Cooled Aerofoil with Segmented Internal Passages
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Solution Overview
Problem
Existing gas turbine engine aerofoil cooling arrangements fail to effectively manage cooling air flow to match varying external static pressures across the aerofoil surfaces, leading to inefficiencies in heat transfer and potential hot gas ingestion, which compromises turbine efficiency and structural integrity.
Innovation Solution
A cooled aerofoil design featuring separate internal passages for pressure and suction surfaces, with a radial multi-pass serpentine passage for the suction side to match cooling air pressure with external hot gas static pressure, and flow-disrupting formations to enhance heat transfer and reduce pressure losses, allowing for independent optimization of flow conditions for each surface.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Device complexity
If a single internal passage is used to supply cooling air to both pressure and suction surfaces, then the device complexity is reduced, but the heat transfer effectiveness deteriorates due to inability to match cooling air pressure with varying external static pressures
Solution Approach 1:
The cooling system is divided into separate internal passages for the pressure side and suction side, allowing independent optimization of cooling air flow for each surface. This segmentation enables the cooling air pressure to be matched with the varying external static pressures on different aerofoil surfaces, improving heat transfer effectiveness without excessive complexity.
Solution Approach 2:
Different cooling air supply conditions are provided for different locations on the aerofoil. The pressure side receives cooling air optimized for its higher external static pressure, while the suction side receives cooling air optimized for its lower external static pressure. This local quality approach ensures optimal heat transfer effectiveness at each location.
2Reliability
If cooling air pressure is increased to improve heat transfer, then the heat transfer effectiveness is improved, but the risk of hot gas ingestion increases due to pressure differential
Solution Approach 1:
The cooling air pressure is locally optimized for each aerofoil surface based on its external static pressure conditions. The pressure side, experiencing higher external pressure, receives higher cooling air pressure to maintain effective heat transfer. The suction side, experiencing lower external pressure, receives lower cooling air pressure to prevent hot gas ingestion. This local quality approach resolves the contradiction between heat transfer effectiveness and hot gas ingestion risk.
3Ease of manufacture
If conventional cooling arrangements are used, then the manufacturing cost is reduced, but the turbine efficiency deteriorates due to inadequate cooling effectiveness
Solution Approach 1:
The cooling system is segmented into separate passages for pressure and suction sides, improving cooling effectiveness and turbine efficiency. While this increases manufacturing complexity compared to conventional single-passage designs, the improved efficiency justifies the additional manufacturing cost by enabling better heat transfer management.
Solution Approach 2:
The cooling system optimizes multiple parameters including cooling air pressure, flow distribution, and passage geometry to maximize turbine efficiency. These parameter changes improve cooling effectiveness but require more sophisticated manufacturing processes, increasing manufacturing cost while enhancing 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 design enhances internal heat transfer and film cooling effectiveness, reduces the risk of hot gas ingestion, and improves the structural integrity of the aerofoil by matching cooling air pressure with external static pressure, thereby increasing turbine efficiency and reducing operational costs.
Implementation Method 1
the second internal passage is a radial multi-pass passage which extends along a serpentine path from its entrance to the passage towards the leading edge of the aerofoil, which has flow-disrupting formations on its internal surface
Implementation Method 2
flow-disrupting formations on its internal surface, thereby matching the cooling air pressure at the externals holes served by the passage to the external hot gas static pressure
Implementation Method 3
Cooling air flows to and from an NGV 1 and a rotor blade 2 are indicated by arrows. The cooling air cools the NGV and rotor blade internally by convection and then exits the NGV and rotor blade through many small exterior holes 3 to form cooling films over the external aerofoil surfaces.
Implementation Method 4
In order to produce a high level of heat transfer on the suction surface an impingement plate 7 is inserted into the passage, holes (not shown) in the plate producing jets of cooling air which impinge on the suction surface exterior wall at a relatively high velocity.
Data Source
Figure 1
Figure 2
Figure 3(a)
AI summary
A cooled aerofoil for a gas turbine engine has an aerofoil section with pressure and suction surfaces extending between inboard and outboard ends thereof. The aerofoil section includes first and second internal passages for carrying cooling air. The aerofoil section further includes a plurality of holes in the external surface of the aerofoil section which receive cooling air from the internal passages. The external holes are arranged such that cooling air exiting a first portion of the external holes participates in a cooling film extending from the leading edge of the aerofoil section over said pressure surface and cooling air exiting from a second portion of the external holes participates in a cooling film extending from the leading edge over said suction surface. The first portion of external holes receives cooling air from the first internal passage, and the second portion of external holes receives cooling air from the second internal passage. The first and second internal passages are supplied with cooling air from respective and separate passage entrances. Each entrance is located at either the inboard end or the outboard end of the aerofoil section.