Rocket Engine Cooling Channels With Sinusoidal Depth 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 increased structural weight due to high turbopump size and propellant consumption, which can result in catastrophic failures and reduced 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 propellants, thereby reducing thermal stratification and the need for pressurization within the combustion chamber.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Ease of manufacture
If conventional straight or helical cooling channels are used, then the structure is simple and easy to manufacture, but thermal transfer efficiency is reduced due to thermal boundary layers that stratify temperatures across the propellant flow
Solution Approach 1:
The patent applies curvature by transforming conventional straight or helical cooling channels into serpentine (sinuous) pathways. This curved configuration disrupts thermal boundary layers and promotes mixing of propellant flow, thereby enhancing thermal transfer efficiency from combustion chamber walls to the propellant without complicating manufacturing processes
Solution Approach 2:
The serpentine cooling channel design introduces dynamic flow patterns within the cooling system. The winding pathway creates varying flow velocities and directions that actively disrupt thermal stratification, transforming the static thermal field into a dynamic mixing process that improves heat transfer
2Temperature
If high pressure flows are used to increase thermal transfer, then cooling effectiveness improves, but thermal boundary layers stratify temperatures and reduce mixing efficiency
Solution Approach 1:
The serpentine cooling channel geometry creates curved flow paths that induce secondary flows and disrupt boundary layers. This curvature effect promotes lateral mixing of propellant, preventing thermal stratification even under high pressure conditions, thereby maintaining both cooling effectiveness and mixing homogeneity
3Reliability
If larger turbopumps are used to increase propellant pressurization, then adequate cooling flow is maintained, but turbopump size and propellant consumption increase, decreasing payload weight
Solution Approach 1:
The patent changes the geometric parameters of cooling channels from straight/helical to serpentine configurations. This parameter change increases the effective heat transfer surface area and improves thermal coupling between combustion chamber walls and propellant, allowing reduced propellant flow rates to achieve the same cooling effect, thereby reducing turbopump size and payload penalty
4Temperature
If propellant flow rate is increased to improve cooling, then thermal transfer increases, but pressure drop and turbopump power requirements increase
Solution Approach 1:
The serpentine channel design optimizes the balance between heat transfer and pressure drop by using curved pathways that enhance mixing and thermal contact time without requiring excessive flow rates. The geometry provides efficient heat transfer at moderate pressure drops
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 approach increases thermal transfer efficiency, reduces turbopump size and propellant consumption, and enhances combustion stability, resulting in increased payload capacity and improved thermal management without complex injector structures.
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
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
propellant flows through cooling channels formed in the combustion chamber walls to remove thermal energy
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.