Turbofan Core Cooling Architecture for High Thrust and EGT Limits
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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 the compressor pressure ratio and leads to prohibitively high exhaust gas temperatures, necessitating a redesign to operate at higher temperatures while maintaining or increasing thrust output and efficiency.
Innovation Solution
Incorporating a cooled cooling air system that reduces the temperature of airflow using a heat exchanger, providing cooled airflow to turbine components, and optimizing engine architecture to accommodate this system without increasing drag or weight, while balancing aerodynamic and structural integrity.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If the compressor pressure ratio is increased to improve thrust output, then the thrust output is improved, but the exhaust gas temperature becomes prohibitively high
Solution Approach 1:
The cooling system is segmented into multiple independent cooling circuits, each serving specific turbine stages. This allows differential cooling strategies for different temperature zones, enabling higher overall pressure ratios without excessive exhaust temperatures by针对性地 cooling specific hot spots in the turbine path.
Solution Approach 2:
Cooling air is extracted from the compressor at multiple stages and cooled before being introduced to the turbine. This preliminary cooling action reduces the temperature of air that will eventually mix with exhaust gases, allowing higher compressor pressure ratios to be achieved while maintaining acceptable exhaust temperature limits.
2Temperature
If a cooled cooling air system is added to reduce airflow temperature, then the core component cooling is improved, but the device complexity increases
Solution Approach 1:
The cooling air system serves multiple functions simultaneously: it cools turbine blades, controls exhaust gas temperature, and provides thermal management for various engine components. By integrating these functions into a unified system with shared infrastructure (ducting, heat exchangers, control mechanisms), the patent reduces overall complexity compared to having separate systems for each function.
Solution Approach 2:
The cooling system is nested within the existing engine architecture, with cooling ducts and heat exchangers integrated into the turbine and compressor structures. The cooling air pathways are nested within the main gas flow path, allowing the cooling system to utilize existing structural spaces rather than requiring separate external systems.
3Productivity
If the compressor pressure ratio is increased, then the efficiency is improved, but the exhaust gas temperature becomes prohibitively high
Solution Approach 1:
The system dynamically adjusts cooling air flow rates and temperatures based on operating conditions to maintain optimal exhaust gas temperature across different efficiency regimes. By changing the parameters of cooling air (temperature, pressure, flow rate) in response to varying engine demands, the system can achieve higher compressor pressure ratios for improved efficiency while preventing exhaust temperature from exceeding limits.
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
Enables turbofan engines to operate at higher temperatures with increased thrust output and efficiency by effectively cooling core components, mitigating vibrations and instability, and improving overall performance.
Implementation Method 1
Incorporating a cooled cooling air system that reduces the temperature of airflow using a heat exchanger
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.