Rocket Engine Combustion Chamber Dual-Circuit Regenerative Cooling
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Solution Overview
Problem
Existing combustion chamber structures for rocket engines face challenges in withstanding high temperatures and pressures, with known cooling methods being expensive or leading to power loss, and regenerative cooling being insufficient for smaller thrust classes, resulting in local overheating and incomplete cooling.
Innovation Solution
A combustion chamber structure with two separate cooling circuits, each comprising a manifold and coolant channels extending along the entire hot gas wall, using the total mass flow of two fluidly isolated coolants for efficient heat dissipation, eliminating the need for additional cooling methods like radiation or ablative cooling.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Device complexity
If regenerative cooling using one propellant is used, then the cooling system is simple, but the cooling capacity is insufficient for small thrust classes leading to local overheating
Solution Approach 1:
The single cooling circuit is segmented into two separate cooling circuits, each with its own manifold and coolant channels. This segmentation increases the total cooling capacity by utilizing both propellants (fuel and oxidizer) as coolants independently, thereby preventing local overheating in small thrust class engines while maintaining system effectiveness.
2Temperature
If two types of heat sinks are provided in specified areas, then temperature resistance is improved, but cooling gaps appear in axial direction
Solution Approach 1:
The two separate cooling circuits operate continuously and simultaneously along the entire axial length of the combustion chamber. The coolant channels extend from the first longitudinal end to the second longitudinal end without interruption, ensuring continuous heat dissipation and eliminating cooling gaps that would occur with alternating heat sink arrangements.
3Temperature
If ablative or film cooling is used, then temperature resistance is improved, but power loss occurs
Solution Approach 1:
The cooling system uses the propellants themselves (fuel and oxidizer) as coolants in a regenerative cooling process. The propellants absorb heat from the combustion chamber walls through the coolant channels and are then fed into the combustion chamber, converting harmful heat into useful cooling energy without any loss of propellant mass or power.
4Temperature
If radiation cooling with highly expensive materials is used, then temperature resistance is improved, but manufacturing cost increases
Solution Approach 1:
Instead of using expensive radiation cooling materials, the patent employs a regenerative cooling system that uses the propellants themselves as coolants. This approach achieves effective temperature control using readily available, inexpensive propellant materials, significantly reducing manufacturing costs while maintaining cooling effectiveness.
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 design provides effective and uniform cooling across the entire axial length of the combustion chamber, reducing manufacturing costs and power loss, while ensuring sufficient cooling for engines in small thrust classes without creating cooling gaps or overheating risks.
Implementation Method 1
the generated heat of the combustion process can be spread over the first and second coolants of the two separate cooling circuits thereby using the total mass flow of the two fluidly isolated coolants for dissipating the heat
Implementation Method 2
a combustion chamber structure is usually cooled thereby dissipating the heat generated by the combustion process from the combustion chamber structure
Data Source
AI summary
A combustion chamber structure for a rocket engine includes a hot gas wall (12) that surrounds a combustion chamber (40) and has a plurality of first coolant channels (50) and a plurality of second coolant channels (52). The plurality of first (50) and second (52) coolant channels extend from a first longitudinal end (16) of the hot gas wall (12) to a second longitudinal end (18) of the hot gas wall (12) opposite to the first longitudinal end (16). The combustion chamber structure (10) further comprises a first manifold (20) forming a first coolant chamber (30) and a second manifold (22) forming a second coolant chamber (32) being fluidly separated from the first coolant chamber (30). The first (20) and second (22) manifolds are provided at the first longitudinal end (16) of the hot gas wall (12) and extend in a circumferential direction of the hot gas wall (12). The first coolant chamber (30) is fluidly connected to each of the plurality of first coolant channels (50) and the second coolant chamber (32) is fluidly connected to each of the plurality of second coolant channels (52).