An early TSTO fully reusable vehicle design used to 'calibrate' Stage 1 combined-cycle hypersonic propulsion systems

2000 ◽  
Author(s):  
William Escher ◽  
Carl Ehrlic
Keyword(s):  
Combined Cycle ◽  
Vehicle Design ◽  
Propulsion Systems ◽  
Hypersonic Propulsion ◽  
Stage 1

Integrated Thermal Management System Design for Advanced Propulsion System

2012 ◽  
Vol 232 ◽  
pp. 723-729
Author(s):  
De Cang Lou ◽  
Wen Guo ◽  
Zhi Guo Wang ◽  
Yong Hong Wang
Keyword(s):  
Thermal Management ◽  
Management System ◽  
Flow System ◽  
Combined Cycle ◽  
Propulsion System ◽  
Propulsion Systems ◽  
Key Technology ◽  
Hypersonic Propulsion ◽  
Thermal Management System ◽  
Aero Engines

Thermal management system (TMS) design is considered to be a key technology for advanced aero engines and supersonic or hypersonic propulsion systems. In this paper, the concepts of coupling flow and thermodynamic networks are proposed for TMS design. In this method, the propulsion system is considered to be a zero-dimensional flow system. Components, subsystems and hence the entire engine system can be modelled using some basic flow and thermodynamics networks. The platform for TMS design, ThermalM, is developed based on this model. As an example, modelling for a Turbine Based Combined Cycle (TBCC) thermal management system is described. Performance of the fuel heat exchanger in the network is discussed in detail. With the TMS design technology, performance of the advanced propulsion system can be analysed.

Aerothermodynamic analysis of combined cycle propulsion systems

1990 ◽  
Author(s):  
A. GANJI ◽  
M. KHADEM ◽  
S. KHANDANI
Keyword(s):  
Combined Cycle ◽  
Propulsion Systems

Numerical Simulation of Supersonic Combustion in Quasi One-Dimensional Flow

1991 ◽  
Author(s):  
T. Gary Yip
Keyword(s):  
Flow Model ◽  
Cone Angle ◽  
Initial Pressure ◽  
Supersonic Combustion ◽  
Propulsion Systems ◽  
One Dimensional ◽  
Hypersonic Propulsion ◽  
Dimensional Flow ◽  
Conical Shock ◽  
Shock System

Abstract Supersonic combustion induced by a two-shock system has been studied using a chemical nonequilibrium, quasi one-dimensional flow model. The combustion of stoichiometric, premixed H2-air is described by a chemistry model which consists of 11 species and 28 reactions. The freestream Mach numbers used in this calculations are 8, 10 and 12. The initial pressure is 0.01 atm and temperature 300 K. The first of the two shocks is a conical shock and the second is its reflection. Supersonic combustion has been predicted to occur at combustor pressures between 0.8 and 2.9 atmospheres, and temperatures between 1500 and 3000 K. The Mach number of the flow in the combustor is between 1.7 and 4. These combustor conditions are typical of the future hypersonic propulsion systems. The results also show the changes in the composition of the flow during the induction and heat release phases. The two-shock system is assumed to be generated by a cone. For Mach 8, 10 and 12, the minimum cone angle for generating a strong enough two-shock system to induce supersonic combustion has also been identified.

Research on control-oriented coupling modeling for air-breathing hypersonic propulsion systems

2019 ◽  
Vol 84 ◽  
pp. 143-157 ◽  
Author(s):  
Dong Zhang ◽  
Shuo Tang ◽  
Lin Cao ◽  
Feng Cheng ◽  
Fan Deng
Keyword(s):  
Propulsion Systems ◽  
Air Breathing ◽  
Hypersonic Propulsion

scholarly journals Ramjet Compression System for a Hypersonic Air Transportation Vehicle Combined Cycle Engine

Energies ◽  
2018 ◽  
Vol 11 (10) ◽  
pp. 2558 ◽  
Author(s):  
Sasha Veeran ◽  
Apostolos Pesyridis ◽  
Lionel Ganippa
Keyword(s):  
Flow Characteristics ◽  
Air Transportation ◽  
Combined Cycle ◽  
Propulsion Systems ◽  
Compression System ◽  
Ramjet Combustor ◽  
Start Up ◽  
Flight Profile ◽  
Self Powered ◽  
Transportation Vehicle

This report assesses the performance characteristics of a ramjet compression system in the application of a hypersonic vehicle. The vehicle is required to be self-powered and perform a complete flight profile using a combination of turbojet, ramjet and scramjet propulsion systems. The ramjet has been designed to operate between Mach 2.5 to Mach 5 conditions, allowing for start-up of the scramjet engine. Multiple designs, including varying ramp configurations and turbo-ramjet combinations, were investigated to evaluate their merits and limitations. Challenges arose with attempting to maintain sufficient pressure recoveries and favourable flow characteristics into the ramjet combustor. The results provide an engine inlet design capable of propelling the vehicle between the turbojet and scramjet phase of flight, allowing for the completion of its mission profile. Compromises in the design, however, had to be made in order to allow for optimisation of other propulsion systems including the scramjet nozzle and aerodynamics of the vehicle; it was concluded that these compromises were justified as the vehicle uses the ramjet engine for a minority of the flight profile as it transitions between low supersonic to hypersonic conditions.

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