Abstract: A basesheet for use in an aircraft gas turbine engine exhaust nozzle includes a plurality of basesheet segments extending between longitudinally extending spaced apart basesheet leading and trailing edges, respectively, of the basesheet. Right and left hand basesheet side edges, respectively, extend longitudinally between the basesheet leading and trailing edges. Each basesheet segment has a panel body defined between a pair of segment side edges longitudinally extending between segment leading and trailing edges, respectively. Stiffeners having leading and trailing edge ribs at the segment leading and trailing edges of aft and forward ones of the basesheet segments, respectively, extend widthwise across the panel body between the segment side edges. The leading and trailing edge ribs are joined together only at intermediate sections of the stiffeners which are centered between left and right hand expansion sections, respectively, of the stiffener.

Abstract: A method of fabricating a test model of a gas turbine engine combustor dome, and the test model produced thereby. The method entails individually stamping a plurality of dome wall segments and first and second mounting band segments. Each wall segment comprises at least one cup between radially inward and outward edges of the wall segment, and an opening in the cup. At least one wall segment and its two corresponding mounting band segments are placed on a fixture that locates the opening of the wall segment, locates the first and second mounting band segments at the radially-inward and outward edges of the wall segment, and orients the wall segment to establish a dome angle of the fixtured dome assembly. The wall segment and mounting band segments are then joined while the fixtured dome assembly remains on the fixture to form at least a unitary sector of the test model.

Abstract: A system for attaching liners to a carrier for forming inner surfaces of turbine combustor. The system may include one or more liners attached to a carrier using connectors engaged to the carrier on a cold side of the liner and may include one or more liners attached to the carrier using connectors engaged to a hot side of the liners. The carrier may include one or more access ports for accessing from the hot side connectors engaged to the carrier on the cold side. In at least one embodiment, the system may include attaching all liners, except for one liner, to a carrier using connectors coupled to the carrier on the cold side and may include attaching the one liner to the carrier using connectors coupled to the carrier on the hot side.

Abstract: Method and arrangement for providing liquid fuel rocket engine member (10). The member forms a body of revolution having an axis of revolution and a cross section that varies in diameter along said axis. The wall structure includes a plurality of cooling channels (11). The outside of the wall structure includes a continuous sheet metal wall (14). The cooling channels (11) are longitudinally attached to the inside of the sheet metal wall. The method for manufacturing the rocket engine member (10) includes the steps of forming a sheet metal wall (14) having a wall section corresponding to the desired nozzle section, providing a plurality of channel members (15), and attaching the channel members (15) to the inside of the sheet metal wall (14).

Abstract: Methods for manufacturing outlet nozzles for use in rocket motors are disclosed. The methods include providing an inner wall, positioning an outer wall around the inner wall, positioning spacers in the space between the inner and outer walls, and laser welding between the spacers and at least one of the inner and outer walls in order to form cooling ducts between the spacers and to provide T-shaped joints between the spacers and at least one of the inner and outer walls, the T-shaped joints including a rounded shape within the cooling ducts.

Abstract: The rocket motor nozzle assembly of this invention includes a throat insert and a carbon or silica protective eyelid. The throat insert has a carbon throat support and a refractory metal shell. The shell is positioned radially inside the throat support to cover the inner surface of the throat support. The protective eyelid covers a sufficient portion of the forward surface region of the shell and the underlying converging portion of the throat support to protect these components against particle impingement. The protective eyelid extends sufficiently far forward along the converging/diverging pathway to cover and protect the forward face or edge of the throat insert and prevent the combustion gases from passing under the throat insert and reaching the radially outer surfaces of the throat insert. However, the protective eyelid leaves the throat surface region of the shell exposed to the converging/diverging pathway.

Abstract: A method of fabricating a modular fuel injector permits the fabrication of the electrical group subassembly outside a clean room while a fuel group subassembly is fabricated inside a clean room. The fuel injector comprises a valve group subassembly and a coil group subassembly. The valve group subassembly includes a tube assembly having a longitudinal axis that extends between a first end and a second end; a seat that is secured at the second end of the tube assembly and that defines an opening; an armature assembly that is disposed within the tube assembly; a member that biases the armature assembly toward the seat; an adjusting tube that is disposed in the tube assembly and that engages the member for adjusting a biasing force of the member; a filter that is located at least within the tube assembly; and a first attachment portion.

Abstract: A method facilitates replacing a portion of a gas turbine engine combustor liner. The combustor includes a combustion zone formed by inner and outer liners. The inner and outer liners each include a series of panels, and at least one of the inner and the outer liner includes an aft support coupled to an aft end of liner. The method comprises cutting through at least one of the combustor inner and the outer liner upstream from the liner aft support, removing the combustor liner aft support from the combustor, and installing a replacement aft support within the combustor such that the liner aft support extends aft ward from the portion of the combustor liner that is upstream from the cut.

Abstract: An air-cooled combustor liner segment, including a plurality of spaced apart radially outwardly extending integral heat exchange protuberances or pins is repaired by providing a replacement edge member for a service damaged edge portion including such pins. The damaged edge portion is removed to provide a segment body with a body repair surface, of a first shape, from which the damaged edge portion was removed. A replacement edge member is provided including a plurality of spaced apart radially outwardly extending integral pins to replace pins in the removed edge portion. The replacement edge member includes a replacement edge member repair surface of a second shape matched with the first shape of the body repair surface. The body repair surface and the replacement edge member repair surface are disposed in juxtaposition at a matched junction therebetween so that their respective radially outer surfaces substantially are coextensive.

Abstract: A combustion chamber structure for a rocket engine or the like, has a combustion chamber liner with cooling channels and at least one manifold for feeding and removing a coolant, particularly a cryogenic fuel. The at least one manifold is brazed together with the combustion chamber liner and the area of the combustion chamber liner that is not covered by the at least one manifold is coated with an electroplated structural jacket. The invention also includes a method for manufacturing the combustion chamber structure.

Abstract: A method for forming an article having a wall that is at least partially fabricated from tubing. The methodology employs a pressurized fluid to deform the tubing such that an outer surface of the tubing substantially conforms to a predetermined first wall configuration, an inner surface of the tubing substantially conforms to a predetermined second wall configuration and each of segment or loop of the tubing is substantially contiguous with at least one other segment or loop between the first and second surfaces.

Abstract: A nanophase composite duct assembly and method of fabricating the same are provided that comprise an ultra-high strength nanophase aluminum alloy duct joined with a ceramic particulate reinforced metal matrix fitting, preferably using solid-state friction welding. The nanophase aluminum alloy duct is fabricated by extruding a billet formed by a process of cryogenic milling the alloy, followed by out-gassing, then hot isostatic pressing. The fitting is fabricated by combining a ceramic particulate with a metal matrix, preferably by powder processing or liquid metal infiltration. Further, the solid-state friction welding may comprise inertial welding, friction stir welding, or a combination thereof. As a result, a lightweight duct assembly is provided for high-pressure liquids such as propellants in rocket engines.

Abstract: A combustor for a gas turbine engine includes a deflector assembly that enhances heat transfer from the combustor and minimizes low cycle fatigue stresses induced within the combustor. The deflector assembly includes a plurality of deflectors secured to a spectacle plate. Each deflector has tapered edges and includes a plurality of cylindrical projections extending outward from the deflector to facilitate heat transfer. The projections include rounded edges and are arranged in a high density pattern. The deflector is coated with a thermal barrier coating and a bondcoat to minimize exposure to hot combustion gases or flame radiation.

Abstract: A method for designing and assembling a high performance catalyst bed gas generator for use in decomposing propellants, particularly hydrogen peroxide propellants, for use in target, space, and on-orbit propulsion systems and low-emission terrestrial power and gas generation. The gas generator utilizes a sectioned catalyst bed system, and incorporates a robust, high temperature mixed metal oxide catalyst. The gas generator requires no special preheat apparatus or special sequencing to meet start-up requirements, enabling a fast overall response time. The high performance catalyst bed gas generator system has consistently demonstrated high decomposition efficiency, extremely low decomposition roughness, and long operating life on multiple test articles.

Abstract: A nanophase composite duct assembly and method of fabricating the same are provided that comprise an ultra-high strength nanophase aluminum alloy duct joined with a ceramic particulate reinforced metal matrix fitting, preferably using solid-state friction welding. The nanophase aluminum alloy duct is fabricated by extruding a billet formed by a process of cryogenic milling the alloy, followed by out-gassing, then hot isostatic pressing. The fitting is fabricated by combining a ceramic particulate with a metal matrix, preferably by powder processing or liquid metal infiltration. Further, the solid-state friction welding may comprise inertial welding, friction stir welding, or a combination thereof. As a result, a lightweight duct assembly is provided for high-pressure liquids such as propellants in rocket engines.

Abstract: In a first assembly step a coolant liner is formed having an outside surface. A plurality of coolant channels are formed on the outside surface. At least two inner conformal throat support sections and at least two outer conformal throat support sections are formed. The two inner conformal throat support sections are formed around the outside surface of the coolant liner to form a closed inner conformal throat outside surface. The inner conformal throat support sections are joined at inner throat seam lines. At least two outer conformal throat support sections are assembled around the closed inner conformal throat outside surface so as to cover the inner throat seam lines, to form a liner throat support assembly. An inlet manifold and an outlet manifold are assembled around the liner throat support assembly to form a hot isostatic press (HIP) assembly. A HIP'ed assembly is formed by HIP'ing the HIP assembly.

Abstract: Method and arrangement for providing a liquid fuel rocket engine member (10). The member has a load bearing wall structure (14) having a plurality of cooling channels (11). For improving the heat transfer, a material with a higher thermal conductivity than the load bearing wall structure (14) is applied to the wall structure.

Abstract: Method and arrangement for manufacturing an outlet nozzle (10) for use in a liquid fuel rocket engine. The nozzle forms a body of revolution having an axis of revolution and a cross section that varies in diameter along said axis. The nozzle has a wall structure comprising a plurality of mutually adjacent cooling channels extending substantially in parallel from the inlet end (12) of the nozzle to its outlet end (13). The method includes providing a plurality of preprocessed profile members, each having a web and flanges in opposite directions from said web. Each profile member is machined to present a longitudinally gradually tapering width. The member is curved to conform with the wall section of the nozzle, and members are joined by welding the flanges to form a bell-shaped nozzle structure with cooling channels (11) formed by adjacent webs and adjacent pairs of flanges.

Abstract: A method enables a portion of combustor dome assembly to be replaced. The dome assembly includes at least one dome cup including at least one swirl cup including a joint that defines a slot therein, and a deflector plate assembly that is coupled to, and extends downstream from, the swirl cup, such that a portion of the deflector plate assembly is securely fixed within the joint slot. The method comprises cutting through an interface formed between the deflector plate assembly and the joint slot, extending the cut substantially between the deflector assembly and the joint slot through the at least one dome cup, removing the deflector plate assembly from the combustor dome assembly, and coupling a replacement deflector plate assembly within the combustor dome assembly that extends aft from the portion of the combustor dome assembly that is upstream from the cut.

Abstract: A method facilitates replacing a portion of a gas turbine engine combustor. The combustor includes a combustor liner and a wire-wrapped cowl assembly that includes an inner cowl and an outer cowl. The method comprises cutting through the wire-wrapped cowl assembly upstream from fastener openings used to couple the inner and outer cowls to the combustor liner, removing a portion of the cowl assembly from the combustor, and coupling a replacement cowl to the portion of the existing cowl assembly that is downstream from the cut, wherein the replacement cowl includes an inner annular portion, an outer annular portion, and a plurality of circumferentially-spaced radial ligaments extending therebetween.

Abstract: A method of manufacturing a fuel injector is disclosed. The method includes assembling a power group assembly with a valve group assembly to form the fuel injector. The power group assembly includes an overmold formed over a coil assembly. The valve group assembly includes a tube assembly formed by connecting an inlet tube to a valve body. The valve body includes a seat, a lower guide, and an armature/ball assembly disposed within the valve body. A retention member can be used to retain the power group assembly to the valve group assembly.

Abstract: Method and arrangement for manufacturing an outlet nozzle (10) for use in a liquid fuel rocket engine. The nozzle forms a body of revolution having an axis of revolution and a cross section that varies in diameter along said axis. The nozzle has a wall structure comprising a plurality of mutually adjacent cooling channels extending substantially in parallel from the inlet end (12) of the nozzle to its outlet end (13). The method includes providing a plurality of preprocessed profile members, each having a web and flanges in opposite directions from said web. Each profile member is machined to present a longitudinally gradually tapering width. The member is curved to conform with the wall section of the nozzle, and members are joined by welding the flanges to form a bell-shaped nozzle structure with cooling channels (11) formed by adjacent webs and adjacent pairs of flanges.

Abstract: A method enables replacement of a gas turbine engine combustor liner. The combustor has a combustion zone formed by inner and outer liners. A plurality of cooling features are formed by overhanging portions of the inner and outer liner liners. The method comprises cutting through at least one of the combustor inner and outer liner aft of an overhanging portion, wherein the portion of the combustor liner upstream from the cut remains coupled within the combustor, removing the portion of the combustor liner that is aft of the cut, and installing a replacement liner within the combustor such that the replacement liner extends aftward from the portion of the combustor liner that is upstream from the cut.

Abstract: A method facilitates replacing a portion of a gas turbine engine combustor liner. The combustor includes a combustion zone formed by inner and outer liners. The inner and outer liners each include a series of panels, and at least one of the inner and the outer liner includes an aft support coupled to an aft end of liner. The method comprises cutting through at least one of the combustor inner and the outer liner upstream from the liner aft support, removing the combustor liner aft support from the combustor, and installing a replacement aft support within the combustor such that the liner aft support extends aft ward from the portion of the combustor liner that is upstream from the cut.

Abstract: Halves of a torus, bridge ring and nozzle are fitted up on tapered fits and clamped one to the other. Weld preps are performed on the torus and nozzle surfaces to be welded. Weld material is applied along the inner and outer radii, initially in a root pass to consume the metal, in subsequent intermediate passes to build up the base metal and finally in a rapid deposition of weld material, e.g., using submerged arc process. The nozzle box assembly half is then stress-relieved, X-rayed and machine-finished.

Abstract: A method for assembling a flap system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates extending a useful life of the exhaust nozzle. The method includes providing a flap basesheet having a width defined between a pair of side edges that are coupled together by a leading edge and a trailing edge, and including at least one stiffener that extends between the basesheet side edges and includes an intermediate portion that has a width that is smaller than that of the basesheet and is at least one of bonded to and formed integrally with the basesheet, and coupling the basesheet to the gas turbine engine with a backbone assembly.

Abstract: A thruster device and associated methods are provided. The thruster device comprises a frustoconical inner layer and a frustoconical outer layer that surrounds the inner layer such that the inner and outer layers cooperate to define a spiral flow passage therebetween. An outlet, such as a nozzle, is attached to an exit end of the thruster device such that the outlet is in fluid communication with the spiral flow passage. In one embodiment, the thruster device includes an end flow cap that is also attached to the exit end and establishes the fluid communication between the spiral flow passage and the outlet. The end flow cap includes at least one channel for directing the propellant from the spiral flow passage to the outlet. The design of the thruster device allows the frustoconical inner and outer layers to be assembled without galling or damaging the spiral flow passage.

Abstract: A method for assembling a flap and seal system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates attaching a basesheet to a backbone. The method includes attaching an attachment system including at least one strap to a basesheet, and coupling the basesheet to a backbone using the attachment system strap.

Abstract: A method of assembly for a gas turbine engine and nacelle includes the steps of aligning the nacelle assembly and the engine substantially parallel to the rotational axis of the engine and translating the nacelle assembly along the axis to engage a first and a second attachment and securing the first attachment.

Abstract: A thrust reverser cascade flow directing element (28) for a gas turbine engine (10) comprises a plurality of plates (30) and a plurality of flow directing vanes (40). At least one vane (40) extends between each pair of plates (30), the plurality of plates (30) are formed from sheet material and a single integral piece of ductile sheet material (32) forms the plurality of flow directing vanes (40). The single integral piece of ductile sheet material (32) has a plurality of longitudinally spaced apart apertures (36) and a plurality of longitudinally spaced apart sheet material portions (38) between the apertures (36) and the ductile sheet material (32) is bent at a plurality of longitudinally spaced positions such that each sheet material portion (38) defines one of the plurality of flow directing vanes (40).

Abstract: Thrust chamber housing for a driving mechanism for space travel applications which is fastened to an injection head with its first end and which includes a combustion chamber housing of a highly heat-resistant steel, a nozzle element of a platinum-iridium alloy, and an expansion nozzle housing of a highly heat-resistant steel, whereby the combustion chamber housing is welded through a first intermediate ring with the nozzle element and the nozzle element is welded through a second intermediate ring with the expansion nozzle housing by means of welded connections, whereby the first and the second intermediate ring are made of a platinum-rhodium alloy.

Abstract: A thruster device and associated methods are provided. The thruster device comprises a frustoconical inner layer and a frustoconical outer layer that surrounds the inner layer such that the inner and outer layers cooperate to define a spiral flow passage therebetween. An outlet, such as a nozzle, is attached to an exit end of the thruster device such that the outlet is in fluid communication with the spiral flow passage. In one embodiment, the thruster device includes an end flow cap that is also attached to the exit end and establishes the fluid communication between the spiral flow passage and the outlet. The end flow cap includes at least one channel for directing the propellant from the spiral flow passage to the outlet. The design of the thruster device allows the frustoconical inner and outer layers to be assembled without galling or damaging the spiral flow passage.

Abstract: A combustion apparatus including a structural jacket, a forward liner, and an aft liner, and an associated method of constructing the same are provided. The structural jacket defines a passage including a first end, a second end, and a neck. The neck is positioned between and separates the first and second ends of the passage. The forward liner is positioned within the first end of the passage and has a throat fitted within the neck of the passage. The aft liner is positioned within the second end of the passage and has an upstream potion abutting the throat of the forward liner. The forward and aft liners are brazed together, and are bonded to the passage of the structural jacket, to form a longitudinal combustion chamber so that combustion of the propellant is contained and directed through the combustion chamber from the forward liner to the aft liner.

Abstract: A method for assembling a gas turbine engine combustor facilitates reducing costs and time required for assembly. The combustor includes a spectacle plate, a plurality of swirlers, and a plurality of deflector plates. The method includes coupling an assembly fixture to at least one swirler, coupling the assembly fixture to the spectacle plate such that the swirler is maintained in alignment with respect to the spectacle plate during assembly of the combustor, and attaching the swirler to the spectacle plate.

Abstract: A rocket motor ablative material (e.g., an insulation liner, bulk material, or the like) of this invention is formed from, as a precursor of the carbon reinforcement structure, one or more polyarylamides configured as a suitable reinforcing structure, such as a yarn, flock, and/or felt. Aramid yarns can be prepared by twisting/spinning aramid filaments, and/or by carding and yarn-spinning staple aramid fibers. In particular, the insulation may be used, for example, for a rocket motor nozzle or as a rocket motor heat shield subjected to conditions comparable to those of continuous filament viscose rayon.

Abstract: A rocket engine having a combustion chamber with a chamber inner wall, a throat with a throat inner wall, and a nozzle with a nozzle inner wall is provided. The chamber inner wall is a vacuum plasma sprayed metal, the throat inner wall is a vacuum plasma sprayed metal, and the nozzle inner wall is a vacuum plasma sprayed metal. The porosity of the vacuum plasma sprayed metal varies in an axial direction of the engine.

Abstract: Methods for manufacturing outlet nozzles for use in rocket motors are disclosed. The methods include providing an inner wall, positioning an outer wall around the inner wall, positioning spacers in the space between the inner and outer walls, and laser welding between the spacers and at least one of the inner and outer walls in order to form cooling ducts between the spacers and to provide T-shaped joints between the spacers and at least one of the inner and outer walls, the T-shaped joints including a rounded shape within the cooling ducts.

Abstract: A replacement method facilitates replacing of a portion of a nuggeted combustor liner within a gas turbine engine combustor in a cost-effective and reliable manner. The combustor includes a combustion zone that is defined by an inner and an outer liner. The inner and outer liners each include a series of panels and a plurality of nuggets formed by adjacent panels. The method includes the steps of cutting between an outer surface and an inner surface of at least one liner panel through at least one nugget, removing at least one panel that is adjacent the area of the liner that was cut, and installing a replacement panel into the combustor for each panel that was removed from the combustor.

Abstract: A combustor baffle includes an outer tube with external threads and a heat shield at opposite ends thereof. An inner tube is disposed inside the outer tube in a unitary assembly therewith. The outer tube is retained in a combustor dome by a retention nut, and the inner tube supports an air swirler with a brazed joint therewith. The brazed joint permits sacrifice of the baffle during disassembly for access to the threaded joint for final disassembly without damage to the dome or air swirler.

Abstract: A replacement method facilitates replacing of a portion of a combustor liner within a gas turbine engine combustor in a cost-effective and reliable manner. The combustor includes a combustion zone that is defined by an inner and an outer liner. The inner and outer liners each include a series of panels and a plurality of nuggets formed by adjacent panels. The method includes the steps of cutting between an outer surface and an inner surface of at least one liner panel, removing at least one panel that is adjacent the area of the liner that was cut, and installing a replacement panel into the combustor for each panel that was removed from the combustor.

Abstract: An annular gas turbine engine combustor liner has a single piece annular shell circumscribed about a first axis of revolution. The shell has a hot side and a cold side, an annular dilution hole section, and a plurality of asymmetrical cylindrical dilution holes extending through the section. Pockets are circumferentially interdigitated between the dilution holes and extend into the shell from the cold side such that the pockets form dilution hole bosses between the pockets through which the dilution holes extend. Each of the dilution holes is circumscribed about a second axis of revolution which is not parallel to the first axis of revolution. The liner may further include forward and aft annular cooling nuggets located forward and aft of the dilution hole section and having annular film cooling slots which are open in the axially aftwardly direction.

Abstract: A regeneratively coolable nozzle for a rocket engine is formed by winding a tubular material (6) onto a winding mandrel to produce a helical composite structure (5), through which a coolant can flow during operation of the nozzle. The spiral composite structure (5) is fixed temporarily and the fixed composite structure (5) is removed from the winding mandrel. Then an inner heat resistant, load bearing layer (9) is applied to the radially inner hot-gas facing surface (H) of the fixed composite structure (5). The present nozzle is lightweight and withstands a high coolant pressure.

Abstract: A first gap formation member and a fixed portion are provided on an element substrate, a movable member is formed on the first gap formation member and the fixing member, and a second gap formation member is formed thereon. The first gap formation member is removed, a wall material is coated and exposed at a pattern mask. The wall material is patterned to form the liquid flow path walls and the liquid supply ports altogether, and removing the second gap formation member, hence making it easier to form the side stopper that supports the movable member stably in a state where the displacement of the movable member is regulated to close the liquid supply port, as well as the minute gap between the movable member and the side stopper in higher precision.

Abstract: A combustor liner has a stepped combustor liner surface defining a combustion zone and an overhang portion forming an air cooling slot. A layer of thermal barrier material is applied to the combustor liner such that at least one portion of the combustor liner receives a layer of thermal barrier material with a thickness greater than 0.01 inches. Thus, the combustor liner absorbs less heat, and the combustor may operate at higher temperatures. As a result, low cycle fatigue and thermal creep are reduced within the combustor and the life cycle for the combustor is extended.

Abstract: A head part of an annular combustor for a gas turbine includes an annular bulkhead at the upstream end of the combustor. The bulkhead includes a first annular section integrally, inwardly and radially extending from the outer annular wall of the combustor, and a second annular section integrally, outwardly and radially extending from the annular inner wall of the combustor. The first and second annular sections are overlapped adjacent to the annular inner wall and secured together by segmented heat shields which have integrated threaded studs engaging with self-locking nuts, and cover the downstream side of bulkhead. The first annular section has a plurality of apertures circumferentially spaced apart from one another to receive fuel burners. Each burner is radial-displaceably positioned in the aperture and sealed by a burner collar, which is clamped between the heat shield and the bulkhead wall.

Abstract: A replacement method facilitates replacing of a portion of a nuggeted combustor liner within a gas turbine engine combustor in a cost-effective and reliable manner. The combustor includes a combustion zone that is defined by an inner and an outer liner. The inner and outer liners each include a series of panels and a plurality of nuggets formed by adjacent panels. The method includes the steps of cutting between an outer surface and an inner surface of at least one liner panel through at least one nugget, removing at least one panel that is adjacent the area of the liner that was cut, and installing a replacement panel into the combustor for each panel that was removed from the combustor.

Abstract: A replacement method facilitates replacing of a portion of a combustor liner within a gas turbine engine combustor in a cost-effective and reliable manner. The combustor includes a combustion zone that is defined by an inner and an outer liner. The inner and outer liners each include a series of panels and a plurality of nuggets formed by adjacent panels. The method includes the steps of cutting between an outer surface and an inner surface of at least one liner panel, removing at least one panel that is adjacent the area of the liner that was cut, and installing a replacement panel into the combustor for each panel that was removed from the combustor.

Abstract: In order to provide a miniaturized ink jet head having a piezoelectric actuator 21 by which ink in a pressure chamber 3 is emitted and to improve its productivity and reliability, a vibration plate 22 is made up of two layers having different Young's moduli, i.e., a layer 27 having a smaller Young's modulus and a layer 28 having a greater Young's modulus. Further, the Young's modulus of each of the layers 27 and 28 is set at values ranging from 50 GPa to 350 GPa and the total thickness of the vibration plate 22 is set at values ranging from 1 &mgr;m to 7 &mgr;m.

Abstract: A thruster device and associated methods are provided. The thruster device comprises a frustoconical inner layer and a frustoconical outer layer that surrounds the inner layer such that the inner and outer layers cooperate to define a spiral flow passage therebetween. An outlet, such as a nozzle, is attached to an exit end of the thruster device such that the outlet is in fluid communication with the spiral flow passage. In one embodiment, the thruster device includes an end flow cap that is also attached to the exit end and establishes the fluid communication between the spiral flow passage and the outlet. The end flow cap includes at least one channel for directing the propellant from the spiral flow passage to the outlet. The design of the thruster device allows the frustoconical inner and outer layers to be assembled without galling or damaging the spiral flow passage.

Abstract: An inkjet print head includes a nozzle, and an ink pool temporarily holding a printing ink and a first screen. The first screen has at least a convex bar column, the bar column having a plurality of bars arranged in a longitudinal direction of the ink pool and spaced apart from each other. A piezo element pumps the printing ink from the ink pool to the nozzle.