Integrated Rocket Combustor Head and Turbopump Design
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Solution Overview
Problem
Conventional rocket engines with turbopumps are bulky, heavy, and complex, requiring additional systems for propellant management and leakage prevention, and have increased manufacturing difficulties due to the need for multiple pipes and separate pumps for oxidizers and fuels.
Innovation Solution
An integrated combustor head and turbopump design where oxidizer and fuel pumps are coupled with turbines, sharing a rotating shaft and connected through gears to minimize weight and volume, with regenerative cooling channels for efficient vaporization and combustion, eliminating the need for separate heat exchangers and reducing pipe complexity.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Weight of moving object
If conventional separate turbopump and combustor head design is used, then functional reliability is maintained, but weight and volume increase
Solution Approach 1:
The patent integrates the turbopump and combustor head into a single unified structure. The pump housing serves dual functions as both the turbopump casing and the combustor head, eliminating the need for separate components and reducing overall weight while maintaining functional reliability through careful design of integrated flow paths and combustion chambers.
2Volume of moving object
If multiple pipes connecting oxidizer tank, fuel tank, and injector manifold are used, then propellant delivery function is achieved, but volume and weight increase
Solution Approach 1:
The patent consolidates multiple propellant delivery functions into the integrated pump housing structure. The housing incorporates internal passages and chambers that serve as both pump casings and combustor components, eliminating the need for extensive external piping and reducing manufacturing steps while achieving compact volume.
3Use of energy by moving object
If separate oxidizer pump and fuel pump operated by one shaft are used, then power efficiency is improved, but propellant mixing risk increases
Solution Approach 1:
The patent divides the single driving shaft into two independent shafts, with each pump (oxidizer pump and fuel pump) having its own dedicated shaft. This segmentation physically separates the propellant delivery systems, eliminating the risk of propellant mixing while maintaining power efficiency through direct coupling of each pump to its own turbine driver.
4Reliability
If additional IPS purge system is added for blocking propellant mixing, then safety is improved, but device complexity increases
Solution Approach 1:
The patent removes the need for additional IPS purge systems by fundamentally redesigning the propellant delivery architecture. By using separate shafts for oxidizer and fuel pumps from the outset, the design inherently prevents propellant mixing, making additional safety systems unnecessary and reducing overall device complexity.
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 reduces the weight and volume of the turbopump, simplifies manufacturing, enhances propellant management, and increases engine efficiency by directly integrating the turbopump with the combustor head, while maintaining a constant combustion ratio and preventing propellant mixing.
Implementation Method 1
a fuel is vaporized when the fuel is heated in a regenerative cooling channel
Implementation Method 2
an oxidizer and a fuel are injected into the thrust chamber 10 and combusted therein
Implementation Method 3
the gas generator 30 receives a portion of the oxidizer and fuel and combuts the oxidizer and fuel to operate the turbine
Implementation Method 4
a turbopump 20 for supplying the oxidizer and fuel to the thrust chamber 10
Implementation Method 5
A rocket engine is a device which obtains propulsion by injecting an oxidizer and a fuel, which are propellants, into a combustor and combusting the oxidizer and the fuel
Data Source
AI summary
The present disclosure relates to a rocket engine, and more particularly, a rocket engine with an integrated combustor head and turbopump in which a turbopump of the rocket engine is formed integrally with a combustor head.