Rocket Engine Cooling Channels for Thermal Stratification Mixing
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Solution Overview
Problem
Liquid rocket engines face challenges in achieving efficient thermal transfer from combustion chamber walls to propellant flows, leading to incomplete mixing and potential catastrophic failures due to thermal stratification and high-pressure issues, which increase turbopump size, weight, and propellant consumption, thereby reducing payload capacity.
Innovation Solution
The implementation of variable depth cooling channels with sinusoidal forms and cross-impinged propellant injection to create toroidal vortex flows, enhancing thermal transfer and mixing of propellant within the combustion chamber, reducing thermal stratification, and simplifying fuel and oxygen injection processes.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Ease of manufacture
If conventional straight or helical cooling channels are used, then the structure is simple to manufacture, but thermal transfer efficiency is reduced due to thermal boundary layers and stratification
Solution Approach 1:
The patent applies curvature by forming cooling channels with sinusoidal or serpentine patterns instead of straight configurations. This curved path increases the surface area contact between cooling propellant and combustion chamber walls, disrupting thermal boundary layers and improving heat transfer efficiency while maintaining manufacturability through standard machining or additive manufacturing processes.
Solution Approach 2:
The patent introduces dimensional complexity by transitioning from two-dimensional planar cooling channels to three-dimensional sinusoidal paths that oscillate through the chamber wall thickness. This multi-dimensional configuration maximizes thermal contact area and promotes turbulent mixing of propellant, thereby enhancing heat rejection without significantly increasing manufacturing complexity.
2Temperature
If high pressure flows are used to improve thermal transfer, then cooling effectiveness increases, but pressure drop increases and turbopump size and weight increase
Solution Approach 1:
The sinusoidal cooling channel geometry creates continuous flow direction changes that induce secondary flows and disrupt boundary layers, enhancing heat transfer coefficients. This allows the system to achieve effective cooling at lower pressure drops, reducing the required turbopump power and weight while maintaining thermal management effectiveness.
Solution Approach 2:
The patent utilizes the hydraulic properties of liquid propellant by designing cooling channels that promote turbulent flow regimes through their curved geometry. The sinusoidal path length and cross-sectional variations create flow separation and reattachment zones that enhance mixing and heat transfer, allowing efficient cooling with reduced flow rates and lower pump requirements.
3Temperature
If cooling channels provide maximum heat flux rejection, then wall temperature control improves, but pressure drop of propellant increases
Solution Approach 1:
The sinusoidal channel configuration optimizes the balance between heat transfer area and flow resistance by using gentle curved paths rather than sharp bends. This curvature design maintains lower pressure drops while maximizing thermal contact area, achieving effective wall temperature control with minimal propellant pressure loss.
4Reliability
If turbopump pressurization is increased to maintain adequate flow, then cooling reliability improves, but propellant consumption increases and payload weight decreases
Solution Approach 1:
The sinusoidal cooling channels promote continuous turbulent mixing and prevent thermal boundary layer formation, maintaining high heat transfer coefficients throughout the channel length. This ensures reliable cooling performance at lower flow rates, reducing propellant consumption while preventing vaporization and maintaining adequate cooling margins.
Solution Approach 2:
The continuous sinusoidal path ensures uninterrupted thermal contact between cooling propellant and combustion chamber walls along the entire channel length. This continuous heat extraction action prevents localized hot spots and maintains stable cooling performance without requiring excessive flow rates or pressure fluctuations, thereby reducing propellant consumption while ensuring reliability.
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
This solution increases payload capacity by reducing turbopump size and propellant consumption, improving thermal transfer efficiency, and promoting stable combustion, while also simplifying the construction and reducing costs of the injection system.
Implementation Method 1
High pressure flows through cooling channels have thermal boundary layers that reduce thermal transfer across the flow from an inner surface proximate the combustion chamber to an outer surface proximate the external environment. The boundary layers stratify temperatures across the flow so that higher temperature fuel located proximate the combustion chamber does not mix efficiently with lower temperature fuel of the rest of the flow.
Implementation Method 2
propellant, such as kerosene fuel and in some cases liquid oxygen, flows through cooling channels formed in the combustion chamber walls to remove thermal energy in the wall next to the combustion chamber and reject the thermal energy through the wall to the external environment.
Implementation Method 3
Variable depth cooling channel surfaces destroy propellant flow thermal stratification by mixing heated layers of propellant from lateral surfaces of a cooling channel proximate a combustion chamber interior wall with cooler layers of propellant distal the combustion chamber interior wall. Reduced propellant flow thermal stratification increases the temperature gradient between the lateral surface proximate the combustion chamber and propellant within the cooling channel relative to the temperature gradient provided by an unmixed propellant flow, thus increasing thermal transfer from the combustion chamber wall material to the propellant.
Data Source
AI summary
A liquid rocket engine cools a thruster body by pumping propellant through cooling channels integrated in the thruster body between internal and external surfaces. One or more of the cooling channel surfaces has a variable depth along a thrust axis to mix propellant flow and destroy thermal stratification, such as a depth that varies with a repeated contiguous sinusoidal form along the thrust axis. Fuel passed through the cooling channels injects from the combustion chamber wall towards a central portion of the combustion chamber to cross impinge with oxygen injected at the combustion chamber head so that a toroidal vortex forms to enhance propellant mixing. The cooling channels are formed by milling at the combustion body exterior, filling the channels with wax, then electroplating metal to enclose the channels.