Investigation Of Matching Process Between Intake Air And Combustion In The Rocket Based Combine Cycle Engine

Rocket based combined cycle(RBCC)integrates many thermal cycle,having many technical characteristics such as wide operating range,large flight envelope,excellent comprehensive performance,which is suitable for space transportation,hypersonic flight vehicle and becomes new scheme for future aerospace propulsion.Wider flight range puts forward higher requirements for the stability of component matching.Engine design needs to consider the multi-mode compatibility,high impulse and large thrust requirements in the climbing process of the aircraft.Moreover,owing to the trajectory requirement,propulsion vehicle often operates under large attack angle and complex dynamic pressure flight conditions.It is extremely difficult to ensure high performance in each flight condition.Therefore,in order to improve the stability and efficiency of RBCC engine for wide flight Mach number,matching relationship and heat release method needs to be researched.In this dissertation,taking into account a central-strut RBCC engine which adapt to the wide operation range as the research object,investigation on the area ratio of RBCC components to the performance of different modes was carried out,using research methods contains the thermodynamic analysis and three dimensional numerical simulation.A RBCC flow path operating wide Mach range was designed to analyze the matching process between heat release and area ratio.The matching difficulty between intake and combustion was solved.The method of thermal adjustment in a wide operating range of RBCC engine was established.The performance of full flowpath in Ma2-7,±4° attack angle and 50 kPa?100kPa dynamic pressure range were achieved based on combustion in different zone.Finally,the operating stability and efficiency were verified by direct-connect tests and freejet tests.The main research contents and conclusions were as follows:(1)Integrated analysis of RBCC flowpath was conducted at the flight condition from Ma2 to Ma7.The matching relationship between intake air and combustion method was compared.Results showed that the increasing Mach of intake air before the pre-combustion shock resulted in inlet pressure ratio increase during shock train's forwarding process.This process balanced the pressure difference intake air and gas.Concentrated heat release could force thermal choking emerged,which accelerate gas and decrease its static pressure.This process realized the pressure matching between combustor and nozzle.With flight Mach increase,thermal choking should be put forward to develop the work capacity of gas,stabilizing flame and increasing combustion efficiency.(2)Flowpath design based on the idea of dual-ramjet combustion was inefficient to meet the requirement of wide range operation of the RBCC engine.Based on thermal cycle analysis,the area ratio of typical component is optimized and the flowpath structure is designed.An inlet with variable contraction ratio and a unilateral expansion combustor were adopted.After integrated numerical simulation,multiple fuel injection and compound flame stabilization realized 850s?1200s impulse performance at the flight condition from Ma3 to Ma6.Synthesizing the performance results of Ma2 and Ma7,this RBCC flowpath could satisfy the requirement of overall cruise missiles.(3)Heat releasing method in the RBCC flowpath was optimized which improved the matching between intake and combustion in a wide flight Mach range.Main combustion zone should be carried out in the isolator forepart.Releasing heat in advance could increase combustion temperature,length and efficiency in the scramjet mode.The method of concentrated combustion in a large area ratio section should be adopted to ensure thermal choking built in the rear part of the combustor and improve gas expansion in the nozzle.Based on the heat release rate,traditional empirical formula was developed and prediction model of heat releasing distribution in the RBCC engine was offered.The accuracy of this model was verified by the comparison to the direct-connect test data.(4)The freejet tests of full flowpath were performed from Ma4 to Ma5,which verified the combustor performance and matching capability between different components.Results showed that high pressure range and peak value consistent with the numerical simulation.Heat releasing method was reliable and high performance was realized.Inlet operated well under high-pressure combustion in different mode.Combustion efficiency could be developed by moving heat release forward with intake velocity increasing.RBCC engine operated with high performance operating in low speed mode by thermal throat helping gas expansion.(5)Effect of flight conditions on the combustion method was analyzed.The attack angle and height characteristic were obtained from Ma3 to Ma6,with attack angle ±4° and dynamic pressure from 50 KPa to 100 KPa.Pressure ratio increasing of inlet caused by attack angle or flight dynamic pressure developing could not match combustion pressure affected by injection increase.Heat release should be decreased in the isolator forepart and moved to combustor with larger expansion.On the contrary,main combustion zone should be moved forward to make up the combustion efficiency and pressure deficiency caused by intake mass decrease when the flight attack angle or dynamic pressure were too low.(6)Dynamic research on the matching were carried out to analyze the flow field in the inlet unstarting process.The location limit of pre-combustion shock train and modified momentum ratio were propsed to judge the matching reliability between intake air and combustion.The adjustment methods to recover inlet operation were also offered.Heat-release adjusting could increase intake's pressure,compressing backward shock wave.The flow separation would be put backward and sucked by the bleed block zone.The internal compression shock system could be reconstructed and inlet recovered,which offered design basis for fuel supply system.