Abstract: A turbine nozzle includes radially outer and inner bands between which extend a plurality of vanes. The vanes include impingement baffles having impingement holes for impinging cooling air inside the vane for cooling thereof. An air circuit is disposed in the inner band and includes a plurality of outlet holes. A transfer tube is disposed in flow communication between respective ones of the baffles and the air circuit for channeling thereto a portion of the cooling air from inside the baffles as pre-impingement air for discharge through the outlet holes.

Abstract: A turbine disk cooling system for a gas turbine engine of a type having a compressor section including first and second disks connected by a seal and defining a compressor interstage volume, a turbine section including first and second stage disks connected by a seal and defining a turbine inner stage volume, and a compressor shaft interconnecting the compressor disks with the turbine disks including orifices located in the compressor seal for allowing air in the compressor section to enter the compressor interstage volume, and a radial inflow impeller for guiding the compressor air towards the compressor shaft. The impeller including a plurality of radially-extending conduits suspended in the compressor interstage volume, the conduits having a plurality of interior, radially-extending passages for conveying the compressor air from the orifice to the compressor shaft.

Abstract: An apparatus and method for minimizing compressor casing distortion due to bleed air extraction wherein the compressor casing includes an annular outer casing having at least one outlet duct, an annular inner casing disposed co-axially with the outer casing and spaced radially inwardly therefrom to define a plenum, two radial discs extending inwardly from the outer casing defining a plenum, a plurality of small area bleed openings located in the inner casing for imparting a radial velocity to the bleed air greater than the circumferential velocity entering the plenum, an air tube located over each bleed opening, and a baffle plate located on top of the air tubes and communicating with the radial discs for supporting the casings, whereby the bleed air impinges on the outer casing and controls distortion by sufficiently reducing the circumferential temperature gradient forming at the outer casing.

Abstract: An aerodynamically efficient flow transfer device having a means to transfer flow from a static element to a rotor element such that the end of the exit flow is substantially parallel to an exit plane that is perpendicular to the axis of rotation of the rotor and substantially tangential to the operational rotational direction of the rotor. The preferred embodiment provides a cooling air flow transfer apparatus, between a stationary compressor and turbine rotor, having an inducer which includes cooling air flow holes or passages that are acutely angled in a tangential manner to the rotational direction of the rotor. The passages include a cylindrical section leading to a downstream flared outlet in the form of an open channel that has a back wall with a portion that curves to be parallel to the exit plane of the inducer at the channel's end.

Abstract: A turbine blade squealer tip has a plurality of holes defined through the end cap of the blade tip portion and located contiguous with an interior surface of the squealer tip wall along at least one of the turbine blade pressure and suction sides of the tip cap for providing effective cooling air on the squealer tip wall. Each tip hole has an inner portion which communicates with the interior of the blade and an outer portion which communicates with a cavity defined by the blade end cap and the peripheral wall of the squealer tip. The cross-section of the inner hole portion is circular and the inner hole portion overall defines a right cylinder. The cross-section of the outer hole portion is rectangular and the outer hole portion overall defines a rectangular trapezoid.

Abstract: An aft entry system in a gas turbine engine is provided for cooling selected one or more of aft high and low pressure turbine stages of the engine. The aft entry system includes an air flow circuit for routing low pressure cooling air radially outward from an interstage region of a multi-stage compressor of a core engine of the gas turbine engine and axially in an aft direction to a plurality of radial passages, such as provided by stationary stator and outlet guide vanes, which provide an aft entry region to the selected high and low pressure turbine stages. The pressure of the cooling air routed by the air flow circuit of the aft entry system can be lower than the pressure of air at a discharge end of the compressor but higher than the pressure of the gas stream discharging from the selected turbine stage.