Gas Turbine Ring Segment Chamfered Surface Design
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Solution Overview
Problem
Gas turbine engines face challenges in managing high temperature environments and heat transfer issues, particularly in the turbine section where rotating blades and vanes operate, leading to inefficiencies and potential damage from abrasive particles and heat-related stress.
Innovation Solution
A ring segment design for gas turbine engines featuring an arcuate body with a chamfered surface and cooling channels, where the chamfered surface directs circumferential flow to reduce heat transfer and abrasive impact, and a coating with high fracture toughness is applied to enhance durability and aerodynamic performance.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If turbine blades and vanes operate in high temperature environment to increase power output, then power is improved, but heat-related stress and potential damage increase
Solution Approach 1:
A coating layer is applied to the turbine blades and vanes as an intermediary protective barrier between the hot gas environment and the metal substrate. This coating mediates the thermal interaction, allowing the component to operate in high temperature environments while protecting against heat-related stress and damage.
Solution Approach 2:
The operating temperature parameters of the turbine components are increased to improve power output, while simultaneously implementing protective measures (coatings and cooling) to manage the resulting increased thermal stress and prevent damage.
2Temperature
If cooling channels are added to turbine blades to reduce heat transfer, then temperature is improved, but device complexity increases
Solution Approach 1:
The turbine blades are segmented to include internal cooling channels within their structure. These channels divide the blade interior into separate flow paths, allowing coolant to be distributed throughout the blade to reduce heat transfer from the hot gas environment to the metal substrate.
Solution Approach 2:
Cooling channels are integrated into the turbine blade structure to circulate coolant through the blade interior. This pneumatic/hydraulic system reduces heat transfer by having cool fluid flow through channels in direct thermal contact with the blade material, absorbing heat from the hot exterior.
3Reliability
If a coating with high fracture toughness is applied to protect from wear, then reliability is improved, but manufacturing precision requirements increase
Solution Approach 1:
A composite structure is created by applying a coating layer with high fracture toughness over the turbine blade substrate. This composite material system combines the structural integrity of the metal blade with the enhanced wear and thermal resistance of the coating, improving overall reliability.
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
The design reduces heat transfer and abrasive damage, improves aerodynamic performance, and protects the coating from wear, leading to increased efficiency and longevity of the gas turbine engine components.
Implementation Method 1
the chamfered surface directs circumferential flow to reduce heat transfer
Implementation Method 2
a coating with high fracture toughness is applied to enhance durability and protect from wear
Data Source
Figure 1
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AI summary
A ring segment (300) for a gas turbine engine (100) includes a forward mate face (302) with respect to a circumferential flow component (502)of a working fluid (310) of the gas turbine engine (100), an aft mate face (304) opposite to the forward mate face (302), an arcuate body (312) extending between the forward mate face (302) and the aft mate face (304), the arcuate body (312) having a first surface (314) facing to the working fluid (310) and a second surface (316) opposite to the first surface (314). The first surface (314) includes an arcuate surface (318) extending from the aft mate face (304) toward the forward mate face (302), the arcuate surface (318) having an arcuate cross section taken in a section plane that is normal to a central axis of the gas turbine engine (100), and a chamfered surface (320) extending from the forward mate face (302) toward the aft mate face (304), the chamfered surface (320) having a non-arcuate cross section taken in the section plane.