Hypersonic Propulsion System with Rocket-SCRAMJet Integration

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Solution Overview

Problem

Current propulsion systems for hypersonic aircraft, such as SCRAMJet and combined cycle SCRAMJet systems, face limitations in operating range, efficiency, and mechanical vibrations, requiring special means for startup and ideal operating conditions, and suffer from overheating and inefficient fuel usage.

Innovation Solution

A propulsion system with a rocket-ignited supersonic combustion RAMjet (RISCRAM jet) configuration, featuring a coaxial arrangement of a bypass duct, air-breathing engine, and rocket within a containment duct, utilizing a supersonic combustor and flame pipe to efficiently convert thermal energy into kinetic energy, allowing autonomous operation from sea level to hypersonic speeds without auxiliary systems.

Engineering Contradictions & Design Principles

VSEngineering Contradiction Analysis

1Speed

If a pure SCRAMJet propulsion system is used, then the aircraft can achieve hypersonic steady state flight, but special means are required to bring the engine to optimal operational condition (transport aircraft for carrying to ideal altitude and speed)

Engineering Contradiction:

Improvehypersonic steady state speedVSAvoidcomplexity of auxiliary systems

Core Design Contradiction:

SpeedVSDevice complexity

Solution Approach 1:

The propulsion system is divided into distinct functional segments: a rocket engine for acceleration and ignition, and a SCRAMJet engine for hypersonic cruise. This segmentation allows each component to operate in its optimal regime, with the rocket providing the necessary push to reach SCRAMJet operating conditions without requiring external transport aircraft.

Inventive Principle:

Principle #1Segmentation

Solution Approach 2:

The rocket engine performs preliminary action by accelerating the aircraft to the required speed and altitude before the SCRAMJet engine takes over. This preliminary acceleration phase prepares the flight conditions necessary for supersonic combustion to occur efficiently in the SCRAMJet combustor.

Inventive Principle:

Principle #10Preliminary action

2Productivity

If supersonic steady state combustion is used, then propulsion efficiency is improved, but the choice of fuel is considerably limited (forces use of hydrogen)

Engineering Contradiction:

Improvepropulsion efficiencyVSAvoidfuel selection flexibility

Core Design Contradiction:

ProductivityVSAdaptability or versatility

Solution Approach 1:

The rocket engine acts as an intermediary system that enables the use of conventional liquid fuels in the SCRAMJet by providing the necessary supersonic flow conditions. The rocket's combustion chamber and nozzle create the high-speed flow environment that allows efficient supersonic combustion of conventional fuels like kerosene, removing the strict limitation to hydrogen.

Inventive Principle:

Principle #24Intermediary (Mediator)

Solution Approach 2:

The system changes the operational parameters by using the rocket to establish supersonic flow conditions (Mach number, pressure, temperature) in the combustor. These parameter changes enable efficient combustion of conventional liquid fuels that would not otherwise be suitable for supersonic combustion, expanding fuel selection beyond hydrogen.

Inventive Principle:

Principle #35Parameter changes

3Volume of moving object

If the combustion flame is close to the combustor walls, then compact design is achieved, but overheating of parts occurs requiring complex insulation and cooling measures

Engineering Contradiction:

Improvecombustor volumeVSAvoidcombustor wall temperature

Core Design Contradiction:

Volume of moving objectVSTemperature

Solution Approach 1:

The rocket exhaust flow acts as an intermediary thermal barrier between the SCRAMJet combustor flame and the combustor walls. The high-velocity rocket exhaust creates a protective layer that reduces direct thermal exposure to the walls, allowing closer flame positioning without excessive heating.

Inventive Principle:

Principle #24Intermediary (Mediator)

Solution Approach 2:

The system replaces complex mechanical cooling and insulation systems with a flow-based thermal management approach. The rocket exhaust flow dynamically manages the thermal environment, substituting for static thermal protection systems and reducing overall system complexity.

Inventive Principle:

Principle #28Mechanics substitution (Replace mechanical system)

4Volume of moving object

If shock waves enter into direct contact with the perimeter walls of the propulsion system, then compact configuration is achieved, but harmful vibrations are generated causing energy dispersion

Engineering Contradiction:

Improvepropulsion system volumeVSAvoidenergy dispersion through vibrations

Core Design Contradiction:

Volume of moving objectVSLoss of energy

Solution Approach 1:

The harmful shock wave interactions with the perimeter walls are extracted or removed from the system by designing the combustor geometry and flow paths to guide shock waves away from the structural boundaries. This extraction eliminates the source of harmful vibrations while preserving the compact configuration.

Inventive Principle:

Principle #2Taking out (Extraction)

Solution Approach 2:

The shock waves that would otherwise cause harmful vibrations are redirected to perform useful work within the combustor, such as enhancing mixing and combustion efficiency. The potentially harmful shock structures are converted into beneficial flow features that improve propulsion performance.

Inventive Principle:

Principle #22Blessing in disguise (Convert harm into benefit)

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

Enables efficient energy conversion, reduced mechanical vibrations, and improved thermal insulation, allowing autonomous hypersonic flight from a fixed point to high altitudes with conventional liquid fuels, enhancing propulsion efficiency and reliability.

Implementation Method 1

the flow of fluid which is greatly energised (thermal energy in particular) is subjected to an expansion through a divergent nozzle of the propulsion system converting the high thermal energy into kinetic energy

Methodology Applied

Scientific EffectShock waves: Shock Wave

Implementation Method 2

allow a combustion process between fuel and comburent

Methodology Applied

Scientific EffectCombustion: Combustion

Implementation Method 3

Methodology Applied

Scientific EffectThermal energy to kinetic energy conversion:

Implementation Method 4

a propulsion system for hypersonic aircraft having a supersonic RAM combustor which is ignited by a rocket

Methodology Applied

Scientific EffectRocket ignition: Rocket

Data Source

PatentUS10927793B2Engine for hypersonic aircrafts with supersonic combustor

Publication Date: 2021.02.23 GABALDO MARCO

AI summary

Described is a propulsion system (1) for hypersonic aircraft, having an air inlet (10) of a fluid (110), a containment duct (20) and an exhaust nozzle (30). The propulsion system (1) comprises a bypass duct (40) for a flow (100) of fluid (110), an air-breathing engine (22) and a rocket (23) configured for processing respective flows (22a, 23a) of fluid (110). The bypass duct (40), the air-breathing engine (22) and the rocket (23) are operatively associated with each other in such a way as to generate a thermodynamic-fluid interaction in a same portion of space (33) between the respective flows (40a, 22a, 23a) processed in an operating configuration of the propulsion system (1) and wherein the portion of space (33) is inside the containment duct (20).