Abstract: An airfoil for use with a gas turbine engine includes a pressure side wall and a suction side wall. The suction side wall is configured to be exposed to less pressure than the pressure side wall during operation of the gas turbine engine. The blade also includes a plurality of ribs forming a plurality of trapezoidal shaped cavities to receive a cooling airflow. The plurality of ribs being shaped in a way to allow for thermal growth of the airfoil, while minimizing stress in the airfoil.

Abstract: A blade for use with a gas turbine engine includes an attachment and an airfoil. The airfoil further includes a suction side wall configured to be exposed to less pressure than the pressure side wall during operation of the gas turbine engine. The blade also includes a plurality of intersecting ribs transitioning from an airfoil cross sectional geometry to an attachment cross section geometry.

Abstract: A blade for a gas turbine engine includes a body that includes an airfoil that extends in a radial direction from a 0% span position near an airfoil base to a 100% span position at an airfoil tip. The airfoil has a leading edge and a trailing edge that define the true chord length. The airfoil includes a first portion near the airfoil base with a first density and a second portion near the airfoil tip with a second density. The second density is less than the first density. The second portion includes an increasing true chord length in the radial direction. The second portion is in the range of 90% span to 100% span.

Abstract: A turbine section for a gas turbine engine includes a first rotor assembly with a first rotor assembly bleed air source and an aft cavity that is in fluid communication with the first rotor assembly bleed air source. A second rotor assembly includes a forward cavity. A vane bleed air source is in fluid communication with the forward cavity. A seal extends between the first rotor assembly and the second rotor assembly.

Abstract: An airfoil for use with a gas turbine engine includes a pressure side wall. The airfoil further includes a suction side wall configured to be exposed to less pressure than the pressure side wall during operation of the gas turbine engine. The airfoil also includes a plurality of ribs defining multiple air passages including at least one triangular air passage partially defined by at least one of the pressure side wall or the suction side wall, and at least one internal air passage bordered on at least three sides by at least two of the plurality of ribs.

Abstract: A blade for a gas turbine engine includes a body that includes an airfoil that extends in a radial direction from a 0% span position near an airfoil base to a 100% span position at an airfoil tip. The airfoil includes a first portion near the airfoil base with a first density and includes a second portion near the airfoil tip with a second density. The second density is less than the first density. The second portion includes an increasing true chord length in the radial direction.

Abstract: An airfoil for use with a gas turbine engine includes a pressure side wall. The airfoil further includes a suction side wall configured to be exposed to less pressure than the pressure side wall during operation of the gas turbine engine. The airfoil also includes a plurality of ribs defining multiple air passages including at least one triangular air passage partially defined by at least one of the pressure side wall or the suction side wall, and at least one internal air passage bordered on at least three sides by at least two of the plurality of ribs.

Abstract: A seal segment for a gas turbine engine includes a first axial span that extends between the first radial span and the second radial span. A second axial span extends between the first radial span and the second radial span, the first radial span, the second radial span, the first axial span and the second axial span forming a torque box.

Abstract: A turbine section for a gas turbine engine includes a first rotor assembly with a first rotor assembly bleed air source and an aft cavity that is in fluid communication with the first rotor assembly bleed air source. A second rotor assembly includes a forward cavity. A vane bleed air source is in fluid communication with the forward cavity. A seal extends between the first rotor assembly and the second rotor assembly.

Abstract: A blade for a gas turbine engine includes a body that includes an airfoil that extends in a radial direction from a 0% span position near an airfoil base to a 100% span position at an airfoil tip. The airfoil includes a first portion near the airfoil base with a first density and includes a second portion near the airfoil tip with a second density. The second density is less than the first density. The second portion includes an increasing true chord length in the radial direction.

Abstract: A seal segment for a gas turbine engine includes a first axial span that extends between the first radial span and the second radial span. A second axial span extends between the first radial span and the second radial span, the first radial span, the second radial span, the first axial span and the second axial span forming a torque box.

Abstract: A robust multiple-walled, multi-pass, high cooling effectiveness cooled turbine vane or blade designed for ease of manufacturability, minimizes cooling flows on highly loaded turbine rotors. The vane or blade design allows the turbine inlet temperature to increase over current technology levels while simultaneously reducing turbine cooling to low levels. A multi-wall cooling system is described, which meets the inherent conflict to maximize the flow area of the cooling passages while retaining the required section thickness to meet the structural requirements. Independent cooling circuits for the vane or blade's pressure and suction surfaces allow the cooling of the airfoil surfaces to be tailored to specific heat load distributions (that is, the pressure surface circuit is an independent forward flowing serpentine while the suction surface is an independent rearward flowing serpentine). The cooling air for the independent circuits is supplied through separate passages at the base of the vane or blade.

Abstract: A robust multiple-walled, multi-pass, high cooling effectiveness cooled turbine vane or blade designed for ease of manufacturability, minimizes cooling flows on highly loaded turbine rotors. The inventive vane or blade design allows the turbine inlet temperature to increase over current technology levels while simultaneously reducing turbine cooling to low levels. The invention comprises a complex multi-wall cooling system, which meets the inherent conflict to maximize the flow area of the cooling passages while retaining the required section thickness to meet the structural requirements. Independent cooling circuits for the vane or blade's pressure and suction surfaces allow the cooling of the airfoil surfaces to be tailored to specific head load distributions (that is, the pressure surface circuit is an independent forward flowing serpentine while the suction surface is an independent rearward flowing serpentine).