Abstract: A hybrid combustor for a gas turbine engine includes a plurality of circularly arrayed ceramic can combustors whose outlets communicate with the inlet of an annular, metal combustor. The combustion process is continuous through the plurality of can combustors and into the single annular combustor. Preferably only fuel-rich combustion occurs within each of the can combustors, and fuel-lean combustion continues within the single annular combustor.

Abstract: The present invention provides a self-sealing segmented deswirl molded from an organic matrix composite. Each deswirl segment is comprised of a length of outer shroud, an integral mounting flange extending radially outward from the shroud, and several integral deswirl vanes extending radially inward from the shroud. A means for sealing between segments is provided at the circumferential ends of the mounting flange. The sealing means includes an integral flexible sealing member extending from one end of the mounting flange, and a mating slot for sealingly receiving the sealing member at the other end of the mounting flange.

Abstract: A control apparatus and method for controlling the start-up of a turbine engine, including a compressor and a fuel control system, uses a sensor for sensing a first parameter associated with the start-up, generally the exit temperature of the turbine engine, along with a second sensor for sensing a second parameter associated with the start-up, generally the rate of change of speed of the compressor. These sensed parameters are compared with desired start-up characteristics to determine a first error and a second error. A first look-up table based on fuzzy logic is utilized to determine a first graded membership that is associated with the first error, while a second look-up table based on fuzzy logic is utilized for determining a second graded membership associated with the second error. A third look-up table based on fuzzy logic is utilized for determining an output compensation factor for the start-up based on the graded memberships.

Abstract: A digital real time x-ray system includes a lead room and a control room. The lead room houses a conventional x-ray source and a microfocus x-ray source and a manipulator for the microfocus x-ray source. The manipulator includes a first bracket which supports the microfocus x-ray source, a second bracket which supports an inspection tray and a third bracket which supports an image intensifier. The first, second and third brackets are controlled for single axis movement and the inspection tray is controlled for two axis movement to allow for accurate positioning of parts being x-rayed.

Abstract: For use in a compressor unit of gas turbine engine, a blade having a tip portion. An abrasive portion is formed on the tip portion with the abrasive portion comprising a dispersion of discrete particles of cubic boron nitride disposed on the tip portion. A shroud is coated with a porous ceramic abradable material based on preferably 8% yttria-stabilized zirconia. The abrasive portion of the tip portion contacts the abradable material. In the preferred embodiment, the abradable material is treated with boron nitride composited polyester that is burned out of the material via thermal exposure to thereby improve porosity within the abradable material.

Abstract: A system for the destruction of volatile organic compounds while generating power. In a preferred embodiment the system comprises a combustor and a reaction chamber connected to an exit of the combustor. A primary inlet to the combustor supplies a primary fuel to the combustor. A secondary fuel, comprising air and an amount of one or more volatile organic compounds, is supplied to a compressor, which compresses the secondary fuel and directs the secondary fuel to the combustor and the reaction chamber. The system is suitably configured to enable the stoichiometric reaction of the two fuels in a manner sufficient to destroy the volatile organic compounds contained in the secondary fuel and power a turbine engine connected to an exit of the reaction chamber.

Abstract: A gas turbine engine having a compressor rotor and turbine rotor in back-to-back relation and annular groove therebetween. A nonrotating seal is disposed in the groove. A dual pilot ring is disposed at the radial inner end of the groove. The dual pilot ring includes an inner annular ring and an outer annular ring radially spaced apart to define a clearance gap. The inner ring is mounted to radially facing piloting surfaces on both the compressor and turbine rotors and maintains concentricity between the two rotors as they grow at different thermal rates. The outer annular ring transfers axial loads between the rotors.

Abstract: Disclosed is an oil film cooled face seal that is comprised of an annular stator having a flat surface in rubbing contact with the flat sealing surface of a rotor. The rotor has a flange portion extending radially from a base portion. The base portion has a plurality of holes extending radially therethrough in fluid communication with a source of cooling oil. One side of the flange portion comprises the flat sealing surface. The opposite side, or backface, has a plurality of external radially spaced apart annular grooves criss-crossing a plurality of external circumferentially spaced apart substantially radial grooves. Each of the radial grooves is aligned with and in fluid communication with a corresponding hole in the base portion, and extends fully across the backface.

Abstract: A transition assembly for directing the gas flow from a combustor to an axial turbine includes a transition liner having a shroud disposed about its outer surface and abutting thereto at a plurality of points to defining a plurality of cooling air passages therebetween. The downstream end of the assembly is circumscribed by the first stage stator and extends axially to just upstream of the first stage turbine rotor. At this end a plurality of circumferentially spaced struts are mounted between the liner and the shroud to define a plurality of axially facing nozzles which are angled to impart a pre-swirl to the cooling air exiting therefrom.

Abstract: A face seal comprising an annular carbon stator having a flat sealing surface in rubbing contact with a flat sealing surface of an annular metallic rotor. The rotor has a head portion connected to a base portion by a neck portion, that is disposed on the same side of the rotor as the rotor's flat sealing surface. The neck portion has two straight radial surfaces, and is disposed on the same side of the rotor as the rotor's flat sealing surface. The formation of adverse thermal gradients within the rotor can be avoided by adjusting the axial thickness of the neck portion, thereby preventing coning of the rotor.

Abstract: A thermal barrier coating for superalloy turbine engine vanes and blades that are exposed to high temperature gas is disclosed. The coating includes a ceramic layer applied to an aluminide or MCrAlY bond coat by electron beam physical vapor deposition. The ceramic layer has a first portion having unstabilized porosity, a second portion, overlying the first portion, with stabilized porosity, and an outer portion wherein the pores are coated with a noble metal. The stabilized porosity portion along with the noble metal coating reduce the thermal conductivity of the ceramic layer. Stabilizing the porosity renders it more resistant to sintering densification at high temperatures.

Abstract: A bleed air buffer seal arrangement is provided for improved buffer sealing of oil sump seals in a gas turbine engine. In a multiple spool gas turbine engine, bleed air for buffer sealing is obtained through an annular bleed slot or annular array of bleed ports formed in the impeller shroud of a high pressure centrifugal impeller at a location spaced aft or downstream from the leading edge thereof. Bleed air from this location exhibits significant pressure at low engine power conditions to provide satisfactory buffer sealing, without subjecting sump seals to excess pressure or temperature at high engine power conditions.

Abstract: To contain high energy fragments generated by a hub burst, a compressor containment assembly in a gas turbine engine includes a diffuser housing having an arcuate portion that defines a channel which traps the high energy fragments and an axial facing U-shaped which imparts flexibility. The assembly also includes a concave back shroud having a series of grooves that create resistance for the fragments as they slide along its surface. At its radial outer end the back shroud has an annular rim that absorbs some of the energy and delays the shearing of bolts. A bumper is disposed between the back shroud and a turbine wheel to limit the deflection of the back shroud during a hub burst.

Abstract: To contain high energy fragments generated by a hub burst, a compressor containment assembly in a gas turbine engine includes a diffuser housing having an arcuate portion that defines a channel which traps the high energy fragments and an axial facing U-shaped which imparts flexibility. The assembly also includes a concave back shroud having a series of grooves that create resistance for the fragments as they slide along its surface. At its radial outer end the back shroud has an annular rim that absorbs some of the energy and delays the shearing of bolts. A bumper is disposed between the back shroud and a turbine wheel to limit the deflection of the back shroud during a hub burst.

Abstract: A self piloting balance arbor for use in the dynamic balancing of gas turbine engine components having shafts extending therefrom. The arbor includes a cylindrical sleeve for receiving the shaft, and an annular flange at one end. A conical compliant member extends axially and radially inward from the flange to an end portion having two flat surfaces perpendicular to each other. Upon the application of a load to the arbor the two flat surfaces come into contact with the component, thereby correctly positioning the arbor relative to the component.

Abstract: A gas turbine vane or blade having a novel internal structure that allows for cooling under low pressure ratios. The vane has an air inlet passage that communicates with an inner cooling cavity positioned between the air passage and the vane's trailing edge. Disposed within this cavity are deflectors, turning members, ribs, and deflecting pins arranged so as to direct the cooling air through the cavity in a manner that minimizes pressure loss. Thus maintaining the velocity and flow of the cooling air.

Abstract: To contain high energy fragments generated by a hub burst, a compressor containment assembly in a gas turbine engine includes a diffuser housing having an arch portion that defines a channel which traps the high energy fragments and an axial facing U-shaped which imparts flexibility. The assembly also includes a concave back shroud having a series of grooves that create resistance for the fragments as they slide along its surface. At its radial outer end the back shroud has an annular rim that absorbs some of the energy and delays the shearing of bolts. A bumper is disposed between the back shroud and a turbine wheel to limit the deflection of the back shroud during a hub burst.

Abstract: An improved deceleration fuel control system and related method are provided for regulating fuel flow to a combustor of a gas turbine engine to permit deceleration as a maximum rate without risk of combustor blowout. The system includes a controller responsive to combustor air flow and engine shaft speed to calculate a minimum fuel-air ratio to sustain a combustor flame during a maximum rate deceleration condition. The controller then regulates fuel supply to maintain a deceleration fuel-air ratio at a selected safety margin higher than the calculated minimum to sustain combustion.

Abstract: A system for the destruction of volatile organic compounds while generating power. In a preferred embodiment the system comprises a combustor and a reaction chamber connected to an exit of the combustor. A primary inlet to the combustor supplies a primary fuel to the combustor. A secondary fuel, comprising air and an amount of one or more volatile organic compounds, is supplied to a compressor, which compresses the secondary fuel and directs the secondary fuel to the combustor and the reaction chamber. The system is suitably configured to enable the stoichiometric reaction of the two fuels in a manner sufficient to destroy the volatile organic compounds contained in the secondary fuel and power a turbine engine connected to an exit of the reaction chamber.

Abstract: An aircraft cabin pressure control system controls aircraft cabin pressure by reference to either the partial pressure or concentration of oxygen in the aircraft cabin. Data on ambient pressure, oxygen content, and total pressure are employed to control air pressure in the cabin.