Gas Turbine Core Cooling for Higher Compressor Exit Temperatures
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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 restricts compressor pressure ratio and exhaust gas temperatures, necessitating the development of new methods to operate at higher temperatures while maintaining or increasing thrust output and efficiency.
Innovation Solution
Incorporation of a cooled cooling air system that reduces the temperature of airflow using a heat exchanger, providing cooled airflow to turbine components, allowing them to withstand higher temperatures and improving engine performance.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If conventional turbofan engine designs are used, then the engine structure is simple and easy to manufacture, but the compressor pressure ratio and exhaust gas temperatures are restricted due to high temperatures at the exit stage of the high-pressure compressor
Solution Approach 1:
The cooling system is segmented into multiple independent cooling channels, each serving specific turbine stages. This allows selective cooling of different components (first stage turbine, second stage turbine, etc.) with optimized cooling airflow paths, enabling higher compressor exit temperatures while managing thermal loads through distributed cooling zones
Solution Approach 2:
Cooling airflow is extracted from the compressor at multiple stages and cooled beforehand using heat exchangers before being directed to turbine components. This preliminary cooling action allows the system to handle higher compressor exit temperatures by pre-conditioning the cooling air to maintain effective temperature differentials for heat transfer
2Power
If higher temperatures are operated to increase thrust output, then the engine power increases, but the compressor and turbine components suffer from excessive thermal stress and temperature damage
Solution Approach 1:
Different cooling strategies are applied to different components based on their specific thermal requirements. The first stage turbine receives cooling from specific compressor stages, while the second stage turbine receives cooling from different compressor stages. This localized cooling approach optimizes thermal stress management for each component while enabling higher overall engine power output
Solution Approach 2:
Heat exchangers serve as intermediary devices between the compressor airflow and turbine cooling requirements. These heat exchangers transfer thermal energy from the cooling airflow to the compressor discharge air, enabling temperature management that protects turbine components from excessive thermal stress while maintaining high power output capability
3Temperature
If a cooled cooling air system is added to reduce airflow temperature, then the temperature of turbine components is controlled, but the size and weight of the engine increase
Solution Approach 1:
The cooling airflow system serves multiple functions simultaneously: it cools turbine components, conditions compressor discharge air through heat exchangers, and manages thermal loads across different engine operating conditions. This multi-functionality reduces the need for separate dedicated cooling systems, thereby minimizing weight increase while achieving effective temperature control of turbine components
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 turbofan engines to operate at higher temperatures with maintained or increased thrust output, reducing the negative impacts of high temperatures on compressor and turbine components, while minimizing size and weight increases.
Implementation Method 1
a cooled cooling air system that reduces the temperature of airflow using a heat exchanger
Implementation Method 2
reduces the temperature of airflow using a heat exchanger, providing cooled airflow to turbine components
Data Source
AI summary
A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). The high-pressure compressor defines a high-pressure compressor exit area (AHPCExit) in square inches. 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). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.