Rocket Ignition System Continuous Flow Homogeneity
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Solution Overview
Problem
In liquid propellant rocket engines, the ignition system is only operational during the initial ignition phase, leading to inefficiencies and safety concerns due to hot gases rising in the igniter channel, which can freeze and block propellant entry, preventing re-ignition, and requires additional neutral gases for purging, impacting performance and safety.
Innovation Solution
The ignition system is integrated into the injection head and comprises a main, secondary, and tertiary ignition channel, allowing for continuous operation during the injection phase, scavenging hot gases with oxidizer and providing a homogeneous fuel and oxidizer flow, eliminating the need for neutral purge gases and enhancing combustion efficiency.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Device complexity
If the ignition system operates only during the initial ignition phase, then the system complexity is reduced, but hot gases rise in the igniter channel and freeze, blocking propellant entry and preventing re-ignition
Solution Approach 1:
The ignition system is designed to operate continuously during both the ignition phase and the stabilized injection phase, rather than shutting down after initial ignition. This continuous operation prevents hot gases from rising and freezing in the igniter channel, maintaining reliable re-ignition capability throughout the engine's operation.
Solution Approach 2:
The ignition system maintains readiness by continuing to supply igniter propellants during the stabilized phase, preparing in advance for any potential re-ignition needs. This preliminary maintenance of operational state prevents the harmful freezing effect that would occur if the system shut down completely.
2Loss of substance
If the ignition system is shut down during the stabilized injection phase, then propellant consumption is reduced, but additional neutral gases are required for purging, impacting performance
Solution Approach 1:
The ignition system uses its own propellant supply during the stabilized phase to maintain operation and prevent hot gas rise, rather than requiring external neutral purge gases. This self-service approach eliminates the need for separate purging systems and maintains performance by using combustion-supporting propellants instead of inert gases.
Solution Approach 2:
Instead of shutting down the ignition system and dealing with the harmful effect of hot gas freezing, the system converts this potential harm into a benefit by maintaining continuous operation. The continuous flow of igniter propellants prevents freezing and blocks hot gas rise, turning what would be a wasted resource into a protective function that enhances reliability.
3Productivity
If the ignition system operates continuously during the injection phase, then combustion efficiency is improved, but the device complexity increases
Solution Approach 1:
The ignition system is integrated with the main propellant supply system, merging the igniter propellant supply with the main oxidizer and fuel supply lines. This integration allows continuous operation during the stabilized phase without requiring entirely separate systems, thereby improving combustion efficiency while limiting the increase in 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 configuration improves combustion efficiency by utilizing additional oxidizer and fuel during the stabilized phase, prevents hot gases from rising in the igniter channel, and eliminates the need for neutral purge gases, enhancing overall rocket performance and safety.
Implementation Method 1
a spark plug 107 for igniting the oxidizer and fuel mixture
Implementation Method 2
where the combustion as such takes place
Implementation Method 3
a combustion chamber body 6 (where the combustion as such takes place) cooled by a cooling circuit 6bis
Data Source
Figure 1
Figure 2
Figure 3~4
AI summary
The invention relates to a rocket propulsion system comprising: an oxidizer supply system (2); a fuel supply system (3); a combustion chamber (6); a plurality of injectors (101); an ignition system (40); the injectors (101) and the ignition system (40) being in communication with the oxidizer supply system and the fuel supply system and opening into the combustion chamber (6), the injectors (101) being configured to, during an injection phase, bring oxidizer and fuel to the combustion chamber (6);the ignition system (40) being configured to, during an ignition phase, bring hot gases to the combustion chamber (6) and to, during the injection phase, bring oxidizer and fuel to the combustion chamber (6) in such a way that the flow rate of oxidizer and fuel from the ignition system (40) is homogeneous with that from the injectors (101).