Aircraft Cooling System Valve Control for Thermal Management
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Solution Overview
Problem
Current gas turbine engine cooling systems face inefficiencies due to high cooling air mass flow requirements, which reduce thermodynamic efficiency and lead to excessive fuel burn or reduced engine service life, especially during descent and shutdown processes.
Innovation Solution
A method to control the cooling system by operating in a first mode during aircraft descent, using a heat exchanger to provide a higher cooling fluid flow and lower temperature, allowing earlier engine shutdown post-landing without compromising service life, and automatically switching between modes based on flight conditions detected by the Electronic Engine Control Unit.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If cooling air mass flow is increased to cool turbine components, then component temperature is reduced and damage is prevented, but thermodynamic efficiency decreases and fuel burn increases
Solution Approach 1:
The cooling air is cooled in advance using a heat exchanger before being supplied to the turbine components. This preliminary cooling action reduces the temperature of the cooling air, thereby increasing its cooling effectiveness and allowing for reduced cooling air mass flow while maintaining adequate component cooling, thus improving thermodynamic efficiency
Solution Approach 2:
The temperature parameter of the cooling air is changed by cooling it to a lower temperature before delivery to the turbine components. This parameter change increases the temperature differential between the cooling air and the hot components, enhancing cooling efficiency and reducing the required cooling air mass flow
2Reliability
If engine is kept running to provide cool down period after high thrust operation, then component thermal stress is reduced and service life is extended, but fuel burn increases
Solution Approach 1:
The engine components are cooled in advance during the descent phase before the engine is shut down. This preliminary cooling action reduces the temperature of the components prior to shutdown, allowing the engine to be turned off earlier while still maintaining adequate cooling to prevent thermal damage and extend service life
Solution Approach 2:
The natural cool-down process that requires continued engine operation is replaced by an active cooling system using a heat exchanger. This substitution allows for controlled cooling during descent, enabling earlier engine shutdown without compromising component thermal management
3Temperature
If cooling system provides high cooling air flow, then component cooling effectiveness is improved, but engine efficiency and productivity are reduced
Solution Approach 1:
The cooling air is pre-cooled using a heat exchanger before being delivered to the turbine components. This preliminary cooling action increases the cooling effectiveness of the air, allowing for reduced cooling air flow rates while maintaining adequate component cooling, thus improving engine efficiency and productivity
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 approach reduces fuel burn, minimizes thermal stress on engine components, and ensures the cooling system is operational for 'go-around' scenarios, extending engine life and reducing maintenance costs.
Implementation Method 1
Compressed cooling air from the high pressure compressor 14 flows through the heat exchanger 22 along a second fluid path 26 which is in thermal contact with the first fluid path 24. Consequently, the cooling air in the second fluid path 26 is cooled prior to delivery to the components which require cooling.
Data Source
Figure 1~2
Figure 3
Figure 3
AI summary
A method of controlling an aircraft gas turbine engine cooling system (20). The cooling system comprises a heat exchanger (22) having a first fluid path (24) through which fan air can flow, and a second fluid path (26) through which relatively hot compressor air can flow. The cooling system (20) further comprises a valve (30) configurable between an open position corresponding to a first operating mode, in which fan air flows through the first path (24) to cool the compressor air in the second path (26) to a lower temperature, and a closed position corresponding to a second operating mode, in which fan air flows through the first path (24) at a reduced rate, or not at all, such that the compressor air in the second path (26) is cooled to a lesser extent, or not at all. The valve (30) is operated in the first mode when the aircraft determined to be in a descent mode, or when turbine entry temperature (T30) of the gas turbine engine is above a predetermined amount. Otherwise, the valve (30) is operated in the second mode.