Axi-Centrifugal Compressor Pressure Distribution
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Solution Overview
Problem
Conventional gas turbine propulsion systems for smaller aircraft face challenges in achieving high pressure rise with fewer stages, lighter weight, and lower cost without compromising compressor stability and durability, as they often require large and heavy centrifugal compressor sections due to differing optimal operating regimes and physical characteristics of axial and centrifugal compressors.
Innovation Solution
An axi-centrifugal compressor configuration is implemented, where the pressure rise is proportionally distributed between axial and centrifugal compressor sections, with a tuning factor of 2.8 to 4.5 and a loading factor of 0.6 to 0.8, minimizing aerodynamic over-loading and optimizing each axial stage to achieve a high pressure rise with fewer stages, reducing weight and cost.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Stress or pressure
If the number of compressor stages is increased to achieve high pressure rise, then the pressure ratio is improved, but the volume, weight, and cost of the engine increase
Solution Approach 1:
The patent changes the pressure ratio distribution parameters between axial and centrifugal sections, optimizing the tuning factor (ratio of pressure ratios) and loading factor (pressure ratio per stage) to achieve high overall pressure ratio with fewer stages. This parameter optimization allows the compressor to deliver the same pressure rise with reduced stage count, thereby reducing weight and size.
2Stress or pressure
If the number of compressor stages is increased to achieve high pressure rise, then the pressure ratio is improved, but the volume and length of the engine increase
Solution Approach 1:
By optimizing the pressure ratio distribution and loading factor parameters, the patent achieves high pressure ratio with fewer axial stages, directly reducing the axial length of the compressor section.
3Stress or pressure
If a large centrifugal compressor section is used to achieve high pressure rise, then the pressure ratio is improved, but the diameter and weight of the centrifugal stage increase
Solution Approach 1:
The patent optimizes the tuning factor (ratio of pressure ratios between axial and centrifugal sections) to reduce the pressure ratio requirement of the centrifugal section. This allows for a smaller impeller diameter while maintaining the same overall pressure rise.
Solution Approach 2:
The patent segments the total pressure rise requirement between axial and centrifugal sections, with the axial section handling a larger portion (optimized tuning factor of 2.8-4.5). This segmentation reduces the burden on the centrifugal section, allowing for a more compact design.
4Productivity
If aerodynamic over-loading is increased in axial stages to reduce stage count, then the productivity is improved, but boundary layer separations, low efficiency, and poor compressor operability occur
Solution Approach 1:
The patent optimizes the loading factor (pressure ratio per axial stage) to a specific range (0.6-0.8) that balances productivity and reliability. This parameter optimization ensures high pressure rise per stage while avoiding aerodynamic over-loading that would cause boundary layer separation and poor operability.
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 configuration results in a compact, efficient compressor with a shorter length, lighter weight, and lower cost, while maintaining high performance and stability, achieving the desired pressure rise with fewer stages and a smaller centrifugal stage diameter.
Implementation Method 1
The fluid is compressed along the flow path in an axial compressor section having one or more axial stages downstream from the first inlet to a first exit in the axial compressor section at a first outlet pressure. A first pressure ratio (PRax) is affected across the axial compressor section.
Implementation Method 2
The fluid is compressed along the flow path in a centrifugal compressor section from the second inlet to a second exit. A second pressure ratio is affected across the centrifugal compressor section.
Data Source
Figure 1
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AI summary
Methods and apparatus are provided for an axi-centrifugal compressor in a gas turbine engine for a business aviation or rotorcraft propulsion unit. The compressor includes an axial compressor section operable to affect a first pressure ratio along the flow path between a compressor inlet and a first section exit, and a centrifugal compressor section operable to affect a second pressure ratio along the flow path between a second section inlet and the compressor exit. The pressure rise across the axial and centrifugal compressor section is configured to have a tuning factor is in a range between 2.8 and 4.5 and a loading factor in a range between 0.6 and 0.8.