Electrical Ignition for Hybrid Rocket Fuel
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Solution Overview
Problem
Current hybrid rocket ignition systems are limited by their inability to initiate multiple restarts, pose safety hazards, and have difficulty igniting stable fuel due to high enthalpy requirements, making them unsuitable for in-space propulsion and CubeSat applications.
Innovation Solution
A restartable hybrid rocket ignition system featuring a housing with multiple flat layers and electrodes that concentrate electrical charges to arc and ignite the fuel, allowing for multiple restarts and reducing hazards, using materials like Acrylonitrile Butadiene Styrene (ABS) and additive manufacturing for efficient thrust generation.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If pyrotechnic ignition methods are used, then combustion can be initiated, but multiple restarts cannot be achieved and safety hazards increase
Solution Approach 1:
The patent replaces the mechanical/pyrotechnic ignition system with an electrical field-based ignition system. Electrodes generate high-voltage electrical discharges to initiate combustion, eliminating the need for pyrotechnic devices. This substitution enables multiple restarts through simple electrical switching while removing the safety hazards associated with pyrotechnic materials and electromagnetic radiation.
Solution Approach 2:
The patent changes the ignition mechanism from thermal/pyrotechnic to electrical. By applying high-voltage electrical fields across electrodes positioned near the fuel surface, the system can reliably initiate combustion multiple times without the limitations of single-use pyrotechnic igniters, thereby improving reliability while maintaining safety.
2Ease of manufacture
If high enthalpy pyrotechnic charges are used to ignite stable fuel, then combustion can be initiated, but device complexity and hazards increase
Solution Approach 1:
The patent replaces complex pyrotechnic charge systems with a simple electrical field generation system. Two electrodes connected to a high-voltage power source create an electrical discharge that ignites the fuel. This substitution dramatically simplifies the igniter design while maintaining the ability to ignite stable fuels, and eliminates the need for complex pyrotechnic compositions and safety mechanisms.
3Force
If conventional hybrid rocket systems are used, then propulsion can be achieved, but in-space propulsion and CubeSat applications are not suitable
Solution Approach 1:
The patent segments the hybrid rocket system into distinct functional components: a solid fuel grain, a separate oxidizer delivery system, and an electrical ignition system. This segmentation allows each component to be independently optimized for space applications, with the electrical ignition system providing precise control suitable for in-space operations and CubeSat missions where reliability and controllability are critical.
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
The system enables safe, efficient, and repeatable ignition of hybrid rockets, overcoming the limitations of traditional pyrotechnic igniters by providing a compact, high-thrust, and moderate-specific-impulse propulsion suitable for CubeSats, with thrust levels up to 5 N and oxidizer flow rates of 5 g/s, and improved volumetric efficiency.
Implementation Method 1
The internal surface with the ridges is configured to concentrate an electrical charge upon being subjected to an electrical potential field between the at least two electrodes
Implementation Method 2
the internal surface with the ridges are configured to concentrate an electrical charge generally between the at least two electrodes upon being subjected to the electrical potential field
Implementation Method 3
igniting the combustion surface with the at least two electrodes to generate a hot-gas, fuel-rich flow through the nozzle to generate thrust
Data Source
AI summary
A hybrid propulsion system includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The electrodes extend into the cavity. The fuel material is free of an oxidizer and is positioned in the cavity. The fuel material has a combustion surface and is exposed to the electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port provides a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber may be dominated by radiative heat transfer. Combustion of the fuel material in the combustion chamber may generate thrust of no more than 5 N at an oxidizer flow rate of no more than 5 g/s.