37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Repair, construction and servicing of gas turbine engines
and their component parts, excluding use in the automotive
industry. Custom manufacturing of component parts for gas turbine
engines, excluding use in the automotive industry. Engineering services in the field of gas turbine engines for
the commercial aviation, military, energy, business and
general aviation sectors, excluding use in the automotive
industry.
2.
Additively Manufactured Tooling for Coating Components Using Electrically Driven Aqueous Processes
A tool for securing a component for a treatment process and methods of making and using same. A method for securing a component for a treatment process includes obtaining a tool for retaining the component during the treatment process. The tool includes a cantilevered plate. The method includes associating the tool with a support bar and securing the component to the tool such that the cantilevered plate is in contact with the component.
Lost wax casting bonding methods and systems. A method for bonding a wax component in a feeder system for a lost wax casting process comprises embedding conductive nano-particles in wax to form a sacrificial susceptor. The method includes coupling the wax component to another component such that the sacrificial susceptor is at an interface of the wax component and the other component. The method comprises inductively heating the sacrificial susceptor to cause the wax of the sacrificial susceptor and at least a portion of wax of the wax component to melt to thereby form a bond between the wax component and the other component.
37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
(1) Repair, construction and servicing of gas turbine engines and their component parts, excluding use in the automotive industry.
(2) Custom manufacturing of component parts for gas turbine engines, excluding use in the automotive industry.
(3) Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry.
5.
Fastening System for Securing a Component to a Mounting Device for a High-Temperature Process
A fastening system for securing a mounting tool configured to retain a component that is to undergo a high-temperature process. The mounting tool has a first portion having a first tab with a first slot and a second portion having a second tab with a second slot. The fastening system comprises an outer clip including outer clip arms and a central portion. The outer clip arms have a gap therebetween. The central portion protrudes into the gap. The fastening system includes an inner clip including inner clip arms. The inner clip is pushed into the gap such that each inner clip arm contacts and pushes one outer clip arm towards the first tab and the second tab. When the fastening system is inserted into the first slot and the second slot, the fastening system limits relative movement between the first tab and the second tab.
B25B 11/00 - Porte-pièces ou dispositifs de mise en position non couverts par l'un des groupes , p. ex. porte-pièces magnétiques, porte-pièces utilisant le vide
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry
A system for securing a component for an electrically-driven process is disclosed. The system includes a mai015134n body having an opening, a retention slot configured for retaining the component, and a channel. The system includes a fastener extending through the opening in the main body, and an insert inserted into the channel. The system has a shank movably coupled to the fastener and configured to contact the insert. The insert is configured to maintain electrical contact between the shank and the component when the component is retained within the retention slot.
Lost wax casting bonding methods and systems. A method for bonding a wax component in a feeder system for a lost wax casting process comprises embedding conductive nano-particles in wax to form a sacrificial susceptor. The method includes coupling the wax component to another component such that the sacrificial susceptor is at an interface of the wax component and the other component. The method comprises inductively heating the sacrificial susceptor to cause the wax of the sacrificial susceptor and at least a portion of wax of the wax component to melt to thereby form a bond between the wax component and the other component.
Systems and methods are provided for repairing compressor blades using hybrid manufacturing. A mounting tool for use in repairing a turbine component includes a slot formed in the mounting tool. The slot is configured to receive the turbine component for securing the turbine component to the mounting tool. The mounting tool includes at least one fastener engageable with the turbine component and capable of retaining the turbine component within the slot; and at least a pair of support bases provided on the mounting tool. Each of the at least a pair of support bases is positioned adjacent a respect edge of the turbine component when the turbine component is retained in the mounting tool. Each of the at least a pair of support bases is configured for additively building a sacrificial support on each of the at least a pair of support bases for enabling repair of the turbine component.
B22F 7/06 - Fabrication de couches composites, de pièces ou d'objets à base de poudres métalliques, par frittage avec ou sans compactage de pièces ou objets composés de parties différentes, p. ex. pour former des outils à embouts rapportés
B22F 5/04 - Fabrication de pièces ou d'objets à partir de poudres métalliques caractérisée par la forme particulière du produit à réaliser d'aubes de turbines
B22F 10/40 - Structures destinées à soutenir des pièces ou des articles pendant la fabrication et retirées par la suite
B22F 10/66 - Traitement de pièces ou d'articles après leur formation par des moyens mécaniques
B33Y 40/20 - Posttraitement, p. ex. durcissement, revêtement ou polissage
12.
Method and apparatus for improving core manufacturing for gas turbine components
A system for producing a core including a core profile for use in casting. The system comprises a cavity block including an upper portion, a lower portion, and a recessed cavity in each of the upper portion and the lower portion; an adapter insert including a first portion and a second portion, the adapter insert sized to fit within the recessed cavity in each of the upper portion and the lower portion of the cavity block; and a core die insert sized to fit and be positioned within the adapter insert. The core die insert includes a sacrificial material and a hollow internal profile, with the hollow profile of the core die insert corresponding to the core profile.
A method is provided for repairing a hollow fan blade that includes a first portion and a second portion secured together by a joint. An interior cavity of the hollow fan blade is under a vacuum. The method includes obtaining a replacement portion having a unitary construction, removing a portion from a cutback region of the hollow fan blade to form a weld foundation, and welding the replacement portion to the weld foundation. The welding is at a temperature at the joint that does not exceed a distortion temperature of the joint.
B23K 37/00 - Dispositifs ou procédés auxiliaires non spécialement adaptés à un procédé couvert par un seul des autres groupes principaux de la présente sous-classe
B23K 9/095 - Surveillance ou commande automatique des paramètres de soudage
B23K 9/16 - Soudage ou découpage à l'arc utilisant des gaz de protection
Systems and methods for repairing hollow fan blades. The hollow fan blade has a concave portion and a convex portion secured together by a joint. An interior cavity of the hollow fan blade is charged with a vacuum. The method includes removing material from a damaged portion of the hollow fan blade and welding the damaged portion such that a temperature of the joint does not exceed a distortion temperature.
B23K 37/00 - Dispositifs ou procédés auxiliaires non spécialement adaptés à un procédé couvert par un seul des autres groupes principaux de la présente sous-classe
Systems and methods for repairing compressor blades using hybrid manufacturing techniques. A tool for repairing a compressor blade has a plurality of support bases, and a slot configured for securement of said compressor blade to said mounting tool. The tool has a pair of support bases, each of which is configured for the building of a sacrificial support thereon to stabilize the compressor blade during repair. The tool is coupled to a hybrid manufacturing machinery and the sacrificial supports stabilize the compressor blade during additive and subtractive manufacturing processes.
B22F 7/06 - Fabrication de couches composites, de pièces ou d'objets à base de poudres métalliques, par frittage avec ou sans compactage de pièces ou objets composés de parties différentes, p. ex. pour former des outils à embouts rapportés
B22F 10/66 - Traitement de pièces ou d'articles après leur formation par des moyens mécaniques
B22F 5/04 - Fabrication de pièces ou d'objets à partir de poudres métalliques caractérisée par la forme particulière du produit à réaliser d'aubes de turbines
B33Y 40/20 - Posttraitement, p. ex. durcissement, revêtement ou polissage
B22F 10/40 - Structures destinées à soutenir des pièces ou des articles pendant la fabrication et retirées par la suite
16.
Spherical fused silica compositions for injection molded ceramic cores and methods of making parts using such compositions
A single crystal ceramic core composition has an inorganic portion and an organic portion. The inorganic portion makes up about 85% by weight of the total weight of the ceramic core composition, and the organic portion makes up about 15% by weight of the total weight of the ceramic core composition. The inorganic portion includes about 94 to 98% by weight spherical fused silica, and about 2 to 6% by weight zircon flour. The organic portion includes about 84 to 88% by weight binder, about 1 to 2% by weight dye, about 6 to 12% by weight surfactant, and about 1 to 5% by weight polymeric fiber.
B22C 1/00 - Compositions des matériaux réfractaires pour moules ou noyauxLeur structure granulaireCaractéristiques chimiques ou physiques de la mise en forme ou de la fabrication des moules
B22C 9/10 - NoyauxFabrication ou mise en place des noyaux
C04B 35/14 - Produits céramiques mis en forme, caractérisés par leur compositionCompositions céramiquesTraitement de poudres de composés inorganiques préalablement à la fabrication de produits céramiques à base d'oxydes à base de silice
A method of using a masking device to mask a portion of a gas turbine engine component for an electroplating process is provided. The masking device includes a main body having sidewalls, a removeable coverplate having an end plate and one or more locking tabs, a fastener located at least partially within the main body, and a shank engaged with the fastener. The method includes placing the gas turbine engine component through an opening in the main body such that the component is held within the main body by a retention slot. The method includes sliding the one or more locking tabs of the removeable coverplate through the opening in the main body such that the removeable coverplate covers the opening of the main body. The method includes securing the one or more locking tabs in one or more relief slots in the sidewalls of the main body.
Lost wax casting bonding methods and systems. A method for bonding a wax component in a feeder system for a lost wax casting process comprises embedding conductive nano-particles in wax to form a sacrificial susceptor. The method includes coupling the wax component to another component such that the sacrificial susceptor is at an interface of the wax component and the other component. The method comprises inductively heating the sacrificial susceptor to cause the wax of the sacrificial susceptor and at least a portion of wax of the wax component to melt to thereby form a bond between the wax component and the other component.
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry
The present disclosure provides for improvements in fixturing parts in preparation for an electroplating process, thus reducing part handling. The system provides a reusable masking tool comprising a main body having an opening with a removeable coverplate positioned within the opening and one or more locking tabs engaging a corresponding relief slot in at least one of the sidewalls of the main body. The system also comprises a fastener moveably secured within the main body and extending through a top surface of the main body. The main body, fastener, and coverplate may be fabricated from a polymer material by way of an additive manufacturing process. A shank, fabricated from a conductive material, extends through the fastener and is engaged with the fastener such that upon rotation of the fastener, the shank is drawn into contact with the gas turbine engine component, thus providing a conduit for the electroplating process.
A tool for reforming a core pattern. A method of reforming a core pattern comprises opening a core reformer tool. The core reformer tool has a first portion and a second portion facing the first portion. The first portion includes a concave member and a first adjustable pin. The second portion includes a convex member and a second adjustable pin. The method includes positioning the core pattern in the core reformer tool and closing the core reformer tool. The method comprises adjusting at least one of the first adjustable pin and the second adjustable pin to alter a surface of the core pattern. The method includes directing cooling air through the core reformer tool to solidify the core pattern and opening the core reformer tool to remove the core pattern.
The present disclosure provides a system and method of repairing a flange of an engine case having a damaged section of attachment tabs and/or flange region. One or more damaged attachments tabs extending away from a flange are removed along with a first portion of the flange, thus leaving a second portion of the flange. A replacement component is produced comprising a replacement flange portion and one or more replacement tabs. The replacement flange portion is placed in contact with the second portion of the flange and the replacement flange portion is secured to the second portion of the flange.
A core pattern reformer system for adjusting and setting a core pattern for use in casting a gas turbine blade has a first portion having a first internal face with a concave portion. The first portion is coupled to a second portion. The second portion has a second internal face with a convex portion. A plurality of adjustable pins is positioned along the internal faces of the first portion and the second portion. The pins have a height that is adjustable with respect to the internal faces. A locking mechanism is included for securing the first portion to the second portion. The system includes one or more air inlets and one or more air exits and a source of cooling air coupled to the one or more air inlets.
A nickel-based alloy is disclosed which is suitable for casting gas turbine components having improved strength and comparative lower density while utilizing commercially available heat treatment cycles. The nickel-based alloy is suitable for providing equiaxed, directionally solidified, and single crystal castings. Methods of providing a cast article of the nickel-based alloy and subjecting the article to heat treatment cycles are also disclosed.
C22C 19/05 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de nickel avec du chrome
C22F 1/00 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid
C22F 1/10 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid du nickel ou du cobalt ou de leurs alliages
36.
METHOD AND APPARATUS FOR IMPROVING CORE MANUFACTURING FOR GAS TURBINE COMPONENTS
An apparatus and method for producing a core for use in casting a gas turbine component is provided. The apparatus comprises a system having a cavity block with an upper portion, a lower portion, and a recessed cavity in each of the upper portion and the lower portion. Positioned within the cavity block is an adapter insert, and within the adapter insert is positioned a core die insert, where the core die insert is fabricated from a sacrificial material and has a hollow internal profile corresponding to an external surface of a core. A ceramic-based material is supplied to the core die insert where it solidifies. The core die insert is placed into a water-based solution and the core die insert is removed from around the core.
An apparatus and method for producing a core for use in casting a gas turbine component is provided. The apparatus comprises a system having a cavity block with an upper portion, a lower portion, and a recessed cavity in each of the upper portion and the lower portion. Positioned within the cavity block is an adapter insert, and within the adapter insert is positioned a core die insert, where the core die insert is fabricated from a sacrificial material and has a hollow internal profile corresponding to an external surface of a core. A ceramic-based material is supplied to the core die insert where it solidifies. The core die insert is placed into a water-based solution and the core die insert is removed from around the core.
A guide vane for use in a compressor discharge plenum of a gas turbine engine is disclosed. The guide vane comprises a guide support system, a first side panel, a second side panel, and a deflector panel secured to the mounting system and extending between the first side panel and the second side panel. The deflector panel has a curved profile substantially in accordance with Cartesian coordinate values of X, Y, and Z as set forth in Table 1 where the X, Y, and Z values are in inches from a center point of a bottom surface of the deflector panel. The coordinate values are connected by smooth continuing arcs to form profile sections and the profile sections are joined together smoothly to form the curved profile of the guide vane.
A turbine nozzle having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, and within an envelope of approximately +/- 0.049 inches, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form the airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.
An apparatus and method for reducing the vibrations to a transition duct and improving the assembly and fit-up of a transition duct relative to a combustion system are disclosed. A damping member is provided between the transition duct and its mounting points to the gas turbine engine. The damping member comprises one or more layers of a composite material positioned between layers of sheet metal. An adjustable mounting system is also provided for the transition duct to allow for movement of a transition duct relative to the combustion liner.
An apparatus and method for mounting a combustion liner within a flow sleeve of a gas turbine combustion system is disclosed. A mounting system comprises a plurality of low-profile mounting tabs secured to a combustion liner where each of the mounting tabs are placed within slots of flow sleeve pegs when the combustion liner is installed in a flow sleeve. A plurality of liner stop brackets are removably secured to a flange of the flow sleeve and have an arm extending to be adjacent to a top contact surface of the mounting tabs. The mounting system reduces blockage to the surrounding airflow.
A system for retaining a cross fire tube in a multi-combustor gas turbine engine is disclosed. The system comprises a flow sleeve having a generally annular body, and a flange at a forward end thereof and having one or more recessed portions. A cross fire tube extends through one or more openings in the flow sleeve and is secured in place by a retention clip. The retention clip includes a plurality of fingers which engage the cross fire tube and a mounting plate engaging the one or more recessed portions of the flow sleeve flange so as to create a clip engagement having a lower profile than prior art configurations.
A system for directing cooling air into a gas turbine combustor is provided. The system comprises a transition duct coupled to a flow sleeve, where air to be used for combustor cooling and in the combustion process enters a bellmouth of the transition duct, passes through a plurality of struts within the bellmouth, and is distributed to a passage located between the combustion liner and flow sleeve.
A heat transfer mechanism is provided comprising a plurality of turbulators located along a surface of a body, such as a combustion liner. The turbulators have a first side with a first ramp angle, a second side with a second ramp angle, a height, and a base width, where the base width is a function of the height and where the turbulators are spaced an axial distance apart that is a function of the turbulator height.
A system and method for cooling a turbine blade tip shroud is provided. The turbine blade comprises a blade attachment, a platform extending radially outward from the attachment, an airfoil extending radially outward from the platform, and a tip shroud extending radially outward from the airfoil. The tip shroud has one or more knife edges extending radially outward from an outer surface of the tip shroud. One or more cooling passages extend through the airfoil and to the tip shroud. The turbine blade also includes one or more tip plates secured at or near the outer surface of the tip shroud thereby forming a plenum between the outer surface and the one or more tip plates. The one or more tip plates also include a plurality of cooling holes for flowing cooling air through the plenum to cool the tip shroud.
B23P 15/02 - Fabrication d'objets déterminés par des opérations non couvertes par une seule autre sous-classe ou un groupe de la présente sous-classe d'aubes de turbine ou d'organes équivalents, en une seule pièce
46.
Diffuser guide vane with deflector panel having curved profile
A guide vane for use in a compressor discharge plenum of a gas turbine engine is disclosed. The guide vane comprises a guide support system, a first side panel, a second side panel, and a deflector panel secured to the mounting system and extending between the first side panel and the second side panel. The deflector panel has a curved profile substantially in accordance with Cartesian coordinate values of X, Y, and Z as set forth in Table 1 where the X, Y, and Z values are in inches from a center point of a bottom surface of the deflector panel. The coordinate values are connected by smooth continuing arcs to form profile sections and the profile sections are joined together smoothly to form the curved profile of the guide vane.
A system for directing cooling air into a gas turbine combustor is provided. The system comprises a transition duct coupled to a flow sleeve, where air to be used for combustor cooling and in the combustion process enters a bellmouth of the transition duct, passes through a plurality of struts within the bellmouth, and is distributed to a passage located between the combustion liner and flow sleeve.
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
F23R 3/02 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la configuration du flux d'air ou du flux de gaz
A heat transfer mechanism is provided comprising a plurality of turbulators located along a surface of a body, such as a combustion liner. The turbulators have a first side with a first ramp angle, a second side with a second ramp angle, a height, and a base width, where the base width is a function of the height and where the turbulators are spaced an axial distance apart that is a function of the turbulator height.
A system and method for improving sealing at a turbine blade tip shroud, while reducing weight associated with the improved sealing is disclosed. The gas turbine blade incorporates a tip shroud having one or more pockets located therein, where the one or more pockets remove weight from the shroud, thus reducing load on the blade attachment generated by additional sealing at the turbine blade shroud. Methods for incorporating the one or more tip shrouds in a new turbine blade or a repaired turbine blade are also disclosed.
F01D 5/20 - Extrémités de pales spécialement façonnées en vue d'obturer l'espace entre ces extrémités et le stator
F01D 5/18 - Aubes creusesDispositifs de chauffage, de protection contre l'échauffement ou de refroidissement des aubes
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
B23P 15/04 - Fabrication d'objets déterminés par des opérations non couvertes par une seule autre sous-classe ou un groupe de la présente sous-classe d'aubes de turbine ou d'organes équivalents, en plusieurs pièces
B33Y 80/00 - Produits obtenus par fabrication additive
50.
Method and apparatus for improving cooling of a turbine shroud
A system and method for cooling a turbine blade tip shroud is provided. The turbine blade comprises a blade attachment, a platform extending radially outward from the attachment, an airfoil extending radially outward from the platform, and a tip shroud extending radially outward from the airfoil. The tip shroud has one or more knife edges extending radially outward from an outer surface of the tip shroud. One or more cooling passages extend through the airfoil and to the tip shroud. The turbine blade also includes one or more tip plates secured at or near the outer surface of the tip shroud thereby forming a plenum between the outer surface and the one or more tip plates. The one or more tip plates also include a plurality of cooling holes for flowing cooling air through the plenum to cool the tip shroud.
A turbine blade having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form an airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.
A turbine blade having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form an airfoil shape. The X, Y, and Z distances may be scaled as a function of the same constant number and the X, Y, and Z distances lie within an envelope of approximately +/−0.032 inches in directions normal to the airfoil.
A turbine blade having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form an airfoil shape. The X, Y, and Z distances may be scaled as a function of the same constant number and the X, Y, and Z distances lie within an envelope of approximately +/- 0.032 inches in directions normal to the airfoil.
A turbine blade having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form an airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.
A turbine nozzle having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, and within an envelope of approximately +/−0.049 inches, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form the airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.
A turbine nozzle having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, and within an envelope of approximately −0.067 to +0.101 inches, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form the airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.
A turbine nozzle having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, and within an envelope of approximately -0.067 to +0.101 inches, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form the airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.
Systems and methods for modal testing of blades. A method for modal testing of a blade comprises providing a generally spherical fixture for retaining the blade in a fixed-free configuration. The method includes using an excitation device for exciting the blade while the blade is retained within a blade retaining portion of the fixture. The method comprises using a measurement device to measure a response of the blade to the excitation. The fixture includes an attachment portion configured for the securement of the excitation device to the fixture.
37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Repair, construction and servicing of gas turbine engines
and their component parts, excluding use in the automotive
industry. Custom manufacturing of gas turbine engines and their
component parts, excluding use in the automotive industry. Engineering services in the field of gas turbine engines for
the commercial aviation, military, energy, business and
general aviation sectors, excluding use in the automotive
industry.
37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Repair and maintenance of gas turbine engines and their
component parts. Custom manufacturing of gas turbine engines and their
component parts. Engineering services in the field of gas turbine engines for
the commercial aviation, military, energy, business and
general aviation sectors.
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
37 - Services de construction; extraction minière; installation et réparation
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
[ Custom ] Manufacturing of gas turbine engines [ and their ] component parts, excluding use in the automotive industry Repair, construction and servicing of gas turbine engines and their component parts, excluding use in the automotive industry Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors, excluding use in the automotive industry
37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
(1) Repair, construction and servicing of gas turbine engines and their component parts
(2) Custom manufacturing of gas turbine engines and their component parts
(3) Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors
37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
(1) Repair, construction and servicing of gas turbine engines and their component parts
(2) Custom manufacturing of gas turbine engines and their component parts
(3) Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors
37 - Services de construction; extraction minière; installation et réparation
40 - Traitement de matériaux; recyclage, purification de l'air et traitement de l'eau
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Repair, construction and servicing of gas turbine engines and their component parts Custom manufacturing of gas turbine engines and their component parts Engineering services in the field of gas turbine engines for the commercial aviation, military, energy, business and general aviation sectors
65.
Turbine blade platform undercut with decreasing radii curve
A turbine blade has an undercut beneath its platform proximate a trailing edge region. The undercut incorporates a curved portion to reduce to reduce undesirable stress concentration. The undercut shape includes a curved portion of decreasing radius with increasing distance from the underside of the platform.
A turbine blade has an undercut beneath its platform proximate a trailing edge region. The undercut incorporates a curved portion to reduce to reduce undesirable stress concentration. The undercut shape includes a curved portion of decreasing radius with increasing distance from the underside of the platform.
A method and apparatus for restoring displaced features on a turbine vane segment for a gas turbine engine, such as a vane segment in a low pressure turbine, and more specifically, the inner shroud thereof relative to the outer shroud thereof to meet the original design position and dimensions.
B21D 9/05 - Cintrage des tubes par utilisation de mandrins ou d'organes analogues coopérant avec des organes de formage
B21D 31/00 - Autres procédés de travail des tôles, tubes ou profilés métalliques
B21D 7/02 - Cintrage des barres, profilés ou tubes sur un organe de formage fixeCintrage des barres, profilés ou tubes par utilisation d'un organe de formage ou d'une butée oscillante
B21D 17/02 - Opérations permettant d'effectuer des rainures individuelles dans des tôles ou dans des objets tubulaires ou creux par pressage
B21D 7/022 - Cintrage des barres, profilés ou tubes sur un organe de formage fixeCintrage des barres, profilés ou tubes par utilisation d'un organe de formage ou d'une butée oscillante sur un organe de formage fixe uniquement
B23P 6/00 - Remise en état ou réparation des objets
68.
PROCESS AND APPARATUS TO RESTORE DISTORTED FEATURES ON GAS TURBINE VANES
A method and apparatus for restoring displaced features on a turbine vane segment for a gas turbine engine, such as a vane segment in a low pressure turbine, and more specifically, the inner shroud thereof relative to the outer shroud thereof to meet the original design position and dimensions.
B23P 6/00 - Remise en état ou réparation des objets
B21D 7/02 - Cintrage des barres, profilés ou tubes sur un organe de formage fixeCintrage des barres, profilés ou tubes par utilisation d'un organe de formage ou d'une butée oscillante
B21D 7/022 - Cintrage des barres, profilés ou tubes sur un organe de formage fixeCintrage des barres, profilés ou tubes par utilisation d'un organe de formage ou d'une butée oscillante sur un organe de formage fixe uniquement
B21D 17/02 - Opérations permettant d'effectuer des rainures individuelles dans des tôles ou dans des objets tubulaires ou creux par pressage
B21D 31/00 - Autres procédés de travail des tôles, tubes ou profilés métalliques
A method and apparatus for restoring displaced features on a turbine vane segment for a gas turbine engine, such as a vane segment in a low pressure turbine, and more specifically, the inner shroud thereof relative to the outer shroud thereof to meet the original design position and dimensions.
B21D 3/10 - Redressage ou remise en forme des barres, tubes ou profilés métalliques, ou des objets déterminés faits à partir de ces matériaux, qu'ils comportent ou non des parties en tôle entre des marteaux et des enclumes ou butées
B21D 3/16 - Redressage ou remise en forme des barres, tubes ou profilés métalliques, ou des objets déterminés faits à partir de ces matériaux, qu'ils comportent ou non des parties en tôle d'objets déterminés faits de barres, tubes ou profilés métalliques, p. ex. de vilebrequins, en utilisant des méthodes ou des moyens spécialement adaptés à ce but
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
F01D 25/28 - Dispositions pour le support ou le montage, p. ex. pour les carters de turbines
B23P 6/00 - Remise en état ou réparation des objets
70.
SURFACE ANALYSIS FOR DETECTING CLOSED HOLES, AND DEVICE
By carrying out laser triangulation measurements on an uncoated component and a coated component with holes, the exact position of the holes to be reopened can be detected following the coating.
G01B 11/24 - Dispositions pour la mesure caractérisées par l'utilisation de techniques optiques pour mesurer des contours ou des courbes
F01D 5/00 - AubesOrganes de support des aubesDispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
F01D 5/18 - Aubes creusesDispositifs de chauffage, de protection contre l'échauffement ou de refroidissement des aubes
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
71.
Process for repairing sulfidation damaged turbine components
A process is provided for cleaning a surface of an internal cavity of a gas turbine component having sulfidation or sulfur bearing deposits comprising: inserting into the internal cavity a fluoride salt; and heating the fluoride salt and the component in an inert atmosphere for a time and at a temperature to clean the sulfidation or sulfur bearing deposits on the surface. In a preferred embodiment the cleaned internal surface is then coated with a metallic coating.
