Abstract: A method of fabrication of a rocket engine nozzle assembly using pressure brazing generally includes initially assembling a rocket engine nozzle liner into a rocket engine nozzle jacket for a rocket engine nozzle assembly. The rocket engine nozzle assembly may then be sealed. Prior to pressure brazing the rocket engine nozzle assembly, pressure brazing parameters may be determined. The pressure brazing may be performed with the determined pressure brazing parameters to complete the fabrication of the rocket engine nozzle. The rocket engine nozzle assembly may include a rocket engine nozzle jacket and a rocket engine nozzle liner having a plurality of channels, with the space between each channel defining a land, and the rocket engine nozzle liner having at least a pair of endlands disposed at each end thereof. The rocket engine nozzle liner is bonded to the rocket engine nozzle jacket by the endlands being bonded to the nozzle jacket and the lands being pressure brazed to the nozzle jacket.

Abstract: A method of fabrication of a rocket engine nozzle assembly using pressure brazing generally includes initially assembling a rocket engine nozzle liner into a rocket engine nozzle jacket for a rocket engine nozzle assembly. The rocket engine nozzle assembly may then be sealed. Prior to pressure brazing the rocket engine nozzle assembly, pressure brazing parameters may be determined. The pressure brazing may be performed with the determined pressure brazing parameters to complete the fabrication of the rocket engine nozzle. The rocket engine nozzle assembly may include a rocket engine nozzle jacket and a rocket engine nozzle liner having a plurality of channels, with the space between each channel defining a land, and the rocket engine nozzle liner having at least a pair of endlands disposed at each end thereof. The rocket engine nozzle liner is bonded to the rocket engine nozzle jacket by the endlands being bonded to the nozzle jacket and the lands being pressure brazed to the nozzle jacket.

Abstract: A method of fabrication of a rocket engine nozzle assembly using pressure brazing generally includes initially assembling a rocket engine nozzle liner into a rocket engine nozzle jacket for a rocket engine nozzle assembly. The rocket engine nozzle assembly may then be sealed. Prior to pressure brazing the rocket engine nozzle assembly, pressure brazing parameters may be determined. The pressure brazing may be performed with the determined pressure brazing parameters to complete the fabrication of the rocket engine nozzle. The rocket engine nozzle assembly may include a rocket engine nozzle jacket and a rocket engine nozzle liner having a plurality of channels, with the space between each channel defining a land, and the rocket engine nozzle liner having at least a pair of endlands disposed at each end thereof. The rocket engine nozzle liner is bonded to the rocket engine nozzle jacket by the endlands being bonded to the nozzle jacket and the lands being pressure brazed to the nozzle jacket.

Abstract: A method is provided for damping vibrations in a turbine. The method includes performing structural dynamics analysis on the turbine to determine at least one area of high vibrational stress on the turbine, and performing thermal analysis of the turbine to determine at least an approximated maximum operating temperature at the area of high vibrational stress. Additionally, the method includes utilizing hysteresis damping to dampen operational vibrations. The hysteresis damping includes selecting a shape memory alloy (SMA) having a martensitic-to-austenite transformation temperature substantially similar to the approximate maximum operating temperature of the component at the area of high vibrational stress, and disposing the selected SMA on the turbine on the related area of high vibratory stress.

Abstract: A method is provided for damping vibrations in a turbine. The method includes performing structural dynamics analysis on the turbine to determine at least one area of high vibrational stress on the turbine, and performing thermal analysis of the turbine to determine at least an approximated maximum operating temperature at the area of high vibrational stress. Additionally, the method includes utilizing hysteresis damping to dampen operational vibrations. The hysteresis damping includes selecting a shape memory alloy (SMA) having a martensitic-to-austenite transformation temperature substantially similar to the approximate maximum operating temperature of the component at the area of high vibrational stress, and disposing the selected SMA on the turbine on the related area of high vibratory stress.