A thermal barrier coating includes: a metal bonding layer made of an alloy containing Al; a thermal barrier layer; and an intermediate layer provided between the metal bonding layer and the thermal barrier layer. The thermal barrier layer contains a compound represented by the following general formula (1), Ln.sub.xTa.sub.yHf.sub.zO.sub.(3x+5y+4z)/2 (1) (wherein Ln is an atom of one element selected from among rare earth elements, x is 0 to 1.0, y is 0.8 to 3.0, and z is 0 to 7.0). The intermediate layer contains at least one layer selected from among the following (A) to (C): (A) a layer containing hafnium oxide (HfO.sub.2) as a main component; (B) a layer containing, as a main component, a compound consisting of tantalum (Ta), hafnium (Hf), and oxygen (O); and (C) a layer containing, as a main component, a compound consisting of a rare earth element, tantalum (Ta), hafnium (Hf), and oxygen (O).

The invention relates to a movable blade made of aluminum and titanium alloy, for a turbojet engine turbine comprising a vane and at least one root at a distal end of the vane. The root has at least one azimuthal contact surface with another directly adjacent blade. A hard abrasion-resistant material, called wear-resistant material, is deposited onto the at least one azimuthal contact surface. A cavity is produced in said at least one azimuthal contact surface, the wear-resistant material being deposited in the cavity.

A structural guide vane including a vane body contact surface and a vane edge, wherein at least one of the vane body contact surface or the vane edge comprises a fiber metal laminate.

Cladding material is applied by laser to a net-shape. A method of cladding a host component includes installing the component in a fixture. A shroud component is located against the host component adjacent a select location for the cladding. Cladding is applied to the host component to the select location and adjacent to shroud component so that the shroud component defines an edge of the cladding as applied. The edge of the cladding as defined by the shroud component defines a desired cladding profile requiring no/approximately no post-cladding processing to remove over-cladded material.

A method for producing a sealing component has a body made of superalloy covered by a coating to be placed in contact with a gas turbine blade tip. Steps are carried out in which: a) the new coating is produced by moulding from an alloy of a cobalt-nickel-chromium-aluminium-yttrium (CoNiCrAlY) type further containing between 0.5 and 5% by mass (wt %) of boron, and b) the superalloy body and the new coating are brazed together in order to obtain the said sealing component.

An airfoil includes a first airfoil piece and a second airfoil piece that is bonded to the first airfoil piece at a joint. The first airfoil piece and the second airfoil piece are formed of aluminum alloys. At least one of the aluminum alloys is an aluminum alloy composition that has greater than 0.8% by weight of zinc. The joint includes a braze element of magnesium, zinc, or combinations thereof in a higher concentration than in other portions of the first airfoil piece and the second airfoil piece.

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 seal assembly for a fluid transfer tube in a gas turbine engine is disclosed. In various embodiments, the seal assembly includes a base member having a first side configured to mate with a casing and a second side opposite the first side, an annular ring configured to mate with the second side of the base member and to surround a portion of the fluid transfer tube, a first O-ring disposed between the annular ring and the fluid transfer tube, a second O-ring disposed between the base member and the annular ring, and an attachment ring configured to secure the annular ring and the base member to the casing.

Coating systems for components of a gas turbine engine, such as a compressor case, are provided. The coating system can include a dense layer disposed along the inner surface of the compressor case as well as an abradable, top coat disposed along the dense layer. The combination of dense layer and abradable top coat can reduce the occurrence of galvanic corrosion of the coating system and thereby increase the lifetime of the coating system and preserve blade clearances within the compressor. Methods are also provided for applying the coating system onto a compressor case.

An airfoil extends in a radial direction a span in a range 3.97-4.27 inch (100.9-108.6 mm). A chord length extends in a chordwise direction from a leading edge to a trailing edge at 50% span in a range 1.28-1.58 inch (32.4-40.0 mm). The airfoil includes at least two of a first mode with a frequency of 24310% Hz, a second mode with a frequency of 37410% Hz, a third mode with a frequency of 74110% Hz, a fourth mode with a frequency of 119810% Hz, a fifth mode with a frequency of 166310% Hz, a sixth mode with a frequency of 241110% Hz, a seventh mode with a frequency of 373410% Hz, an eighth mode with a frequency of 673810% Hz and a ninth mode with a frequency of 897710% Hz.