Aerospace Engine with Augmenting Turbojet Flowpath Control
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Solution Overview
Problem
Gas turbine engines face inefficiencies due to fixed rotational relationships between sections, leading to tradeoffs in fan and core efficiencies during different flight conditions, and thrust augmenters like afterburners are fuel-inefficient and require suboptimal pressure ratio adjustments.
Innovation Solution
A gas turbine engine system with a flowpath control mechanism that allows selective restriction of fluid flow through a secondary engine core, enabling dynamic adjustment of engine core size and configuration, including a primary reverse-flow core and a secondary augmenting core with axial or reverse-flow configurations, shared fan, compressor, and turbine sections, and optional torque coupling and flow mixing to reduce exhaust signature.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Adaptability or versatility
If a fixed rotational relationship is maintained between fan section and engine core sections, then mechanical simplicity is preserved, but fan efficiency and core efficiency cannot be optimized simultaneously for different flight conditions
Solution Approach 1:
The engine is divided into multiple independent core sections (first engine core and second engine core) that can operate independently or together. The flowpath control mechanism segments the airflow to direct it through different core sections based on flight conditions, allowing each core to be optimized for specific operating regimes without being constrained by a fixed rotational relationship with the fan section.
Solution Approach 2:
The engine configuration is made dynamic through the flowpath control mechanism that can selectively restrict or open flowpaths through different core sections. This allows the engine to adapt its effective core size and configuration in real-time based on flight conditions, transitioning between using only the first core, only the second core, or both cores together.
2Power
If thrust augmenters like afterburners are used to provide additional thrust, then high-power capability is achieved, but fuel efficiency deteriorates and pressure ratio adjustments become suboptimal
Solution Approach 1:
Instead of using a single thrust augmenter, the invention divides the thrust augmentation capability into two separate engine cores. Each core can be independently activated based on power requirements, allowing for more efficient fuel consumption by engaging only the necessary core capacity rather than always requiring afterburner-level power settings.
Solution Approach 2:
The invention changes the operational parameters by allowing independent control of multiple core sections rather than relying on afterburner combustion. This enables optimal pressure ratio adjustments for each core based on specific flight conditions, avoiding the suboptimal pressure ratios required when operating an afterburner.
3Productivity
If engine core size is reduced for cruise conditions, then fan efficiency improves, but core efficiency deteriorates at high-power modes
Solution Approach 1:
The engine dynamically adjusts its effective core size by selectively activating the first engine core, second engine core, or both cores based on power requirements. During cruise conditions, only the smaller first core may be needed, maintaining high fan efficiency. During high-power modes, the second core is activated to provide additional core capacity without requiring the fan to be oversized for maximum thrust conditions.
Solution Approach 2:
The core capability is segmented into two separate engine cores with different size characteristics. This allows the fan to be sized for optimal cruise performance with the smaller first core, while the larger second core provides additional power capability when needed, eliminating the need to oversize the fan for maximum thrust conditions.
Data Source
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AI summary
A gas turbine engine system includes a fan section, a low pressure compressor section downstream of the fan section, a first engine core downstream from the low pressure compressor section, a second engine core downstream from the low pressure compressor section, and a flowpath control mechanism configured to selectively restrict fluid flow through the second engine core. The first engine core includes a first engine core compressor section, a first engine core combustor downstream of the first engine core compressor section, and a first engine core turbine section downstream of the first engine core combustor. The second engine core includes a second engine core compressor section, a second engine core combustor downstream of the second engine core compressor section, and a second engine core turbine section downstream of the second engine core combustor.