Tandem Rotor Blades for Gas Turbine Compressor Heat Reduction
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Conventional gas turbine engines experience increased temperatures due to windage heat-up in the high-pressure compressor section, which reduces the durability of aerospace components, particularly in the last stages, due to knife edge seals in stator vane stages.
Innovation Solution
The implementation of a tandem blade stage configuration that combines two discrete blade stages into a single stage, eliminating the need for an intervening stator vane stage and its associated shrouded cavity, thereby reducing windage heat-up and metal temperatures.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If conventional stator vane stages with knife edge seals are used in the high pressure compressor, then the compressor can condition and guide fluid flow through multiple stages, but windage heat-up increases temperatures and reduces component durability
Solution Approach 1:
The patent combines two discrete blade stages into a single tandem blade stage, eliminating the intervening stator vane stage and its associated knife edge seal. This merging reduces the number of seal interfaces with the rotor disk, thereby reducing windage heat-up and component temperatures while maintaining compression functionality.
Solution Approach 2:
The invention extracts and removes the problematic stator vane stage with knife edge seals from between the two blade stages. By taking out this intermediate stage, the design eliminates the source of windage heat-up that was causing increased temperatures and reduced seal durability.
2Ease of operation
If multiple stator vane stages are used to condition fluid flow, then flow conditioning is achieved, but the compressor length increases due to necessary gaps between stages
Solution Approach 1:
The tandem blade stage merges two compression functions into a single integrated stage, eliminating the need for an intervening stator vane stage and the associated gaps between stages. This reduces the overall compressor length while maintaining the ability to condition and guide fluid flow through the blade pairs.
3Adaptability or versatility
If conventional discrete blade stages with intervening stator vanes are used, then each stage can independently condition flow, but the overall system complexity increases
Solution Approach 1:
The tandem blade stage combines two blade stages into one integrated unit with blade pairs operating in sequence without intervening stator vanes. This reduces the number of discrete components and assemblies required, simplifying the overall compressor stage architecture while maintaining flow conditioning capability.
Data Source
AI summary
A gas turbine engine includes a compressor section and a compressor case with a low pressure compressor (LPC) and a high pressure compressor (HPC). The HPC is aft of the LPC. The compressor case defines a centerline axis. The compressor section also includes a rotor disk defined between the compressor case and the centerline axis. A plurality of stages are defined radially inward relative to the compressor case. The plurality of stages include at least one tandem blade stage. The tandem blade stage includes a plurality of blade pairs. Each blade pair is circumferentially spaced apart from the other blade pairs, and is operatively connected to the rotor disk. Each blade pair includes a forward blade and an aft blade. The aft blade is configured to further condition air flow with respect to the forward blade without an intervening stator vane stage shrouded cavity therebetween.