Abstract: Antigen binding agents (e.g., single domain antibodies) that bind guanylyl cyclase C (GCC) are disclosed. Nucleic acids, recombinant expression vectors, host cells, antigen binding fragments, and pharmaceutical compositions comprising these antigen binding agents and fragments thereof are also disclosed. The invention also provides therapeutic methods for utilizing the antibodies and antigen-binding molecules provided herein.

Abstract: Embodiments of the invention relate to various cooling systems for a turbine vane made of stacked ceramic matrix composite (CMC) laminates. Each airfoil-shaped laminate has an in-plane direction and a through thickness direction substantially normal to the in-plane direction. The laminates have anisotropic strength characteristics in which the in-plane tensile strength is substantially greater than the through thickness tensile strength. Such a vane construction lends itself to the inclusion of various cooling features in individual laminates using conventional manufacturing and forming techniques. When assembled in a radial stack, the cooling features in the individual laminates can cooperate to form intricate three dimensional cooling systems in the vane.

Abstract: An airfoil (44) formed of a plurality of pre-fired structural CMC panels (46, 48, 50, 52). Each panel is formed to have an open shape having opposed ends (54) that are free to move during the drying, curing and/or firing of the CMC material in order to minimize interlaminar stresses caused by anisotropic sintering shrinkage. The panels are at least partially pre-shrunk prior to being joined together to form the desired structure, such as an airfoil (42) for a gas turbine engine. The panels may be joined together using a backing member (30), using flanged ends (54) and a clamp (56), and/or with a bond material (36), for example.

Abstract: Aspects of the invention relate to a modular turbine vane assembly. The vane assembly includes an airfoil portion, an outer shroud and an inner shroud. The airfoil portion can be made of at least two segments. Preferably, the components are connected together so as to permit assembly and disassembly of the vane. Thus, in the event of damage to the vane, repair involves the replacement of only the damaged subcomponents as opposed to the entire vane. The modular design facilitates the use of various materials in the vane, including materials that are dissimilar. Thus, suitable materials can be selected to optimize component life, cooling air usage, aerodynamic performance, and cost. Because the vane is an assemblage of smaller sub-components as opposed to one unitary structure, the individual components of the vane can be more easily manufactured and more intricate features can be included.

Abstract: Ceramic tile (32) insulation for protecting a substrate material (34) in a high temperature environment. A plurality of ceramic tiles (78) may be used in combination with a monolithic layer of ceramic insulation (80) to protect a fillet region (76) and an airfoil section (80), respectively, of a gas turbine vane (72). Individual ceramic tiles (84) may be applied to repair a damaged area of the monolithic insulating layer. Ceramic tile insulation may be applied in two layers (56, 58) with the material properties of the two layers being different, and with the gaps (38) of the two layers being misaligned.

Abstract: A stacked ceramic matrix composite lamellate assembly (10) including shear force bearing structures (48) for resisting relative sliding movement between adjacent lamellae. The shear force bearing structures may take the form of a cross-lamellar stitch (50), a shear pin (62), a warp (90) in the lamellae, a tongue (104) and groove (98) structure, or an inter-lamellar sealing member (112), in various embodiments. Each shear force bearing structure secures a subset of the lamellae, with at least one lamella being common between adjacent subsets in order to secure the entire assembly.

Abstract: Embodiments of the invention relate to a robust turbine vane made of stacked airfoil-shaped CMC laminates. Each laminate has an in-plane direction and a through thickness direction substantially normal to the in-plane direction. The laminates have anisotropic strength characteristics in which the in-plane tensile strength is substantially greater than the through thickness tensile strength. Thus, the laminates can provide strength in the direction of high thermal gradients and, thus, withstand the associated high thermal stresses. The laminates are relatively weak in through thickness (interlaminar) tension, but, in operation, relatively low through thickness tensile stresses can be expected. The laminates can be strong in through thickness compression; accordingly, the laminate stack can be held in through thickness compression by one or more fasteners.

Abstract: Thermally insulating layer incorporating a distinguishing agent and method for inspecting the insulating layer are provided. The distinguishing agent may be used for determining a remaining thickness of the thermally insulating layer.

Abstract: A method of manufacturing a hybrid structure (100) having a layer of CMC material (28) defining an interior passageway (24) and a layer of ceramic insulating material (18) lining the passageway. The method includes the step of casting the insulating material to a first thickness required for effective casting but in excess of a desired second thickness for use of the hybrid structure. An inner mold (14) defining a net shape desired for the passageway remains in place after the casting step to mechanically support the insulating material during a machining process used to reduce the thickness of the insulating material from the as-cast first thickness to the desired second thickness. The inner mold also provides support as the CMC material is deposited onto the insulating material. The inner mold may include a fugitive material portion (20) to facilitate its removal after the CMC material is formed.

Abstract: The present invention relates to weapon systems that accelerate projectiles using gases generated by the rapid combustion of a solid propellant, in particular, such a weapon system is able to vary the barrel exiting velocity of the projectile through a barrel venting means. In one embodiment, a front venting means exhausts gas generated by combusting propellant from behind the accelerating projectile and redirects a portion of the exhausted gas either to at least one fixed volume, to the front of the projectile, or to a combination of at least one fixed volume and to the front of the projectile. Redirecting some of the exhausted gas to the front of the projectile restrains the projectile, thereby slowing the projectile, and thus further decreasing the muzzle velocity of the projectile. In another embodiment, gas from behind the projectile is exhausted into a fixed volume, thereby decreasing projectile acceleration, and thus, the muzzle velocity of the projectile.

Abstract: The present invention relates to weapon systems that accelerate projectiles using gases generated by the rapid combustion of a solid propellant, in particular, such a weapon system is able to vary the barrel exiting velocity of the projectile through a barrel venting means. In one embodiment, a front venting means exhausts gas generated by combusting propellant from behind the accelerating projectile and redirects a portion of the exhausted gas either to at least one fixed volume, to the front of the projectile, or to a combination of at least one fixed volume and to the front of the projectile. Redirecting some of the exhausted gas to the front of the projectile restrains the projectile, thereby slowing the projectile, and thus further decreasing the muzzle velocity of the projectile. In another embodiment, gas from behind the projectile is exhausted into a fixed volume, thereby decreasing projectile acceleration, and thus, the muzzle velocity of the projectile.