Abstract: A flight vehicle scramjet combustor is provided having two spaced-apart, generally opposing, and generally longitudinally extending walls extending forward and aft. One of the walls includes a generally aft-facing step, a forward wall portion extending generally longitudinally forward of the step, and an aft wall portion extending generally longitudinally aft of the step. The step further includes a first section and an interconnected second section, where in the first section is attached to the forward wall portion and the second section is attached to the aft wall portion. The second section includes a main or primary fuel injector discharge orifice and a plurality of secondary fuel injector discharge orifices positioned adjacent thereto, the primary and secondary fuel injector discharge orifices each having a fuel discharge axis aligned generally perpendicular to the second section which projects both generally toward the other of the walls and longitudinally aft.

Abstract: A tubular wear sleeve includes radially outer and inner surfaces, with the inner surface having at least one protuberance extending radially inwardly therefrom for being received in a complementary indentation in a fluid carrying tube for restraining relative movement therebetween.

Abstract: A method and an apparatus for damping the vibrations of circumferentially extending and generally cantilevered tubing segments of a cooling manifold of a gas turbine engine which is disposed in a generally concentric relationship with and attached to an annular engine casing. The vibration damping means comprises a plurality of clip elements, with each clip element being attached to a pair of axially adjacent ones of the tubing segments at a location proximate free ends of the tubing segments which engage an axially extending manifold plenum in slip joint connections. Each clip element comprises first and second arcuate retainers connected by an arcuate spacer section which establishes the proper axial spacing between the retainers. At least one of the retainers engages the corresponding tubing segments in an interference fit, with each retainer establishing an interference fit in a preferred embodiment.

Abstract: A cyclonic precombustion mixing chamber or prechamber has an axially extending annular cyclonic chamber to vaporize liquid fuel therein and a centerbody centrally disposed within. A fuel injection means introduces liquid fuel radially outward into the cyclonic chamber. The centerbody extends axially through an annular exit of the prechamber to define a throat and throat exit therebetween which is optimally designed for preventing backflash and autoignition. In one embodiment the diameter of the throat is sized to substantially match the diameter of the potential core.

Abstract: A coupling for splining together first and second shafts which may be inner and outer, co-axial annular members, abutting co-axial annular members, or intermeshing gears on parallel shafts. One embodiment is a coupling for splining together inner and outer annular members having corresponding inner and outer cylindrical surfaces concentrically disposed about an axis and having splines which are generally circumferentially disposed on the cylindrical surfaces along an axial length of at least one of the members such that the splines having a varying pitch so as to be curved with respect to the axis. The pitch is defined by a function which varies parabolically with respect to a shaft axis and, in a more particular embodiment, the function has a non-zero fixed pitch component and a parabolic component so that the shape of the splines are both helical and parabolic.

Abstract: A coolable airfoil for use in gas turbine engine component such as a turbine blade or vane is provided with a radially extending airfoil having an outer wall surrounding at least one radially extending cavity and extending chordwise through at least a portion of the airfoil. The outer wall has ribs disposed on an inner surface of the outer wall and cooling passages formed between the outer wall and an insert disposed in the cavity against and in abutting sealing relationship with the ribs, wherein the outer wall is formed from two radially and chordwise extending sections, preferably suction and pressure side sections, bonded together while the insert was disposed inside the cavity. The insert provides an inner wall surrounding a hollow chamber generally conforming to the cavity the insert is disposed in and the insert includes a spring to force the inner wall out against the ribs during the bonding process.

Abstract: A gas turbine engine airfoil includes a tip corner disposed at the juncture of the airfoil tip and trailing edge, and a cooling hole extends therethrough in direct flow communication with an internal cooling passage thereof. A cross-hole extends perpendicularly into the tip corner and into the cooling hole for discharging from the airfoil a portion of cooling air channeled into the cooling hole. The airfoil tip corner is therefore cooled by the air channeled through the cross-hole received from the intersecting cooling hole.

Abstract: An airfoil includes preformed first and second sides joinable together at a bond line extending between first and second opposite edges thereof. The first side includes a first extension having a proximal end forming a first end of the bond line, an intermediate section, and a distal end. The second side includes a second extension extending from the airfoil first edge which is fixedly joinable to the distal end of the first extension. The intermediate section of the first extension is spaced from the airfoil first edge to define therewith a working void so that pressure loading on the intermediate section is carried in part through the bond line first end for diffusion bonding thereof.

Abstract: A gas turbine engine rotor blade includes an airfoil having first and second sidewalls joined together at leading and trailing edges. First and second tip walls extend from adjacent the leading edge along the respective first and second sidewalls to adjacent the trailing edge and are spaced apart to define a tip cavity therebetween. A first notch is disposed in the first tip wall adjacent to the leading edge for channeling into the tip cavity a portion of combustion gases flowable over the airfoil for reducing the heating effect of the gases on the blade tip. In a preferred embodiment, a second notch is disposed adjacent to the trailing edge for promoting flow through the tip cavity from the first notch.

Abstract: A gas turbine engine turbine airfoil includes first and second opposite sidewalls joined together at leading and trailing edges and extending from a root to a tip thereof. The sidewalls have respective outer surfaces over which is flowable combustion gases during operation. The sidewalls are spaced apart to define an internal flow passage for receiving a cooling fluid. A plurality of ejection slots are arranged in a column through the first sidewall and have flow inlets for receiving cooling fluid from the internal flow passage and ejecting the fluid from respective flow outlets along the outer surface. At least some of the ejection slots are inclined at an acute ejection angle which varies in magnitude for preferably matching the respective inclination of combustion gas streamlines flowable over the airfoil surface.

Abstract: A gas turbine engine airfoil includes first and second sidewalls joined together at opposite leading and trailing edges and spaced apart from each other to define at least one internal flow channel for channeling cooling air therein. A plenum is defined by an inner surface of the first sidewall and by an internal wall adjoining the flow channel. A pair of injection holes extend through the internal wall to feed the cooling air into the plenum. A plurality of film cooling holes extend through the first sidewall in flow communication with the plenum for channeling the cooling air in a film along an outer surface of the first sidewall to effect film cooling. The injection holes are aligned to converge toward each other so that cooling air jets discharged therefrom collide with each other inside the plenum to reduce pressure thereof. This regulates the pressure across the film cooling holes to improve film cooling.

Abstract: An active liner for attenuating noise includes a rigid backplate supporting a piezoelectric panel. A pressure transducer is disposed in the panel for sensing acoustic pressure of noise being propagated against a face surface of the panel. A controller includes a predetermined schedule of acoustic impedance for controlling a displacement driver joined to the panel to effect a displacement velocity of the panel face surface for obtaining a predetermined acoustic impedance at the sensed acoustic pressure for attenuating noise.

Abstract: A method of detecting hydraulic leaks from a high pressure fuel hydraulic system is useful in systems where a pump, actuators, and hydraulic lines are filled with engine fuel at high pressure. The method comprises the steps of calculating pump flow based on a core engine RPM and a pump displacement, and calculating actuator fluid flow. System condition signals are then generated to indicate whether the core engine RPM is constant and whether the actuators are quiescent. A servo flow value is then provided, typically at predetermined system conditions, as indicated by the system condition signals, using the pump flow calculation and the actuator fluid flow calculation. Finally, leak flow is determined based on the pump flow calculation, the actuator fluid flow calculation, and the servo flow value, and the leak flow is responded to accordingly.

Abstract: A variable stator vane includes an airfoil with an integral outer trunnion having a seat extending integrally therefrom. A threaded stem extends from the seat and includes a coextensive alignment surface that cooperates with a complementary mounting hole in a lever arm which restrains rotation of the lever arm about the stem during assembly for ensuring a predetermined rotational orientation between the lever arm and the airfoil.

Abstract: A blade support includes first and second abutting disks, with the first disk having a plurality of axial dovetail slots at one end thereof, and a circumferential dovetail first half-slot at an opposite end thereof, with the second disk having a circumferential dovetail second half-slot therein. A plurality of rotor blades have axial dovetails disposed in respective ones of the axial dovetail slots, and circumferential dovetails disposed in part in each of the first and second half-slots. The first and second disks are retained together so that the first and second half-slots collectively support the circumferential dovetails.

Abstract: A blade containment system is disclosed which includes an annular casing positioned radially outward of the blades and in surrounding relationship therewith, the annular casing including a plurality of ribs on the outer radial surface thereof, wherein the ribs limit the propagation of cracks and holes formed in the annular casing.

Abstract: The present invention provides a cooling apparatus with two spaced apart walls, one of which is a coolable wall, having a cooling flowpath between them wherein the cooling flowpath is converging in the direction of the cooling flow. In one embodiment the coolable wall is an afterburner cooling liner that includes a film cooling means on the hot side of the liner. Another embodiment provides a corrugated or wavy wall afterburner cooling liner wherein each wave includes a forward facing surface, from crest to trough of the wave in the direction of the hot gas flow, and an aft facing surface, from trough to crest. A high density pattern of multi-hole film cooling holes is disposed on the hotter of the two surfaces and a lower density pattern on the cooler one.

Abstract: A coolable airfoil for use in gas turbine engine component such as a turbine blade or vane is provided with a one-piece integrally formed double shell outer wall surrounding at least one radially extending cavity. The inner and the outer shells are integrally formed of the same material together with tying elements in the form of continuous ribs which space apart the shells, mechanically and thermally tie the shells together, and form convective cooling passages therebetween.

Abstract: A turbine frame includes outer and inner coaxially disposed annular casings having a plurality of circumferentially spaced apart struts extending therebetween is provided with at least one and preferably two axially spaced apart polygonal stiffening rails, a forward rail and an aft rail, circumferentially disposed on the outer annular casing. The rails have at least a first section with a constant cross-sectional area normal to the casing in a circumferential direction around the outer casing, a linear centroid distribution of the cross-sections in the circumferential direction, and the centroids lie in a first plane P1 which is parallel to and spaced apart from a second plane P2 that is tangential to the outer casing mid-way between adjacent struts.

Abstract: A self bleeding rotor blade and method of operation are disclosed for reducing boundary layer thickness for improved performance. The rotor blade includes an outer surface configured for pressurizing air flowable thereover with bleed apertures being disposed therein for bleeding a portion of boundary layer air from the outer surface during operation and thereby decreasing its thickness for improving aerodynamic performance of the blade.