Spacecraft Heat Exchanger Thruster for High-Isp Propulsion
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Solution Overview
Problem
Existing spacecraft propulsion systems face inefficiencies and complexities, such as high propellant mass requirements, low thrust levels, and operational complexities, particularly in combustion-based and electric propulsion systems, limiting their maneuverability and suitability for various space missions.
Innovation Solution
A non-combustive rocket propulsion system utilizing a heat exchanger that transfers heat from a working fluid to a propellant, expelling it through a nozzle to generate thrust, with the heat exchanger manufactured via additive manufacturing using refractory metals like niobium alloys, enabling high-temperature propellant heating without combustion.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Force
If combustion-based propulsion systems are used, then high thrust is achieved, but propellant mass requirement increases and specific impulse decreases
Solution Approach 1:
The patent changes the thermal parameter of the propellant by heating it to high temperatures (above 1000°C) through heat transfer from a working fluid in a heat exchanger, rather than using combustion. This parameter change enables the propellant to achieve high specific impulse (above 400 seconds) while maintaining sufficient thrust, resolving the contradiction between thrust and propellant mass requirement
Solution Approach 2:
The patent replaces the chemical combustion mechanism with a thermal heat transfer mechanism. Instead of combusting the propellant to generate thrust, the system uses a heat exchanger to transfer thermal energy from a working fluid to the propellant, then expels the heated propellant through a nozzle. This substitution eliminates the need for separate oxidizer systems and reduces propellant mass while maintaining thrust capability
2Weight of moving object
If cold gas propulsion systems are used, then system weight is reduced, but thrust level becomes insufficient for orbit-raising and deorbiting maneuvers
Solution Approach 1:
The patent changes the temperature parameter of the propellant from cold gas temperatures to high temperatures (above 1000°C) through heat exchange with a working fluid. This parameter change increases the specific impulse to above 400 seconds and generates sufficient thrust for orbit-raising and deorbiting maneuvers, while the system remains lighter than combustion-based systems because it eliminates heavy combustion chambers, oxidizer tanks, and associated safety systems
3Quantity of substance
If electric propulsion systems are used, then specific impulse is increased, but thrust capability becomes insufficient for significant in-orbit maneuvers and electrical power requirements increase
Solution Approach 1:
The patent changes the thermal state of the propellant by heating it to high temperatures (above 1000°C) using a heat exchanger system that can be powered by solar arrays or radioisotope thermal generators. This thermal parameter change enables the system to achieve specific impulse above 400 seconds comparable to electric propulsion, while generating much higher thrust levels suitable for orbit-raising and deorbiting maneuvers without requiring complex electrical power systems
4Force
If combustion-based systems use multiplehypergolic propellants, then efficiency and thrust are increased, but system complexity and dry mass increase due to separate fluid systems
Solution Approach 1:
The patent extracts and eliminates the oxidizer component from the propulsion system entirely. Instead of using multiple hypergolic propellants (fuel and oxidizer) that require separate storage tanks, pumps, and mixing systems, the system uses a single propellant that is heated by a working fluid in a heat exchanger. This extraction of the oxidizer function simplifies the system architecture, reduces dry mass, and eliminates the complexity of managing multiple propellant systems while maintaining thrust capability through the heated propellant expansion in the nozzle
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 achieves efficient thrust generation with reduced propellant mass and electrical power requirements, allowing for high delta-V maneuvers and operation in diverse orbital environments, including Low Earth Orbit, with improved thrust efficiency and reduced system complexity.
Implementation Method 1
the heat exchanger can transfer heat from a solar collector or other heat source to a propellant
Implementation Method 2
the heat exchanger body may be configured to transfer heat from the working fluid to propellant in the propellant channels
Implementation Method 3
The heated propellant can be expelled through a nozzle, producing thrust
Implementation Method 4
The nozzle may be configured to expel the propellant
Implementation Method 5
a solar collector may be configured to capture solar energy and heat the working fluid before it enters or after it exits the heat exchanger
Implementation Method 6
a solar collector may be configured to capture solar energy
Implementation Method 7
the heat exchanger may be constructed via an additive manufacturing process
Data Source
AI summary
The present disclosure relates to systems, devices, and methods for spacecraft propulsion. In an embodiment, the present disclosure relates to an apparatus comprising a heat exchanger body defining a plurality of propellant channels configured to contain a propellant, a central cavity configured to contain a working fluid and fluidically connected to a plurality of working fluid channels that extend along a radial dimension of the apparatus, and a nozzle fluidically connected to the plurality of propellant channels and configured to expel the propellant.