Gas Turbine Engine Cooled Cooling Air for Turbine Heat Control
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Solution Overview
Problem
Conventional turbofan engine designs are limited by high temperatures at the exit stage of the high pressure compressor, which restrict compressor pressure ratios and result in prohibitively high exhaust gas temperatures, necessitating the development of a system to cool core components while maintaining or increasing thrust output.
Innovation Solution
Incorporation of a cooled cooling air system that receives airflow from the compressor section, reduces its temperature using a heat exchanger, and supplies it to turbine components, such as high pressure turbine rotor blades, allowing them to withstand higher temperatures.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If compressor pressure ratio is increased to improve thrust output, then engine power is improved, but exhaust gas temperature becomes prohibitively high
Solution Approach 1:
The cooling system is segmented into multiple independent components: a cooled cooling air system with its own ductwork and a separate heat exchanger system. This segmentation allows the cooling function to be added without redesigning the entire engine, enabling higher compressor pressure ratios while controlling exhaust gas temperatures.
Solution Approach 2:
A heat exchanger is introduced as an intermediary component between the compressor outlet and turbine inlet. This heat exchanger transfers heat from the cooling air to the engine core airflow, enabling temperature control that permits higher compressor pressure ratios without prohibitively high exhaust gas temperatures.
2Power
If compressor pressure ratio is increased to improve thrust output, then engine power is improved, but compressor exit temperature becomes prohibitively high
Solution Approach 1:
The cooling air is cooled in advance using a heat exchanger before being introduced into the combustor or turbine sections. This preliminary cooling action allows the system to accommodate higher compressor pressure ratios by pre-controlling the temperature of air that will be used in high-temperature zones.
Solution Approach 2:
The heat exchanger serves as an intermediary that transfers heat from the compressor exit airflow to the cooling air system, thereby reducing the temperature of the main airflow path and enabling higher compressor pressure ratios without excessive compressor exit temperatures.
3Temperature
If heat exchanger is added to cool the airflow, then temperature is reduced, but device complexity increases
Solution Approach 1:
The heat exchanger is integrated into the existing engine structure by utilizing the engine core airflow as the heat transfer medium. This merging approach allows the cooling function to be added without requiring completely separate cooling systems, thereby reducing the increase in device complexity.
Solution Approach 2:
The engine's own core airflow is used to provide the cooling function through the heat exchanger, rather than requiring entirely external cooling systems. This self-service approach reduces system complexity by utilizing resources already present in the engine design.
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 cooled cooling air system enables the turbofan engine to operate at higher temperatures with maintained or increased thrust output, addressing the limitations of conventional designs by reducing gas temperatures and optimizing engine performance.
Implementation Method 1
a cooled cooling air system in thermal communication with a heat exchanger for reducing a temperature of the cooling airflow
Data Source
AI summary
A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT/(AHPCExit2×1000).