Supersonic Missile Turbojet Engine with Variable Nozzle
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Solution Overview
Problem
Current gas turbine engines for supersonic missiles are limited by size, weight, and complexity, requiring increased propulsion thrust while minimizing weight and cost, and must address thermal management and efficient propulsion across subsonic, transonic, and supersonic flight speeds.
Innovation Solution
A compact turbojet engine design incorporating a core engine, an afterburner, and a converging-diverging exhaust nozzle with a controller for scheduling fuel flow, featuring a single-stage high-pressure turbine, minimal compressor stages, and variable stator vanes for optimized airflow and thrust, along with a simplified exhaust system and integrated power generation.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Weight of moving object
If a compact turbojet engine design is used for supersonic missiles, then engine size and weight are minimized, but propulsion thrust capability must be increased to maintain performance across subsonic, transonic, and supersonic flight speeds
Solution Approach 1:
The patent implements a variable geometry exhaust nozzle with movable vanes that can adjust their angle and position dynamically during flight. The nozzle includes a converging section with variable angle vanes and a diverging section with adjustable vanes that change configuration based on flight regime (subsonic, transonic, supersonic), allowing the same compact engine to deliver optimized thrust across different speed ranges without increasing engine size or weight
Solution Approach 2:
The engine employs an afterburner system that can be activated or deactivated based on flight conditions, and the exhaust nozzle geometry parameters (area ratio, angles) are changed dynamically. The controller adjusts fuel flow to the afterburner and positions of exhaust vanes to optimize the thrust-to-weight ratio at different Mach numbers, enabling high thrust from a compact configuration
2Device complexity
If a simple engine configuration is used for missile applications, then device complexity is reduced, but thermal management becomes more challenging at supersonic speeds
Solution Approach 1:
The patent extracts and separates the afterburner system from the main core engine, allowing the afterburner to be activated only when thermal management permits and when additional thrust is needed. The afterburner can be isolated and deactivated during supersonic flight when the core engine temperature is already high, simplifying the overall thermal management strategy while maintaining the ability to provide high thrust when required
Solution Approach 2:
The variable geometry exhaust nozzle with movable vanes provides dynamic thermal management by adjusting the exhaust flow patterns and heat dissipation characteristics. The vanes can be positioned to optimize cooling airflow through the nozzle structure during high-temperature supersonic operation, actively managing thermal loads without adding complex external cooling systems
3Power
If fuel flow is scheduled for afterburner operation during transonic flight, then propulsion thrust is increased, but engine complexity and control requirements increase
Solution Approach 1:
The controller implements feedback control by monitoring flight parameters (Mach number, altitude, thrust requirements) and automatically adjusting afterburner fuel flow and exhaust nozzle geometry accordingly. The system uses sensor inputs to modulate the afterburner ignition and fuel injection rates, maintaining optimal thrust while simplifying pilot control requirements
Solution Approach 2:
The controller is programmed with predetermined fuel flow schedules and nozzle configuration parameters for different flight regimes. Before transonic flight conditions are reached, the system pre-positions the afterburner and exhaust vanes according to predicted requirements, smoothing out control transitions and reducing real-time control complexity
Data Source
AI summary
A turbojet engine includes a core engine, an afterburner, and a converging-diverging exhaust nozzle in serial flow communication. A controller is operatively joined to the core engine and afterburner and configured for scheduling fuel thereto for operating the afterburner dry during subsonic flight operation of the engine, wet during transonic flight, and dry during supersonic flight.