Abstract: The invention relates to a gas turbine seal, comprising a metallic component (1) with a durable or erosion-resistant ceramic coating (2) and an abradable ceramic layer (18) which is arranged thereon in locally delimited fashion, a bond layer (5) being arranged on the surface (4) of the metallic component (1), to which bond layer the durable or erosion-resistant ceramic coating (2) is applied as a covering layer. The invention is characterized in that the bond layer (5) comprises separate, adjacent spherical rivets (6) or mushroom-shaped rivets (6) which have a web (7) and a head (8). These rivets form individual metal islands with numerous undercuts (12), around which there is a continuous ceramic network. It is possible to produce very thick layers which have a good bond strength, the ceramic not being flaked away even in the event of introduction of radial/tangential forces, sickle-shaped contact or locally high overheating/frictional heat, and a good sealing action being achieved.

Abstract: A tip treatment bar (16) for a gas turbine engine has a damping coating (24) applied to one or more sides. The coating (24) may be a hard ceramic, for example Magnesia Alumina Spinel, and this may be plasma sprayed directly on to the bar (16) or applied to a substrate (26) which may then be applied to the bar (16).

Abstract: A sealing element (30) for a turbine of a gas turbine engine includes a radially inner surface region provided with an integrally formed seal structure (38) comprising a plurality of radially inwardly projecting walls (40). The walls may be abradable and may define cells (44) for receiving an abradable sealing material.

Abstract: A compressor casing with a volute-shaped flow duct is fastened by means of a rigid fixing arrangement to the bearing casing of the turbomachine, with simple and low-cost means in such a way that the emergence of fragments of a burst compressor impeller from the compressor casing can be prevented. The compressor casing has an outer spiral casing, which surrounds a duct section of the flow duct deflected towards the outside into the radial direction (B), and an inner casing insert piece, which is provided in the radial direction (B) between the spiral casing and the compressor impeller and whose inner contour, together with the outer contour of the hub of the compressor impeller, forms the duct section of the flow duct extending essentially in the axial direction (A).

Abstract: The coating of abradable material presents cavities opening out into the surface of the coating, the cavities extending over a fraction of the thickness of the coating and being defined by walls which form a plurality of continuous paths at the inside surface of the annular portion between the axial end faces thereof, and which form an angle of not less than 5° relative to the general direction of end portions of blades that might come into contact with the abradable material.

Abstract: The invention relates to a gas turbine seal, comprising a metallic component (1) with a durable or erosion-resistant ceramic coating (2) and an abradable ceramic layer (18) which is arranged thereon in locally delimited fashion, a bond layer (5) being arranged on the surface (4) of the metallic component (1), to which bond layer the durable or erosion-resistant ceramic coating (2) is applied as a covering layer. The invention is characterized in that the bond layer (5) comprises separate, adjacent spherical rivets (6) or mushroom-shaped rivets (6) which have a web (7) and a head (8). These rivets form individual metal islands with numerous undercuts (12), around which there is a continuous ceramic network. It is possible to produce very thick layers which have a good bond strength, the ceramic not being flaked away even in the event of introduction of radial/tangential forces, sickle-shaped contact or locally high overheating/frictional heat, and a good sealing action being achieved.

Abstract: The present invention provides for a squealer tip to include some proportion of a highly oxidation-resistant material, and a method for casting same, such that if any environmental coating were removed, the tip would retain some increased level of environmental resistance. The oxidation-resistant material optionally may also be a high abrasion resistance material, such that recession of the tip due to rubbing against a stator would be reduced. In a preferred embodiment, an abrasion-resistant and/or oxidation-resistant material is placed and suitably anchored into the tip region of a wax precursor used to cast a turbine airfoil. During the casting operation, the abrasion-resistant and/or oxidation-resistant material is not completely melted. As the alloy used to form the majority of the turbine blade solidifies, the abrasion and/or oxidation resistant material is incorporated into the turbine airfoil by the solidification of the alloy around it.

Abstract: The present invention provides for a method to reduce the strength of the honeycomb of a jet turbine stator, increasing its machinability, with a resultant reduction in measured peak tooth temperature, while maintaining or even improving its high temperature capability, so as not to limit its operating environment. The air seal functionality is unaffected, and even improved in some instances. The machinability of the honeycomb is increased by using a light element diffused into the honeycomb ribbon to produce the effect of reducing its strength and ductility while maintaining the environmental resistance needed. The present invention also includes the stator honeycomb produced by the foregoing method.

Abstract: A joint between a turbine ring. The ring is jointed to a spacer of the turbine by a joint including hooks which fit together on one side, and an abutment of flanges as well as a mortise and tenon joint on the other side. According to the joint, the tenon and mortise are separated from the flanges in abutment to limit the plays in the axial direction, produced by design or by expansions, to increase mechanical strength and simplify manufacturing.

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor (32) carrying a plurality of radially extending fan blades (34). A fan blade containment assembly (38) surrounds the fan blades (34) and the fan blade containment assembly (34) comprises a generally annular, or frustoconical, cross-section casing (52). At least one corrugated sheet metal ring surrounds the casing (52) wherein the corrugations extend with axial and/or circumferential components.

Abstract: A honeycomb structure is arranged on a stator of a gas turbine and is arranged opposite the moving blades of the gas turbine. The honeycomb structure is composed of metal sheets in such a way that webs having double the sheet thickness are obtained and connections of single thickness are obtained between the webs. The angle &agr; between the circumferential direction of the stator and the webs of double thickness is 5° to 175°. This serves to avoid undesirable efforts such as, for example, overheating during friction at the honeycomb structure of the moving blades of the gas turbine.

Abstract: The present invention provides for a squealer tip to include some proportion of a highly oxidation-resistant material, and a method for casting same, such that if any environmental coating were removed, the tip would retain some increased level of environmental resistance. The oxidation-resistant material optionally may also be a high abrasion resistance material, such that recession of the tip due to rubbing against a stator would be reduced. In a preferred embodiment, an abrasion-resistant and/or oxidation-resistant material is placed and suitably anchored into the tip region of a wax precursor used to cast a turbine airfoil. During the casting operation, the abrasion-resistant and/or oxidation-resistant material is not completely melted. As the alloy used to form the majority of the turbine blade solidifies, the abrasion and/or oxidation resistant material is incorporated into the turbine airfoil by the solidification of the alloy around it.

Abstract: A rotatory pump has a housing (20) and an impeller wheel (40) which is mounted on a driving shaft (30) to rotate integrally therewith. The driving shaft is rotatably supported in the housing. The driving shaft also has a disk (44) disposed to be concentric with the driving shaft. Radially extending blades (46) are directed along the axial direction of the driving shaft (30) being provided on the disk, and the blades, together with the inner wall portions (22a) of the housing (20) which face the blades (46), form flow channels for the fluid to be pumped. At least one raised portion (48) is on each of the radially extending edges (46a) of preferably three blades (46), to bear against the inner wall portions (22a) of the housing (20) which face the blades (46) of the impeller wheel housing.

Abstract: The present invention provides, in one embodiment, an annular turbine seal for disposition in a turbine between a rotatable component having an axis of rotation and a turbine housing about the same axis of rotation. The turbine seal has a plurality of arcuate seal carrier segments that have an abradable portion secured to the seal carrier segments. In addition, at least one spring is disposed on the seal carrier segment to exert a force and maintain the seal carrier segment adjacent to the rotatable component.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) which has an inner surface (62), an outer surface (64), an upstream flange (42) and a downstream flange (52). The outer surface (62) of the metal casing (40) is provided with a pattern of blind apertures (66) which extend radially inwardly from the outer surface (62) of the metal casing (40) between the flanges (42,52). The fan blade containment has reduced weight of the metal casing (40) of the fan blade containment assembly (38) for a large diameter turbofan gas turbine engine (10) but has unimpaired stiffness and penetration resistance.

Abstract: A gas turbine engine abradeable seal system is provided comprising a seal assembly and a cooperating interacting turbine blade. The turbine blade has a tip portion containing cubic boron nitride abrasive particles and the seal assembly has a superalloy substrate with a bond coat thereon having a surface roughness of at least 300 RA and a porous ceramic abradeable seal material on the bond coat having a porosity of from 5 to 15 volume %.

Abstract: A gas turbine engine abradeable seal system is provided comprising a seal assembly and a cooperating interacting turbine blade. The turbine blade has a tip portion containing cubic boron nitride abrasive particles and the seal assembly has a superalloy substrate with a bond coat thereon having a surface roughness of at least 300 RA and a porous ceramic abradeable seal material on the bond coat having a porosity of from 5 to 15 volume %.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a blade containment portion (54) and a transition portion (58). The upstream portion (56) has a flange (52) connecting the metal containment casing (40) to an axially adjacent casing. The upstream portion (56) of the casing (40) has a greater diameter (D1) than the diameter (D2) of the blade containment portion (54) and the blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54). The transition portion (58) connects the blade containment portion (54) and the upstream portion (56) to transmit loads from the blade containment portion (54) to the upstream flange (42).

Abstract: The present invention provides, in one embodiment, an annular turbine seal for disposition in a turbine between a rotatable component having an axis of rotation and a turbine housing about the same axis of rotation. The turbine seal has a plurality of arcuate seal carrier segments that have an abradable portion secured to the seal carrier segments. In addition, at least one spring is disposed on the seal carrier segment to exert a force and maintain the seal carrier segment adjacent to the rotatable component.

Abstract: Self-supporting wear rings (1a, 1b) for fluid flow machines are arranged at a radial sealing gap (18) to a rotor (9), and have an outer jacket surface (3) provided with a wear surface (2). The wear rings (1a, 1b), which are fixed in the axial direction, consist of a material with lower thermal expansion coefficients than their support and lie hollowly on their inner jacket surface, with a first conical holder surface (8) encountering with the apex of its cone a second holder surface (8″), which is designed as a straight shoulder, in the plane of the latter (19) on the axis of rotation or encountering a second holder surface, which is designed as a conical holder surface (8′), at its cone apex on the axis of rotation (20).

Abstract: An improved gas turbine blade is disclosed that utilizes a blade tip shroud with tapered shroud pockets to remove excess weight from the shroud while not compromising shroud bending stresses. Excess shroud weight removed by the tapered pockets reduces overall blade pull while maintaining material necessary to increase shroud stiffness. The resulting rib created between the tapered pockets provides a surface, away from possible areas of stress concentration, for drilling radial cooling holes.

Abstract: An abradable turbine seal having a novel honeycomb cellular structure fabricated from metal foils or sheets showing good manufacturability, optimized brazeability and especially good structural integrity and oxidation resistance after brazing to metal support structures. MCrAlY (M=Ni, Fe, Co or combinations thereof) foil and sheet metals are particularly suitable to produce such a structure.

Abstract: A turbomachine for moving air comprising includes a shroud disposed about a longitudinal axis and a rotor assembly mounted for rotation about the longitudinal axis. The rotor assembly has a plurality of blades and tips of the blades are coupled to an annular band. The annular band is disposed with respect to the shroud so as to define a gap extending continuously between an outer surface of the annular band and an inner surface of the shroud. A seal structure extends from the inner surface of the shroud and into the gap. The seal structure has a density sufficient to reduce swirl in recirculating airflow and to minimize air leakage across the gap.

Abstract: A unitary interstage sealing ring for bridging and sealing the space between the blade shanks of adjacent turbine rotors in a multi-stage turbine. The sealing ring serves to confine relatively cool air conveyed to the turbine so the cooling air flows into the turbine blades for maintaining the blades at a desired temperature. The sealing ring includes a disk-shaped body member that includes a pair of axially-spaced, radially-extending arms that terminate in respective seal teeth for engagement with a stationary annular sealing member. Each of the radial arms includes an outwardly-extending axial arm that terminates in a disk engagement member for sealing engagement with the face of an adjacent turbine disk to confine cooling air within the space between the adjacent turbine disks. Additionally, one of the disk engagement members includes at least one, projection that contacts a respective turbine rotor to limit relative rotation between the sealing ring and the turbine.

Abstract: A curved section (12) is provided at a proximal end of a fan (10), so that a rotation locus thereof is configured in a shape conforming to a flow of a cooling air such as center-dented. Accordingly, the cooling air can be smoothly sucked in by the fan (10) as well as fan shroud along an arcuate section on an inlet side. Further, since natural flow toward radially outward flow of the cooling air can be guided by an arcuate section on an outlet side, thus smoothly discharging the cooling air. Therefore, the flow of the cooling air can be made smooth entirely from upstream to downstream, thus increasing flow rate and reducing noise.

Abstract: In a fluid machine such as multistage centrifugal compressor, the pressure of a working fluid is raised with the rotation of a impeller. A shaft sealing device is mounted at the end of the rotary shaft in order to prevent the working gas of the raised pressure from leaking out of the fluid machine and then polluting the surroundings. In a case where the fluid machine operates to rotate at high speed, a labyrinth sealing device is employed for the purpose of reducing a frictional loss in the shaft sealing device. The labyrinth sealing device includes a plurality of stages of labyrinth fins which are formed on a casing or the rotary shaft, and a layer of abradable coating to-be-bitten which is formed in the surface part of the rotary shaft or the casing opposing to the labyrinth fins. The clearances between the labyrinth fins and the rotary shaft or the casing are made narrower on the low-pressure side of the labyrinth sealing device than on the high-pressure side thereof.

Abstract: A gas turbine engine includes a non-rotatable member that includes a honeycomb seal that reduces wear to rotor seal teeth disposed within the gas turbine engine. The gas turbine engine also includes a rotatable annular member including a sealing assembly disposed between rotor and stator components. The rotatable annular member includes seal teeth that extend radially outward from the rotatable annular member. The stator components include a non-rotatable member that includes a honeycomb seal that extends radially inward. The honeycomb seal is fabricated from a material that has a melting temperature less than approximately 2000° F.

Abstract: A stator assembly having an inner shroud assembly formed of an inner shroud and rubstrip is disclosed. Various construction details and features are developed which relate to durability and manufacturing feasibility are developed. In one detailed embodiment, at the rubstrip is bonded to the thermoplastic shroud. The method is applicable to joining an elastomer, such as silicone rubber material, to a thermoplastic substrate.

Abstract: A material useful for forming high temperature coatings. The material contains a MCrAlY powder wherein M is selected from the group consisting of iron, nickel, cobalt, iron-base alloys, nickel-base alloys and cobalt-base alloys. An aluminum layer coats the powder. The method forms a high temperature coating with the powder. Thermally spraying the MCrAlY powder towards a substrate reacts the aluminum coating with the MCrAlY powder to metallurgically bond the MCrAlY powder and coat the substrate.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprising a generally cylindrical, or frustoconical, first metal casing (40) and a plurality of generally cylindrical, or frustoconical axially spaced second metal casing is (54, 56 & 58). The second metal casings (54, 56, 58) surrounded and abut the first metal casing (40). The first metal casing (40) is relatively hard and tough to contain a fan blade (34). The second metal casings (54, 56, 58 are relatively soft and lightweight to stiffen the first metal casing (40) and fan blade containment regions (A, B & C). The second metal casing (54, 56, 58) are arranged coaxially with the first metal casing (40). One of the second metal casing means (56) is in a plane (Y) containing the fan blades (34) of the gas turbine engine (10). The first metal casing (40) comprises steel and the second metal casing is (54, 56, 58) comprise aluminum. The fan blade containment assembly (38) is lighter for large diameter turbofan gas turbine engine (10).

Abstract: A casing for the fan section of an aircraft gas turbine engine of the turbofan type. The casing includes an inner filler material and is defined by an outer shell and an inner liner that is spaced radially inwardly from the outer shell. The inner liner is held in position relative to the outer shell by a plurality of radially-inwardly-extending support pins. The pins are carried by the outer shell and are slidably received in and extend at least partially into the inner liner. The support pins retain the inner liner against axial and rotational movement relative to the outer shell, while allowing relative radial movement therebetween in response to thermal expansion and contraction, as well as in response to aircraft maneuver loads imposed on the engine.

Abstract: An air seal for use in a gas turbine engine is disclosed. The seal includes a seal substrate, and an abradable layer on the substrate, with the abradable including a thermoset polymer and a thermoplastic polymer. The abradable may also include a filler to add porosity, or provide lubrication, to enhance abradability.

Abstract: A low pressure turbine shroud includes a backplate and abradable material attached to said backplate. The backplate includes an upstream end, a downstream end, and a body extending between the upstream end and the downstream end. The body includes a first portion and a second portion. The body first portion has a first thickness greater than a body second portion second thickness. The abradable material attaches to an inner surface of the backplate and includes at least one segment having a thickness greater than remaining segments of abradable material.

Abstract: A rotary machine includes a rotatable rotor, blades radially mounted on and spaced apart about the rotor, a stationary casing disposed about and radially outward from the rotor and blades defining an annular gap between the casing and tips of the blades, and a flexible abradable seal assembly attached on the casing and disposed within the annular gap such that during differential growth of the rotor and blades relative to the casing the seal assembly flexes and abrades in response to contact by the blade tips and thereby clearance between the casing and blade tips is controlled as to avoid contact by the moving blade tips with the casing.

Abstract: A casing structure of metal construction for the rotor blade area of axial flow compressor and turbine stages includes a closed, stable outside wall, a segmented inside wall that is interrupted by multiple expansion joints, and a load-transmitting connecting structure between inside and outside wall. A multiple sub-divided hollow chamber structure with a plurality of thin wall elements standing upright on the inside wall and outside wall, at least the majority of which are directly connected together, is arranged as connecting structure. The hollow chamber structure is connected to the inside wall and/or the outside wall by brazing.

Abstract: A sealing element (30) for a turbine of a gas turbine engine includes a radially inner surface region provided with an integrally formed seal structure (38) comprising a plurality of radially inwardly projecting walls (40). The walls may be abradable and may define cells (44) for receiving an abradable sealing material.

Abstract: In a fluid machine such as multistage centrifugal compressor, the pressure of a working fluid is raised with the rotation of a impeller. A shaft sealing device is mounted at the end of the rotary shaft in order to prevent the working gas of the raised pressure from leaking out of the fluid machine and then polluting the surroundings. In a case where the fluid machine operates to rotate at high speed, a labyrinth sealing device is employed for the purpose of reducing a frictional loss in the shaft sealing device. The labyrinth sealing device includes a plurality of stages of labyrinth fins which are formed on a casing or the rotary shaft, and a layer of abradable coating to-be-bitten which is formed in the surface part of the rotary shaft or the casing opposing to the labyrinth fins. The clearances between the labyrinth fins and the rotary shaft or the casing are made narrower on the low-pressure side of the labyrinth sealing device than on the high-pressure side thereof.

Abstract: An improved containment structure for use with turbomachinery, such as turbofan engines. A casing is provided with at least one impact resistant section of a thickness sufficient to resist being pierced upon impact by a blade. The impact resistant section has a smoothly contoured thickness along its axial length and is free from features likely to cause a local failure in the event of impact by a blade released from a rotating disk.

Abstract: The cast part (1) comprises a basic body (2) of a first material and a coating of at least one second material. The second material is applied to a co-cast skeleton structure (3) on the surface (20) of the basic body, with it forming a covering (4) or a covering layer (42).

Abstract: A centrifugal compressor (1) including a single rotating shaft (5), a first (upstream) and second (downstream) impellers (2, 3) mounted on ends of the rotating shaft (5) respectively, and a casing (4) for accommodating these impellers. An air path (35) is provided for introducing an air pressurized by the upstream impeller (2) to the downstream impeller (3). An abradable layer (34) is embedded in the casing inner wall (4) such that it faces and is cut by the second downstream impeller (3). The expensive abradable layer (34) is only provided for the downstream impeller (3) since the effect of the abradable layer (34) is significant when provided for the downstream impeller.

Abstract: There is provided a hardwall fan case for encasing a forward fan in a gas turbine engine. The fan case has a stiff annular shell spaced radially outward from the tips of the fan blades and a flexible ring which is an integral structural part of the shell. The flexible ring has a frusto-conical shape with a lip adjacent to the blade trailing edge. The ring extends axially rearwardly from the fixed root to the free inner edge forming a cantilevering resilient ring. A hollow cavity defined between an inner surface of the shell and an outer surface of the flexible ring provides clearance for the flexible ring to deform radially outwardly on impact with a released blade, or to elastically flex on contact with the trailing edge blade tip during bird strike events.

Abstract: A turbine airfoil includes pressure and suction sidewalls extending between leading and trailing edges and from root to tip. The tip includes a floor bounding an internal cooling channel within the airfoil which channels cooling air. The airfoil tip includes a first rib adjacent the pressure sidewall, a second rib spaced therefrom to define a first slot, and a third rib adjacent the suction sidewall to define a second slot with the second rib. The tip floor includes feed holes extending between the cooling channel and the first slot for supplying cooling air therein for discharge over the second rib towards the third rib.

Abstract: A turbine blade squealer tip includes an airfoil shaped tip cap having a squealer tip wall extending radially outwardly from and around the perimeter of the airfoil shaped tip cap to define a radially outwardly open tip cavity. The tip wall has an inboard side facing the interior of the cavity and an outboard side facing away from the cavity and the tip cap has an outer tip side on a bottom of the cavity. Thermal barrier coatings are disposed on the inboard and outboard sides of the squealer tip wall and on the outer tip side of the tip cap. One embodiment provides the tip cap with cooling holes disposed therethrough to flow cooling air into the cavity.

Abstract: An improved rotary engine, engine lining and engine casing and lining combination are disclosed. The engine is preferable a turbofan engine and includes a fan blade, rotatably mounted within a generally cylindrical casing for rotation about a lengthwise extending central axis. An annular lining is mounted within the casing between a tip of the fan blade and the casing. The lining seals the tip of the fan blade within the casing. The lining includes a brush seal that extends around an inner circumference of the casing, and has a plurality of bristles that extend radially inward from the seal. A retaining membrane extends around the brush seal, preventing the bristles from extending substantially in a radial direction from the brush seal. The retaining membrane is adapted to release the bristles to extend radially inward to occupy a region between the tip of the fan blade and the casing upon a radial excursion of the fan blade.

Abstract: A compressor blade that has a blade root, an airfoil having a first end, and a second end opposite the first end, the second end having at least one edge, and the airfoil is made of a first material having a first modulus of elasticity. A blade platform connects the blade root to the first end of the airfoil, and a flexible seal is connected to the airfoil adjacent the second end, and the seal is made of a second material having a modulus of elasticity that is substantially less than the first modulus of elasticity.

Abstract: A blade containment structure for turbomachinery, such as a high bypass turbofan engine. The blade containment structure includes a first casing member having a wall that defines an inner containment shell that immediately surrounds the blades of the turbomachine, and a second casing member assembled to the first casing member. The second casing member has a wall that defines an outer containment shell that surrounds the inner containment shell, such that a cavity is defined by and between the inner and outer containment shells. The first casing member is formed from a high-toughness material such as steel or a nickel-base alloy, while the second casing member is formed from a lower-weight and potentially lower-toughness material as compared to the material of the first casing member.

Abstract: A rotary machine (10) having rotor elements (54-66), (114) and seal lands (86, 112) is disclosed. Various construction details are developed for facilitating assembly of the machine. In one embodiment clearances are established between rotor blades (66) and the seal land (86) which enable assembly of the components and provide for acceptable clearances at a steady-state operative condition of the engine.

Abstract: A gas turbine engine seal system includes a rotating member having an abrasive tip disposed in rub relationship to a stationary, abradable seal surface. The abrasive tip comprises a zirconium oxide abrasive coat having a columnar structure that is harder than the abradable seal surface such that the abrasive tip can cut the abradable seal surface.

Abstract: A containment casing for a gas turbine engine comprising a substantially rigid casing shell arranged in operation coaxially with an axis of rotation of the gas turbine engine. The casing extending circumferentially around an array of fan rotor blades which are arranged to rotate about the engine axis. In the region of the casing where it is predicted blade impact in the event of a fan blade failure, there is at least one reinforcing rib extending substantially radially from the casing and circumscribing the outer periphery of the casing shell. The ribs providing improved strengthening of the casing. Preferably a first rib is positioned about the casing between the operational position of the trailing edge and the mid chord point of the fan blade. A second rib is axially positioned between the operational position of the mid chord point and the leading edge of the fan blade. A third rib may also be positioned axially forward of the operational position of the leading edge of the fan blade.