Abstract: A die cast system in which an external shell and an internal core usable to form a component of a gas turbine engine are formed together is disclosed. In at least one embodiment, the external shell and internal core may be formed from at the same time via a selective laser melting process, thus eliminating the need for using the conventional lost-wax casting system. In at least one embodiment, the external shell and internal core may be formed a ceramic material that may support receiving molten metal to form a turbine component. Once formed, the external shell and internal core may be removed to reveal the turbine component.

Abstract: Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with asymmetric forward and aft ridge surface area density. The forward ridges have greater surface area density than the aft ridges to compensate for greater ridge erosion in the forward zone during engine operation and reduce blade tip wear in the aft zone. Some abradable component embodiments increase forward zone ridge surface area density by incorporating wider ridges than those in the aft zone.

Abstract: A compressor (10) configured for use in a gas turbine engine (12) and having a rotor assembly (14) with a pumping system (16) positioned on a rotor drum (18) to counteract reverse leakage flow at a gap (20) formed between one or more stator vane tips (22) and a radially outer surface (24) of the rotor drum (18). The pumping system (16) may be from pumping components (26) positioned radially inward of one or more stator vane tips (22) to reduce, if not completely eliminate, reverse leakage flow at the stator vane tips (22). In at least one embodiment, the pumping component (26) may be formed from one or more cutouts (28) in the outer surface (24) of the rotor drum (18). In another embodiment, the pumping component (26) may be formed from at least one pumping fin (30) extending from the radially outer surface (24) of the rotor drum (18). In at least one embodiment, rows (32) of pumping components (26) may be aligned with rows (34) of stator vanes (36) within the compressor (10).

Abstract: A shroud cooling system (100) configured to cool a shroud (50) adjacent to an airfoil within a gas turbine engine (10) is disclosed. The turbine engine shroud (50) may be formed from shroud segments (34) that include a plurality of cooling air supply channels (40) extending through a forward shroud support (52) for impingement of cooling air onto an outer radial surface of the shroud segment (34) with respect to the inner turbine section of the turbine engine (10). The channels (40) may extend at various angles (42) to increase cooling efficiency. The backside surface (62) may also include various cooling enhancement components configured to assist in directing, dispersing, concentrating, or distributing cooling air impinged thereon from the channels (40) to provide enhanced cooling at the backside surface (62).

Abstract: A method of interpreting a rule and a rule-interpreting apparatus for rule-based security apparatus, and an apparatus implementing the method. The method includes the following steps: designating a suspicious timeslot; if any packet does not present in the designated timeslot, capturing current incoming packets or capturing other incoming packets in the designated timeslot next time; automatically associating the packets in the designated timeslot to form at least one traffic flow corresponding to a connection or call; analyzing the at least one traffic flow to select at least one suspicious target traffic flow; and outputting the at least one selected suspicious target flow.

Abstract: An abradable turbine component, a method of creating a turbine component with an abradable mesh structure, and a gas turbine engine are provided. The abradable turbine component includes a turbine component surface for coupling to a turbine casing, and a deposited abradable mesh structure coupled to the turbine component surface. The abradable mesh structure includes interlacing strands of material, each strand including a height relative to the turbine component surface. At least two of the plurality of interlacing strands include a height different from each other. The method includes applying a bond coat layer followed by a thermal barrier coating layer. An abradable mesh structure is deposited on top of thermal barrier coating wherein the abradable mesh structure includes interlacing strands of material wherein at least two of the interlacing strands include a height different from each other. A gas turbine engine including the abradable turbine component is also provided.

Abstract: Gas turbine engine blade squealer tips incorporate cooling slots formed in the suction side rail downstream of the leading edge for directing cooling gas flow along an inside edge of the squealer tip pressure side rail. Some embodiments incorporate a tip fin on the suction side rail proximal a cooling slot. Segmented suction side rail embodiments abrade opposing turbine casing abradable surfaces prior to potential contact with the pressure side rail, reducing likelihood of pressure side rail friction heating. During turbine engine operation cooler pressure side rails reduce likelihood of squealer tip erosion.

Abstract: A component in a gas turbine engine includes an airfoil extending radially outwardly from a platform associated with the airfoil. The airfoil includes opposed pressure and suction sidewalls, which converge at a first location defined at a leading edge of the airfoil and at a second location defined at a trailing edge of the airfoil opposed from the leading edge. The component includes a built-up surface adjacent to the leading edge at an intersection between the pressure sidewall and the platform, and at least one cooling passage at least partially within the built-up surface at the intersection between the pressure sidewall and the platform. The at least one cooling passage is in fluid communication with a main cooling channel within the airfoil and has an outlet at the platform for providing cooling fluid directly from the main cooling channel to the platform.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite, non-inflected, bi-angle, “hockey stick” like pattern abradable surface ridges and grooves. Some embodiments include distinct forward upstream and aft downstream composite multi orientation groove and vertically projecting ridges planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. In some embodiments the grooves are split or divided into multiple sections to interrupt flow traveling inside the groove and cause a local pressurization that reduces tip leakage flow. Some ridge or rib embodiments also have first lower and second upper wear zones. The lower zone optimizes engine airflow characteristics while the upper zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Abstract: A rim seal arrangement for a gas turbine engine includes a first seal face on a rotor component, and a second seal face on a stationary annular rim centered about a rotation axis of the rotor component. The second seal face is spaced from the first seal face along an axial direction to define a seal gap. The seal gap is located between a radially outer hot gas path and a radially inner rotor cavity. The first seal face has a plurality of circumferentially spaced depressions, each having a depth in an axial direction and extending along a radial extent of the first seal face. The depressions influence flow in the seal gap such that during rotation of the rotor component, fluid in the seal gap is pumped in a radially outward direction to prevent ingestion of a gas path fluid from the hot gas path into the rotor cavity.

Abstract: A method of forming an airfoil (12), including: abutting end faces (72) of cantilevered film hole protrusions (64) extending from a ceramic core (50) against an inner surface (80) of a wax die (68) to hold the ceramic core in a fixed positional relationship with the wax die; casting an airfoil including a superalloy around the ceramic core; and machining film cooling holes (34) in the airfoil after the casting step to form an pattern of film cooling holes comprising the film cooling holes formed by the machining step and the cast film cooling holes (102) formed by the film hole protrusions during the casting step.

Abstract: A cooling system (10) positioned within a turbine airfoil (12) useable in a turbine engine and having cooling channels (16) positioned within a platform (18) of the turbine airfoil (12) with exhaust outlets (20) at the pressure and suction side edges (22, 24) to prevent hot gas ingestion under the platform (18) is disclosed. The cooling channels (16) may be formed from main channels (26) extending from cooling fluid supply channels (64) aligned with the airfoil (12) and branch channels (30) extending between the main channels (26) and the pressure or suction side edges (22, 24). The cooling system (10) reduces the cooling surface area adjacent to the airfoil fillet (32) at the intersection (34) of the platform (18) and airfoil (12) and increases cooling surface area adjacent to the pressure side and suction side mate faces (22, 24) as compared with conventional designs.

Abstract: A turbine rotor blade includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channel connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite, non-inflected, bi-angle, “hockey stick” like pattern abradable surface ridges and grooves. Some embodiments include distinct forward upstream and aft downstream composite multi orientation groove and vertically projecting ridges planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. In some embodiments the grooves are split or divided into multiple sections to interrupt flow traveling inside the groove and cause a local pressurization that reduces tip leakage flow. Some ridge or rib embodiments also have first lower and second upper wear zones. The lower zone optimizes engine airflow characteristics while the upper zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Abstract: A composite core die includes a reusable core die; and a disposable core die. The disposable core die is in physical communication with the reusable core die and surfaces of communication between the disposable core die and the reusable core die serve as barriers to prevent the leakage of a slurry that is disposed in the composite core die.

Abstract: A method to augment a plurality of IPS or SIEM evidence information is provided. The method may include monitoring a plurality of processes associated with a computer system. The method may also include identifying a plurality of processes that have network activity. The method may further include capturing the identified plurality of processes that have network activity. The method may also include storing the identified captured plurality of processes that have network activity. The method may include monitoring a plurality of selected programs associated with an operating system of the computer system. The method may also include identifying a plurality of selected programs that have network activity. The method may further include capturing a plurality of screen capture images associated with the identified plurality of selected programs. The method may include storing, by the second component the captured plurality of system process activity.

Abstract: A cooling channel (36, 36B) cools an exterior surface (40 or 42) or two opposed exterior surfaces (40 and 42). The channel has a near-wall inner surface (48, 50) with a width (W1). Interior side surfaces (52, 54) may converge to a reduced channel width (W2). The near-wall inner surface (48, 50) may have fins (44) aligned with a coolant flow (22). The fins may highest at mid-width of the near-wall inner surface. A two-sided cooling channel (36) may have two near-wall inner surfaces (48, 50) parallel to two respective exterior surfaces (40, 42), and may have an hourglass shaped transverse sectional profile. The tapered channel width (W1, W2) and the fin height profile (56A, 56B) increases cooling flow (22) into the corners (C) of the channel for more uniform and efficient cooling.

Abstract: A resonance chamber (42) has an outer wall (32) with coolant inlet holes (34A-C), an inner wall (36) with acoustic holes (38), and side walls (40A-C) between the inner and outer walls. A depression (33A-C) in the outer wall has a bottom portion (50) that is close to the inner wall compared to peaks (37A-C) of the outer wall. The coolant inlet holes may be positioned along the bottom portion of the depression and along a bottom portion of the side walls to direct coolant flows (44, 51) toward impingement locations (43) on the inner wall that are out of alignment with the acoustic holes. This improves impingement cooling efficiency. The peaks (37A-C) of the outer wall provide volume in the resonance chamber for a target resonance.

Abstract: A method to augment a plurality of IPS or SIEM evidence information is provided. The method may include monitoring a plurality of processes associated with a computer system. The method may also include identifying a plurality of processes that have network activity. The method may further include capturing the identified plurality of processes that have network activity. The method may also include storing the identified captured plurality of processes that have network activity. The method may include monitoring a plurality of selected programs associated with an operating system of the computer system. The method may also include identifying a plurality of selected programs that have network activity. The method may further include capturing a plurality of screen capture images associated with the identified plurality of selected programs. The method may include storing, by the second component the captured plurality of system process activity.

Abstract: A method to augment a plurality of IPS or SIEM evidence information is provided. The method may include monitoring a plurality of processes associated with a computer system. The method may also include identifying a plurality of processes that have network activity. The method may further include capturing the identified plurality of processes that have network activity. The method may also include storing the identified captured plurality of processes that have network activity. The method may include monitoring a plurality of selected programs associated with an operating system of the computer system. The method may also include identifying a plurality of selected programs that have network activity. The method may further include capturing a plurality of screen capture images associated with the identified plurality of selected programs. The method may include storing, by the second component the captured plurality of system process activity.