BACKGROUND

1. Field

The present invention relates to rotating blades in a turbomachine, and in particular, to a row of turbine blades with one or more internal cavities having a defined frequency mistuning for improved flutter resistance.

2. Description of the Related Art

Turbomachines, such as gas turbine engines, include multiple stages of flow directing elements along a hot gas path in a turbine section of the gas turbine engine. Each turbine stage comprises a circumferential row of stationary vanes and a circumferential row of rotating blades arranged along an axial direction of the turbine section. Each row of blades may be mounted on a respective rotor disc, with the blades extending radially outward from the rotor disc into the hot gas path. A blade includes an airfoil extending span-wise along the radial direction from a root portion to a tip of the airfoil.

Typical turbine blades at each stage are designed to be identical aerodynamically and mechanically. These identical blades are assembled together into the rotor disc to form a bladed rotor system. During engine operation, the bladed rotor system vibrates in system modes. The blade displacement amplitudes caused by this vibration may be more severe in large blades, such as in low pressure turbine stages. For mechanically and aerodynamically identical blades, the aeroelastic modes are patterns of blade vibration with a constant phase angle between adjacent blades which affects the unsteady flow and aerodynamic work done on the blades. In most cases this serves to damp the vibration of adjacent blades. However, under certain conditions, the aerodynamic damping in some of the modes may become negative, which may cause the blades to vibrate in a self-excited manner, called flutter. When this happens, the vibratory response of the system tends to grow exponentially until the blades either reach a limit cycle or break. Even if the blades achieve a limit cycle, their amplitudes can still be large enough to cause the blades to fail from high cycle fatigue.

Frequency mistuning can cause system modes to be distorted by changing the phase angles of adjacent blades, so that the resulting new, mistuned system modes are stable, i.e., they all have positive aerodynamic damping. It may be desirable in some cases to be able to design blades with a certain amount of defined mistuning. Mistuning may be realized by varying the blade frequencies along the rotor disc in a defined manner. Defined mistuning can be a challenge in cooled turbine blades due to casting variation and core movement during the casting process.

Conventionally, mistuning has been implemented on solid blades by removing material on the blade tip, for example, by grinding, to change the frequency of some blades.

SUMMARY

Briefly, aspects of the present invention are directed to an improved technique for implementing defined mistuning in a row of turbine blades with one or more internal cavities.

According to a first aspect of the invention, a bladed rotor system for a turbomachine is provided, which comprises a circumferential row of blades mounted on a rotor disc. Each blade comprises an airfoil having an outer wall delimiting an airfoil interior. The airfoil interior comprises one or more internal cavities. The row of blades comprises a first set of blades and a second set of blades. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of the outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at least one internal cavity, which is unique to blades of a given set. Thereby, the natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. Blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.

According to a second aspect of the invention a method is provided for producing a bladed rotor system. The method comprises forming a plurality of blades, each blade being formed, at least partially, by a casting process. Each blade comprises an airfoil having one or more internal cavities produced by respective core elements during the casting process. The plurality of blades includes a first set of blades and a second set of blades. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of the outer wall of the respective airfoils. The casting process for forming the first set of blades differs from the casting process for forming the second set of blades, in that, the respective core element for producing at least one internal cavity has a different geometry and/or position during casting of a blade belonging to the first set, in relation to a blade belong to the second set. The geometry and/or position of the respective core element is kept substantially identical for forming blades of a given set. Thereby, the natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount.

