End-Burning Hybrid Propulsion for CubeSat Restartable Ignition
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Solution Overview
Problem
Current hybrid rocket systems face challenges with ignition reliability, safety hazards, and inefficient combustion due to the use of pyrotechnic igniters, which are one-shot devices and susceptible to electromagnetic radiation, limiting their application in small spacecraft like CubeSats, and traditional designs have high aspect ratios that reduce volumetric efficiency.
Innovation Solution
A low-thrust, end-burn hybrid propulsion system with a housing, electrodes, and a solid-grain fuel material that uses radiative heat transfer for combustion, allowing multiple restarts and efficient thrust generation between 0.3 and 1.2 N with specific impulse greater than 100 seconds, utilizing materials like ABS and oxidizers such as GOX.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If pyrotechnic igniters are used to initiate combustion, then sufficient heat for pyrolysis and combustion activation is provided, but the system becomes susceptible to electromagnetic radiation hazards and loses restart capability
Solution Approach 1:
The patent replaces pyrotechnic igniters with an electrical discharge ignition system. electrodes are positioned to create an electric arc across the fuel grain surface, initiating combustion through electrical energy rather than chemical pyrotechnic reactions. This substitution eliminates electromagnetic radiation hazards associated with pyrotechnics while enabling controlled restart capability through electrical switching.
Solution Approach 2:
The patent changes the ignition method from high-temperature pyrotechnic initiation to controlled electrical discharge. The electrical system allows precise control of ignition parameters including voltage, current, and pulse duration, enabling reliable ignition without the hazards of pyrotechnic materials and facilitating multiple restart cycles.
2Force
If traditional core-burning hybrid rocket designs are used, then thrust levels are achieved, but the length-to-diameter ratio becomes excessively high reducing volumetric efficiency
Solution Approach 1:
The patent transitions from a traditional core-burning cylindrical geometry to an end-burning configuration where combustion occurs at the closed end of the fuel grain. This dimensional change in combustion geometry allows the thrust vector to align with the longitudinal axis while maintaining a compact, low aspect ratio fuel grain suitable for CubeSat volume constraints.
Solution Approach 2:
Instead of combustion occurring at the center core and progressing outward, the patent inverts the combustion approach by initiating burning at the closed end of the fuel grain. The combustion wave propagates through the fuel grain in reverse compared to traditional designs, achieving thrust generation in a compact configuration with improved volumetric efficiency.
3Object-affected harmful factors
If propellant materials are made less volatile for safety, then toxicity and explosiveness are reduced, but ignition difficulty increases
Solution Approach 1:
The patent replaces pyrotechnic ignition with electrical discharge ignition, which provides precise control over energy delivery to the fuel surface. The electrical arc concentrates energy at the ignition point, achieving reliable ignition of low-volatility fuels without requiring highly volatile or toxic propellant formulations.
Solution Approach 2:
The patent applies ignition energy locally at the fuel surface through positioned electrodes rather than requiring bulk heating. The electrical discharge creates a focused arc that locally pyrolyzes the fuel surface, initiating combustion in low-volatility materials without needing to heat the entire propellant mass, thus enabling safe, easy-to-ignite formulations.
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 provides a safe, reliable, and efficient propulsion solution for small spacecraft, enabling multiple ignition cycles and improved volumetric efficiency, filling the technology gap for CubeSats with thrust levels between tens of millinewtons and five Newtons.
Implementation Method 1
Combustion of hybrid propellant must be initiated by an igniter that provides sufficient heat to cause pyrolysis of the solid fuel grain at the head end of the motor, while simultaneously providing sufficient residual energy to overcome the combustion activation energy to initiate combustion.
Implementation Method 2
provides sufficient heat to cause pyrolysis of the solid fuel grain at the head end of the motor
Implementation Method 3
the combustion chamber is configured to combust the solid-grain fuel material and oxidizer in a manner dominated by radiative heat transfer
Implementation Method 4
delivering a flow of oxidizer through the oxidizer port and into the combustion chamber, and 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.