Hybrid-Cycle Rocket Engine Turbine Segmentation
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Solution Overview
Problem
Current liquid propellant rocket engine technologies, such as gas generator and expander cycles, face inefficiencies due to partial propellant diversion for turbine power, resulting in lower thrust-to-weight ratios and increased atmospheric losses, as they either vaporize propellants or combust a portion for exhaust, limiting the use of propellants in generating thrust.
Innovation Solution
A hybrid rocket engine design that combines elements of gas generator and expander cycles, utilizing a gas generator to power pumps and turbines while also employing an expander cycle to increase combustion chamber pressure, allowing for higher thrust and efficient use of propellants by isolating and connecting turbines and pumps to optimize fuel and oxidizer flow, enabling operation as either a hybrid or expander cycle engine.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If a gas generator process is used to power turbines, then thrust is generated, but a portion of each propellant is combusted in the gas generator and discarded as exhaust, reducing propellant efficiency
Solution Approach 1:
The rocket engine is divided into multiple independent turbine systems (first turbine for fuel pump, second turbine for oxidizer pump, third and fourth turbines for additional pumping) that operate separately. Each turbine is powered by its own portion of propellant combustion, allowing independent optimization of each power generation pathway and reducing the need to discard propellant for a single centralized gas generator.
Solution Approach 2:
The engine design incorporates multiple turbines that can simultaneously perform different functions: the first two turbines power the main fuel and oxidizer pumps, while the third and fourth turbines provide additional pumping capability. This multi-functional turbine system allows the engine to generate sufficient power while using propellants more efficiently across multiple pathways rather than sacrificing a large portion in a single gas generator.
2Power
If propellants are vaporized by heating in the combustion chamber, then thrust is generated, but this method has limitations to the types of propellants used and the thrust generated
Solution Approach 1:
The engine design changes the thermal parameters by incorporating multiple turbine systems that can handle different temperature and pressure conditions. This allows the use of various propellant combinations with different thermal properties, as each turbine-propellant pathway can be optimized for specific propellant types, thereby increasing adaptability while maintaining thrust generation.
3Quantity of substance
If mechanical pumps are used to pressurize propellants, then propellant flow is driven into the combustion chamber, but the pumps require power from turbines that consume propellant
Solution Approach 1:
The propellant flow system is segmented into multiple independent pumping pathways, each with its own turbine-powered pump. The first pump handles fuel pressurization with its dedicated first turbine, the second pump handles oxidizer pressurization with its dedicated second turbine, and additional pumps are supported by third and fourth turbines. This segmentation reduces the propellant sacrifice ratio by distributing power generation across multiple efficient pathways rather than using a single large gas generator.
4Device complexity
If a single turbine system is used to power pumps, then the system is simpler, but the thrust-to-weight ratio is lower and atmospheric losses increase
Solution Approach 1:
The turbine system is divided into multiple independent units (first, second, third, and fourth turbines) that operate in parallel rather than as a single centralized system. Each turbine is mechanically connected to its specific pump, creating independent power pathways that reduce mechanical losses and improve overall efficiency, thereby reducing atmospheric losses while managing system complexity through modular design.
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
The hybrid rocket engine achieves higher combustion chamber pressures, resulting in increased thrust and reduced atmospheric losses, with the ability to switch between modes to optimize performance during different phases of flight, such as liftoff and later stages of launch, ensuring efficient propellant use and improved rocket propulsion.
Implementation Method 1
combusting, at a gas generator, the first portion of the fuel and the first portion of the oxidizer to power the first pump and the second pump
Implementation Method 2
The first pump is mechanically connected to a first turbine and the second pump is mechanically connected to a second turbine
Implementation Method 3
Propellants may be vaporized by heating, and thus expanding, the one or more propellants in a nozzle wall as combustion occurs in the rocket engine nozzle
Implementation Method 4
Propellants may be vaporized by heating, and thus expanding, the one or more propellants in a nozzle wall
Implementation Method 5
pumping, at a first pump, a first portion of a fuel to a gas generator
Implementation Method 6
combusting, at the combustion chamber, the second portion of the fuel and the second portion of the oxidizer to provide thrust from the nozzle
Implementation Method 7
Liquid propellant rocket engines (LPREs) use a chemical reaction between one or more propellants under pressure to generate thrust to propel a vehicle
Data Source
Figure 1
Figure 2
Figure 3A
AI summary
Systems and methods are described herein for a hybrid liquid propellant rocket engine. In an embodiment, the engine includes a first pump powered by a first turbine, a second pump powered by a second turbine, and a gas generator. An output of the gas generator is connected to the first turbine and the second turbine. The engine further includes a third pump powered by a third turbine, a fourth pump powered by a fourth turbine, and a nozzle having an expander cycle in a wall and a combustion chamber. An output of the third pump is connected to the expander cycle and an output of the wall is connected to the third turbine and the fourth turbine. An output of the fourth pump, an output of the third turbine, and an output of the fourth turbine are connected to the combustion chamber.