (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C416S18600A, C416S232000
Reexamination Certificate
active
06227798
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to aircraft gas turbine engine turbine nozzle segments and, more particularly, to cooling of bands between which airfoils of the segments are mounted.
2. Discussion of the Background Art
In a typical gas turbine engine, air is compressed in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow downstream through a high pressure turbine (HPT) having one or more stages including a HIT turbine nozzle and LPT rotor blades. The gases then flow to a low pressure turbine (LPT) which typically includes multi-stages with respective LPT turbine nozzles and LPI rotor blades. Each LPT turbine nozzle includes a plurality of circumferentially spaced apart stationary nozzle vanes supported between radially outer and inner bands. Each turbine rotor stage includes a plurality of circumferentially spaced apart rotor blades extending radially outwardly from a rotor disk which carries torque developed during operation.
The LPT nozzles are typically formed in arcuate segments having a plurality of vanes integrally joined between corresponding segments of the outer and inner bands. Each nozzle segment is supported at its radially outer end by a flange bolted to an annular outer casing. Each vane has a cooled hollow airfoil disposed between radially inner and outer band panels which form the inner and outer bands, a flange portion along a mid portion of the outer band panel, and a cooling air intake duct for directing cooling air through an opening in the flange portion into the hollow interior or cooling circuit of the airfoil. The intake duct has a 90 degree bend between an axially forward intake duct inlet and an axially aft and radially inward intake duct outlet. The 90 degree bend curves radially inwardly from the opening which in the flange portion toward the hollow airfoil and axially as it extends aft and ends at an intake duct outlet. The airfoil, inner and outer band portions, flange portion, and intake duct are typically cast together such that each vane is a single casting. The vanes are brazed together along interfaces of the flange segments, inner band panels, and outer band panels to form the nozzle segment. The intake duct has a significant amount of convective cooling which it conducts to the band locally but not to the middle of the band between the intake ducts or airfoils. This region of the band between intake ducts operates significantly hotter.
Low pressure turbine nozzle bands are often not cooled, however, advanced engine designs with increased thrust to weight ratios operate at higher turbine inlet temperatures that require more cooling. Cooling schemes that use cooling air from the compressor enhance band cooling for a given amount of cooling flow but also have significant negative effects on engine performance. Impingement baffles, film holes, pin banks and trailing edge holes are all cooling features that have been used in production engines for cooling HPT nozzle bands. Cooling holes disposed through the flange have been used to direct cooling air from a cooling air plenum onto the braze joint along the interface between the outer band panels of the flange portions of an LPT nozzle band. Location of the cooling holes that are disposed through the flange include avoiding obstruction by bolt heads in a flange joint with the engine casing and the cooling air jet from the holes travels far before impinging the band at highly stressed areas further aft of the flange.
It is highly desirable to improve LPT band cooling while minimizing the amount of cooling flow used to do so. It is also highly desirable to improve LPT band cooling to prevent cracking along brazed joints to extend the life of the part and time between repairs of the nozzle segments and vane assemblies. It is desirable to have the impingement jet strike the band farther aft than currently possible without increasing the distance the jet target travels to impingement, thus, minimizing jet velocity decay and improving the convection over the band from the jet. It is also desirable to allow more flexibility in choosing impingement jet location, orientation and angle to the band surface, thus, permitting maximization of cooling effect.
SUMMARY OF THE INVENTION
A gas turbine engine nozzle segment includes at least two circumferentially adjacent vanes joined together along an interface between the vanes. Each of the vanes includes a hollow airfoil disposed between radially inner and outer band panels and a cooling air intake duct leading to a hollow interior of the airfoil for directing cooling air into the hollow interior. The intake duct has a duct wall protruding radially outward from the outer band panel and at least one impingement cooling hole disposed through the intake duct wall and circumferentially and radially inwardly angled.
A flange portion extends circumferentially along a mid portion of the outer band panel and an opening in the flange portion is in fluid communication with the intake duct. The hollow airfoil, radially inner and outer band panels, intake duct wall, and flange portion are integrally formed and preferably integrally cast such that the vane is a single piece integrally cast vane.
REFERENCES:
patent: 4017207 (1977-04-01), Bell et al.
patent: 4187054 (1980-02-01), Landis, Jr. et al.
patent: 4214851 (1980-07-01), Tuley et al.
patent: 5344283 (1994-09-01), Magowan et al.
patent: 5593277 (1997-01-01), Proctor et al.
patent: 5762471 (1998-06-01), Cunha
patent: 5772398 (1998-06-01), Noiret et al.
patent: 5813832 (1998-09-01), Rasch et al.
patent: 5997245 (1999-12-01), Tomita et al.
patent: 6019572 (2000-02-01), Cunha
“F404 Training Guide”, Published by General Electric Company, Aircraft Engine Business Group, Technical Training Operation, Third Issue, Jul., 1982, 6 pages.
Currin Aureen C.
Demers Daniel E.
Gledhill Mark D.
Tsai Gene C. F.
General Electric Company
Hess Andrew C.
Ryznic John E.
Young Rodney M.
LandOfFree
Turbine nozzle segment band cooling does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with Turbine nozzle segment band cooling, we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and Turbine nozzle segment band cooling will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-2526467