A turbine vane airfoil and a turbine nozzle that includes the turbine vane airfoil. The turbine vane airfoil includes an airfoil defining a chamber proximate a surface, and an impingement tube. The impingement tube includes a first wall and a second wall. The first wall defines a plurality of first apertures, and each has a first upstream surface portion opposite a first downstream surface portion. The second wall defines a plurality of second apertures, and each has a second upstream surface portion opposite a second downstream surface portion. Each of the first apertures cooperate with a respective one of the second apertures to direct a cooling fluid onto the surface, and the first downstream surface portion of at least one of the first apertures includes a first point that is collinear with a second point of the second downstream surface portion of the respective one of the second apertures.

A rotor for a gas turbine having a rotor disk on which there are a plurality of rotor components distributed around the circumference. The rotor disk has a circumferential securing shoulder with a contact face. Retaining faces come to bear against the contact face, each of the retaining faces have a retaining shoulder of the respective rotor component and are designed with a form that complements the contact face. In order to optimize the bearing stresses between the retaining shoulder and the securing shoulder, the retaining face has a smaller radius than the contact face, namely the retaining radius is at least 0.99 times and at most 0.995 times the contact radius. Also provided is an axially extending aperture in the rotor component, the width of which in the circumferential direction is 25% to 75% of the rotor component width in the circumferential direction.

A turbine vane of a turbine engine is described. The turbine vane includes a blade and a root. The root includes a stilt having lateral flanks with a curvilinear profile. The stilt includes a frangible zone suitable for undergoing a breakage of the stilt if radial forces higher than a threshold are exerted on the vane, in particular centrifugal forces during an overspeed state of the turbine. The frangible zone includes at least one oblong frangibility recess formed on at least one of the lateral flanks of the stilt, the oblong recess extending in an axial direction of the stilt along a longitudinal axis parallel to or included in a minimum cross-sectional plane which contains a minimum cross-section of the stilt.

A compressor aerofoil for a turbine engine includes a tip portion which extends in a first direction from a main body portion defined by a suction surface wall having a suction surface and a pressure surface wall having a pressure surface. The suction and pressure surface walls meet at a leading edge and a trailing edge. The tip portion includes a tip wall which extends continuously along a camber line of the aerofoil, the camber line extending from the leading edge to the trailing edge. A shoulder is provided on each of the suction and pressure surface walls. A transition region tapers from each of the shoulders in a direction towards the tip wall. The tip wall defines a squealer with a tip surface which increases in width from the leading edge to a point of maximum width, and then decreases in width all the way to the trailing edge.

Stator vanes including curved trailing edges are disclosed. The stator vanes may include a body including a central section, a tip section positioned radially above the central section, and a root section positioned radially below the central section. The body of the stator vanes may also include a leading edge extending radially adjacent the root section, central section, and tip section, respectively, and a trailing edge positioned opposite and aft to the leading edge. The trailing edge may include a concave contour including a first portion radially aligned with the central section of the body. The first portion may be axially offset and forward of a reference line that may be perpendicular to an axial direction and intersects the concave contour at the tip section and the root section. A concavity of the first portion of the concave contour may be formed radially aft of the central section.

A guide vane for a variable turbine geometry and a turbine assembly provided with same are described. The guide vane has an outer face that is at least partly concave in design. The inner face opposite the outer face may also be at least partly concave in design. This gives the guide vane a good functional capability when arranged in a vane ring of a turbine, since an overflow through the gap between the guide vane and the neighboring walls is reduced.

An airfoil vane includes an airfoil section including an outer wall that defines an internal cavity; and a baffle situated in the internal cavity, the baffle including a baffle wall that defines a central cavity having a leading end and a trailing end corresponding to a leading end and a trailing end of the airfoil section, and a tail extending from the baffle wall, the tail including at least one feature configured to disturb an airflow surrounding the tail. A baffle for the airfoil vane assembly and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

A nozzle vane for a variable geometry turbocharger satisfies 0.45<(Xp/L)<0.60, where L is a chord length of the nozzle vane, and Xp is a distance between a leading edge of the nozzle vane and a rotation center of the nozzle vane.

A turbine component includes a hot wall, a cold wall spaced apart from the hot wall and a conduit defined between the hot wall and the cold wall. A cooling system is defined in the conduit. The cooling system includes a plurality of cooling pins, each including a first end having a first cross-sectional area and a second end having a second cross-sectional area. Each cooling pin includes a body extending between the first end and the second end, with a pin leading edge defined along the body from the first end to the second end. The pin leading edge is defined by a first diameter and a pin trailing edge is defined by a second diameter. At least one first cooling pin has the first end coupled to the hot wall and the second end coupled to the cold wall with a fillet.

A gas turbine engine component includes a vane arc segment that defines an axis. The vane arc segment includes an airfoil section and first and second platforms. The airfoil section has a first radial end, a second radial end, a first side, a second side, a leading edge, and a trailing edge. The airfoil section has associated characteristics, including a center of pressure and an aerodynamic load vector through the center of pressure. The first platform defines a first side chevron face that has a leading leg and a trailing leg that meet at an angle. The leading leg is elongated along a centerline that is non-axially oriented with respect to the axis, the leading leg meets the leading face at a leading first side corner. The leading first side corner is located outside of the aerodynamic load vector relative to the leading edge of the airfoil section.