A turbomachine airfoil element includes an airfoil that has pressure and suction sides spaced apart from one another in a thickness direction and joined to one another at leading and trailing edges. The airfoil extends in a radial direction a span that is in a range of 3.84-4.14 inch (97.5-105.1 mm). A chord length extends in a chordwise direction from the leading edge to the trailing edge at 50% span and is in a range of 1.16-1.46 inch (29.5-37.1 mm). The airfoil element includes at least two of a first mode with a frequency of 32610% Hz, a second mode with a frequency of 125210% Hz, a third mode with a frequency of 247510% Hz, a fourth mode with a frequency of 295110% Hz, a fifth mode with a frequency of 435610% Hz and a sixth mode with a frequency of 608610% Hz.

A seal comprises a housing. A coating has at least two layers with a bond layer to be positioned between a housing and a second hard layer. The second hard layer is formed to be harder than the bond layer. The bond layer has a bond strength greater than or equal to 200 psi and less than or equal to 2000 psi. A gas turbine engine, and a method of forming a coating layer are also disclosed.

A mount structure for mounting an ancillary engine unit to a gas turbine engine is provided. The mount structure has plural elongate struts which each extend from a connector portion at one end of the strut to a fastening portion at the other end of the strut. The housing of the ancillary engine unit is formed of a first material having a first coefficient of thermal expansion, and the elongate struts are formed of a second material having a second coefficient of thermal expansion. Each elongate strut extends away from its connector portion in a direction which is crosswise to the direction of the hypothetical differential thermal strain at that connector portion. The mount structure further has a containment bracket which is configured to contain each connector portion.

A blade containment system includes a plurality of circumferentially-arranged rotatable blades. Each blade has a blade compliance. An annular containment structure is arranged around the rotatable blades. The containment structure includes a liner that has a liner compliance. The blade compliance and the liner compliance are configured such that a strain induced on a respective one of the blades upon impact with the liner is less than a threshold critical strain beyond which the rotatable blades fracture.

A turbine component assembly is disclosed, including a first component, a second component, and an interface shield. The first component is arranged to be disposed adjacent to a hot gas path, and includes a ceramic matrix composite composition. The second component is adjacent to the first component and arranged to be disposed distal from the hot gas path across the first component. The interface shield is disposed on a contact region of the first component, and directly contacts the second component.

In one exemplary embodiment, a turbofan engine comprises a fan section. A core section includes a turbine section arranged fluidly downstream from the compressor section. A combustor is arranged fluidly between the compressor and turbine sections. The fan and core sections are configured to produce a thrust in a range 27,000-35,000 pounds-f (120,102-156,688 N). An airfoil is arranged in the fan section. The airfoil has first and second modes each having a frequency. The first mode has the lowest frequency, and the second mode has the second lowest frequency wherein the second mode frequency is 140 Hz or less at a redline engine speed.

A turbomachine airfoil element includes an airfoil that has pressure and suction sides spaced apart from one another in a thickness direction and joined to one another at leading and trailing edges. The airfoil extends in a radial direction a span that is in a range of 2.58-2.88 inch (65.4-73.1 mm). A chord length extends in a chordwise direction from the leading edge to the trailing edge at 50% span and is in a range of 1.59-1.89 inch (40.3-48.0 mm). The airfoil element includes at least two of a first mode with a frequency of 208810% Hz, a second mode with a frequency of 309910% Hz, a third mode with a frequency of 689010% Hz, a fourth mode with a frequency of 720710% Hz, a fifth mode with a frequency of 1124110% Hz, a sixth mode with a frequency of 1191610% Hz and a seventh mode with a frequency of 1260010% Hz.

The invention provides a stator-side member which is arranged in a vacuum pump and which, without the provision of a heat insulation material, prevents the deposition of products at the lower side of a threaded groove pump unit, with this lower side being an area of high pressure where the deposition of products (deposits) occurs easily, and also provides a vacuum pump equipped with this stator-side member. A threaded groove spacer configured to have a coefficient of thermal conductivity lower than a predetermined value is arranged in a vacuum pump equipped with a threaded groove pump unit. (1) The threaded groove spacer is manufactured from a material having a coefficient of thermal conductivity lower than that of a member which opposes or comes into contact with the threaded groove spacer. Specifically, this material has a coefficient of thermal conductivity lower than that of aluminum or aluminum alloy, and is preferably any one of stainless steel, fiber-reinforced plastic, polyetherimide, and polyetheretherketone. (2) The threaded groove spacer is constituted by at least two or more parts.

A method of forming an airfoil includes placing a material onto a die that is heated to a predetermined temperature to pre-heat the material to a first temperature, while the die is in an open position. The method further includes closing the die at a predetermined rate and holding the die in a closed position for a predetermined period of time at a first force. The method still further includes removing the part from the die, cooling the die, placing the part onto the die, and closing the die at a second force.

A gas turbine engine for an aircraft has an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a hub; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has an engine length and a gearbox location relative to a forward region of the fan, and a gearbox location ratio of: gearbox location/engine length is in a range from 0.19 to 0.45.