Outer Band Cooling Passages for Interdigitated Turbine Blades
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Solution Overview
Problem
Existing gas turbine engine designs face limitations in thermal management for outer diameter mounted turbine blades, particularly in interdigitated turbine sections, which restrict their application due to high combustion gas temperatures and structural deterioration, limiting their use to low-pressure turbines and preventing effective cooling.
Innovation Solution
A gas turbine engine design featuring an annular outer band with airfoil cooling passages that extend radially and circumferentially, providing thermal attenuation and enabling interdigitation of turbine rotors by directing cooling air through apertures and serpentine structures to reduce thermal gradients across the blades.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If interdigitated turbine sections are used to take advantage of high fluid velocities, then operational efficiency is improved, but thermal management capability deteriorates due to lack of effective cooling structures
Solution Approach 1:
The outer band is segmented into multiple sections with discrete cooling passages distributed circumferentially. Each cooling passage is a separate channel that can be independently designed and controlled, allowing precise thermal management across different regions of the turbine blade while maintaining the interdigitated configuration for high fluid velocity exploitation
Solution Approach 2:
Cooling air is introduced as an intermediary substance to transfer heat away from the turbine blade. The cooling passages serve as conduits for this intermediary cooling medium, which absorbs thermal energy from the blade structure and transports it to cooling apertures where it is discharged, thereby enabling thermal management in the interdigitated turbine section
2Power
If outer diameter mounted turbine blades are used in high temperature environments, then power output is improved, but structural deterioration increases due to thermal damage
Solution Approach 1:
Cooling air is supplied to the turbine blade before it is exposed to high temperature combustion gases. The cooling passages are pre-filled with cooler air from the compressor, creating a thermal barrier that protects the blade structure from immediate thermal damage when the blade enters the high temperature environment, thereby extending its service life while maintaining power output capability
Solution Approach 2:
The temperature parameter of the air within the cooling passages is actively changed and controlled. Cooler air is introduced at the inlet and circulated through the passages, creating a temperature gradient that protects the blade. The cooling system dynamically adjusts the cooling air parameters to maintain structural integrity under varying operating conditions and temperature exposures
3Temperature
If cooling passages are added to provide thermal attenuation, then thermal management is improved, but device complexity increases
Solution Approach 1:
The outer band structure serves multiple functions simultaneously: it provides the mechanical mounting structure for the turbine blades, acts as a containment structure for the cooling passages, and serves as the discharge path for cooling air through circumferential apertures. This multi-functionality reduces the need for separate cooling components, thereby limiting the increase in device complexity while achieving effective thermal management
Solution Approach 2:
The cooling passages are nested within the outer band structure itself. The passages are formed as internal cavities or channels within the existing outer band geometry, rather than being separate external components. This nesting approach integrates the cooling function into the structural component, minimizing additional complexity while providing effective thermal attenuation
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This design enhances fuel efficiency, operational efficiency, and power output while reducing weight, part count, and packaging, enabling interdigitation of high-pressure turbines and improving thermal management, thus extending the application of interdigitated turbine sections to higher temperature environments.
Implementation Method 1
The outer band defines a plurality of airfoil cooling passages in which the plurality of airfoil cooling passages are extended at least partially in the radial direction in fluid communication with the plurality of airfoils
Implementation Method 2
directing cooling air through apertures and serpentine structures to reduce thermal gradients across the blades
Data Source
AI summary
The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a first turbine rotor. The first turbine rotor includes an annular outer band disposed outward of the core flowpath along the radial direction. The first turbine rotor further includes a plurality of airfoils coupled to an inner diameter of the outer band in which the plurality of airfoils are extended generally inward along the radial direction. The outer band defines a plurality of airfoil cooling passages in which the plurality of airfoil cooling passages are extended at least partially in the radial direction in fluid communication with the plurality of airfoils.