Modular Transvane Assembly Reduces Turbulence in Gas Turbine Combustors
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Solution Overview
Problem
Conventional gas turbine combustion engine configurations experience inefficiencies due to energy loss through turbulence and seal wear, requiring complex and costly cooling structures, and result in mechanical stress on turbine blades from uneven gas flow distribution and redirection.
Innovation Solution
A modular transvane assembly system that unites individual gas flows from combustor cans into a singular annular gas flow with a circumferential component, eliminating trailing edges and reducing aerodynamic losses by using fewer seals and no flow redirecting vanes, while maintaining the dynamic isolation of can annular combustor configurations.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Loss of energy
If conventional transition, vane, and seals structures are used to transport gas flow, then the gas flow can be directed to turbine blades, but energy is lost through turbulence and seal wear
Solution Approach 1:
The patent combines the transition duct and vane functions into a single integrated component. The transition duct is shaped to directly guide gas flow from the combustor to the turbine blades without requiring separate vanes to redirect the flow, thereby eliminating the energy losses associated with multiple components and reducing overall structural complexity.
Solution Approach 2:
The patent removes the traditional vane component from the gas flow path. Instead of using vanes to redirect flow after the transition duct, the design integrates the flow directionality directly into the transition duct geometry, extracting the redundant vane element that caused energy losses through turbulence and multiple seals.
2Loss of energy
If multiple seals are used between components, then gas flow can be contained, but gas flow loss increases through seal wear
Solution Approach 1:
By merging the transition duct and vane into one component, the patent eliminates the seals that would normally be required between separate vanes and ducts. This single-component design reduces the number of sealing surfaces, minimizing gas flow loss through seal wear and improving overall reliability.
3Shape
If flow is redirected by vanes, then gas flow can be oriented properly, but significant energy is lost
Solution Approach 1:
The patent merges the flow redirection function into the transition duct geometry itself. The duct is shaped with curved surfaces and optimized cross-sections that guide the gas flow smoothly from the combustor to the turbine blades, eliminating the need for separate vanes that cause energy loss through abrupt redirection.
4Reliability
If cooled components are used to handle heat, then component durability is improved, but manufacturing cost and complexity increase
Solution Approach 1:
The patent removes the need for complex cooling structures by optimizing the gas flow path to minimize thermal exposure. The integrated transition duct design reduces the surface area of hot components and eliminates the need for active cooling systems, thereby simplifying manufacturing while maintaining component durability through reduced thermal stress.
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 energy delivery to turbine blades, reduces mechanical stress, and lowers manufacturing and maintenance costs by minimizing pressure, temperature, and flow direction gradients, while allowing for easier assembly and reduced component wear.
Implementation Method 1
Gas flow energy is lost through turbulence created in the flow as the flow transitions from one component to the next
Data Source
AI summary
An arrangement for delivering gasses from can combustors of a can annular gas turbine combustion engine to a turbine first stage section including a first row of turbine blades, the arrangement including a flow-directing structure for each combustor, wherein each flow-directing structure includes a straight path and an annular chamber end, wherein the annular chamber ends together define an annular chamber for delivering the gas flow to the turbine first stage section, wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, and wherein the annular chamber is configured to unite the respective straight gas flows along respective shear planes to form a singular annular gas flow, and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber to the first row of blades.