The present invention relates to a protective material, preferably a ballistic protection material, having a protective function against ballistic active bodies (launch bodies), in particular against reinforcement-penetrating and/or armor-piercing projectiles, thrust bodies or penetrating bodies, and to the use thereof.

In an aspect, a cap liner comprises a sintered fluoropolymer layer; and a backing layer; wherein the sintered fluoropolymer layer is in direct physical contact with the backing layer with no intervening layer located there between. In another aspect, a method of forming the cap liner of comprises plasma etching the sintered fluoropolymer layer to form a sintered plasma etched layer; and laminating the sintered plasma etched layer and the backing layer to form the cap liner.

Disclosed is a joined ceramic body comprising a first ceramic portion comprising a first ceramic, a second ceramic portion comprising a second ceramic, and a joining layer formed between the first ceramic portion and the second ceramic portion. The joining layer has a bond thickness of from 0.5 to 20 um and comprises silicon dioxide having a total impurity content of 20 ppm and less. A method of making the joined ceramic body and a joining material are also disclosed.

Disclosed is a display apparatus including a display panel and a cushion plate disposed under the display panel. The cushion plate may include a porous member and a reinforcing plate. A step in a side area of the display apparatus may be removed, thereby improving impact absorption ability and rigidity.

One aspect of the present invention provides a composite sheet which comprises a nitride sintered body having a porous structure and a semi-cured product of a thermosetting resin composition impregnated into the nitride sintered body, the line roughness Rz specified by JIS B 0601:2013 of at least one main surface being 10 μm or less.

Airfoils for gas turbine engines are provided. In one embodiment, an airfoil formed from a ceramic matrix composite material includes opposite pressure and suction sides extending radially along a span and defining an outer surface of the airfoil. The airfoil also includes opposite leading and trailing edges extending radially along the span. The pressure and suction sides extend axially between the leading and trailing edges. The leading edge defines a forward end of the airfoil, and the trailing edge defining an aft end of the airfoil. Further, the airfoil includes a trailing edge portion defined adjacent the trailing edge at the aft end of the airfoil; a plenum defined within the airfoil forward of the trailing edge portion; and a cooling passage defined within the trailing edge portion proximate the suction side. Methods for forming airfoils for gas turbine engines also are provided.

Disclosed herein are relates to a method for compressing a laminate and a method for manufacturing a ceramic electronic component including a laminate. The method for compressing a laminate includes: preparing a laminate; pressurizing the laminate from a first pressure to a second pressure; heating the laminate from a first temperature to a second temperature; maintaining compression of the laminate at the second pressure and the second temperature for a predetermined time; cooling the laminate from the second temperature to a third temperature; and depressurizing the laminate from the second pressure to a third pressure, wherein the second temperature is 70° C. to 150° C.

A method for manufacturing a turbine nozzle vane made of ceramic matrix composite material, wherein the vane is manufactured using a first fibrous preform including a hollow central section intended to form a fibrous reinforcement of an airfoil of the vane to be obtained, and a pair of second fibrous preforms each having an opening with a shape of the airfoil of the vane to be obtained.

A method for producing a heater with a co-sintered multilayer construction for a system for providing an inhalable aerosol, including providing at least one first substrate layer, arranging at least one first insulating layer at least in areas on the first substrate layer, arranging at least one heating element at least in areas on the first insulating layer, arranging at least one second substrate layer and at least one second insulating layer at least in areas on the heating element. The second insulating layer is arranged at least in areas on the second substrate layer, and the second insulating layer is in contact at least in areas with the heating element and/or with the first insulating layer. The method includes pressing the layers and the heating element, and firing the pressed layers in order to co-sinter the layers of the multilayer construction.

Ceramic matrix composite airfoils for gas turbine engines are provided. In an exemplary embodiment, an airfoil includes opposite pressure and suction sides extending radially along a span. The pressure and suction sides define an outer surface of the airfoil. The airfoil further includes opposite leading and trailing edges extending radially along the span, the pressure and suction sides extending axially between the leading and trailing edges. The airfoil also includes a filler pack defining the trailing edge; the filler pack comprises a ceramic matrix composite material. Moreover, the airfoil includes a plenum defined within the airfoil for receiving a flow of cooling fluid, and a cooling passage defined within the filler pack for directing the flow of cooling fluid from the plenum to the outer surface of the airfoil. Methods for forming airfoils for gas turbine engines also are provided.