Flame temperature and spectral emissivity were the important parameters characterizing the sufficient degree of fuel combustion and the particle radiative characteristics in the Rocket Based Combined Cycle(RBCC)combus...Flame temperature and spectral emissivity were the important parameters characterizing the sufficient degree of fuel combustion and the particle radiative characteristics in the Rocket Based Combined Cycle(RBCC)combustor.To investigate the combustion characteristics of the complex supersonic flame in the RBCC combustor,a new radiation thermometry combined with Levenberg-Marquardt(LM)algorithm and the least squares method was proposed to measure the temperature,emissivity and spectral radiative properties based on the flame emission spectrum.In-situ measurements of the flame temperature,emissivity and spectral radiative properties were carried out in the RBCC direct-connected test bench with laser-induced plasma combustion enhancement(LIPCE)and without LIPCE.The flame average temperatures at fuel global equivalence ratio(a)of 1.0b and 0.6 with LIPCE were 4.51%and 2.08%higher than those without LIPCE.The flame combustion oscillation of kerosene tended to be stable in the recirculation zone of cavity with the thermal and chemical effects of laser induced plasma.The differences of flame temperature at a=1.0b and 0.6 were 503 K and 523 K with LIPCE,which were 20.07%and42.64%lower than those without LIPCE.The flame emissivity with methane assisted ignition was 80.46%lower than that without methane assisted ignition,due to the carbon-hydrogen ratio of kerosene was higher than that of methane.The spectral emissivities at 600 nm with LIPCE were 1.25%,22.2%,and 4.22%lower than those without LIPCE at a=1.0a(with methane assisted ignition),1.0b(without methane assisted ignition)and 0.6.The effect of concentration in the emissivity was removed by normalization to analyze the flame radiative properties in the RBCC combustor chamber.The maximum differences of flame normalized emissivity were 50.91%without LIPCE and 27.53%with LIPCE.The flame radiative properties were stabilized under the thermal and chemical effects of laser induced plasma at a=0.6.展开更多
To reduce the drag generated by the recirculation flow at the rocket base in a RocketBased Combined Cycle(RBCC)engine operating in the ramjet/scramjet mode,a novel annular rocket RBCC engine based on a central plug co...To reduce the drag generated by the recirculation flow at the rocket base in a RocketBased Combined Cycle(RBCC)engine operating in the ramjet/scramjet mode,a novel annular rocket RBCC engine based on a central plug cone was proposed.The performance loss mechanism caused by the recirculation flow at the rocket base and the influence of the plug cone configuration on the thrust performance were studied.Results indicated that the recirculation flow at the rocket base extended through the entire combustor,which creates an extensive range of the"low-kineticenergy zone"at the center and leads to an engine thrust loss.The plug cone serving as a surface structure had a restrictive effect on the internal flow of the engine,making it smoothly transit at the position of the large separation zone.The model RBCC engine could achieve a maximum thrust augmentation of 37.6%with a long plug cone that was twice diameter of the inner isolator.However,a shorter plug cone that was half diameter of the inner isolator proved less effective at reducing the recirculation flow for a supersonic flow and induced an undesirable flow fraction that diminished the thrust performance.Furthermore,the effectiveness of the plug cone increased with the flight Mach number,indicating that it could further broaden the operating speed range of the scramjet mode.展开更多
Ejector mode is a unique and critical phase of wide-range rocket-based combined cycle(RBCC)engine.In this paper,a quasi-one-dimensional thermodynamic performance modeling method,with more detailed model treatments for...Ejector mode is a unique and critical phase of wide-range rocket-based combined cycle(RBCC)engine.In this paper,a quasi-one-dimensional thermodynamic performance modeling method,with more detailed model treatments for the inlet-diffuser system,pri-mary/secondaryflow interaction,and pressure feedback matching,was developed for operating characteristics studies and multi-objective optimization analysis of the ejector mode of an actual RBCC engine.A series of three-dimensional simulations of separate inlet and fullflowpath was completed to validate the modeling study and provide further insight into the operating charac-teristics.The primary/secondary equilibrium pressure ratio functions a significant effect on ejector mode performance,a higher performance augmentation can be obtained by lower rocket pressure ratio,larger mixing section area ratio,smaller throttling throat and higher equivalence ratio,within an appropriate range.The positive performance augmentation can be realized at lowflight Mach conditions,the coordination and trade-off relationships between specific im-pulse,performance augmentation ratio and thrust-to-area ratio during ejector mode are present by the Pareto-front from MOP analysis.It is further verified by CFD simulation that,the operating back-pressure at the exit of inlet-diffuser system functions a decisive influence on the airbreathing characteristics,the pressure feedback and matching should be well-controlled for secondaryflowrate and performance augmentation.The thermodynamic modeling analysis re-sults are basically consistent with those of numerical simulation,to validate the rationality and effectiveness of the modeling method.展开更多
Combined-cycle pulse detonation engines are promising contenders for hypersonic propulsion systems.In the present study,design and propulsive performance analysis of combined-cycle pulse detonation turbofan engines(PD...Combined-cycle pulse detonation engines are promising contenders for hypersonic propulsion systems.In the present study,design and propulsive performance analysis of combined-cycle pulse detonation turbofan engines(PDTEs)is presented.Analysis is done with respect to Mach number at two consecutive modes of operation:(1)Combined-cycle PDTE using a pulse detonation afterburner mode(PDA-mode)and(2)combined-cycle PDTE in pulse detonation ramjet engine mode(PDRE-mode).The performance of combined-cycle PDTEs is compared with baseline afterbuming turbofan and ramjet engines.The comparison of afterburning modes is done for Mach numbers from 0 to 3 at 15.24 km altitude conditions,while that of pulse detonation ramjet engine(PDRE)is done for Mach 1.5 to Mach 6 at 18.3 km altitude conditions.The analysis shows that the propulsive performance of a tubine engine can be greatly improved by replacing the conventional afterbumer with a pulse detonation afterburner(PDA).The PDRE also outperforms its ramjet counterpart at all flight conditions considered herein.The gains obtained are outstanding for both the combined-cycle PDTE modes compared to baseline turbofan and ramjet engines.展开更多
基金supported by the National Natural Science Foundation of China (Grant Nos.52276185,52276189 and 51976057)the Fundamental Research Funds for the Central Universities (Grant No.2021MS126)+1 种基金the Natural Science Foundation of Jiangsu Province (Grant No.BK20231209)the Proof-of-Concept Project of Zhongguancun Open Laboratory (Grant No.20220981113)。
文摘Flame temperature and spectral emissivity were the important parameters characterizing the sufficient degree of fuel combustion and the particle radiative characteristics in the Rocket Based Combined Cycle(RBCC)combustor.To investigate the combustion characteristics of the complex supersonic flame in the RBCC combustor,a new radiation thermometry combined with Levenberg-Marquardt(LM)algorithm and the least squares method was proposed to measure the temperature,emissivity and spectral radiative properties based on the flame emission spectrum.In-situ measurements of the flame temperature,emissivity and spectral radiative properties were carried out in the RBCC direct-connected test bench with laser-induced plasma combustion enhancement(LIPCE)and without LIPCE.The flame average temperatures at fuel global equivalence ratio(a)of 1.0b and 0.6 with LIPCE were 4.51%and 2.08%higher than those without LIPCE.The flame combustion oscillation of kerosene tended to be stable in the recirculation zone of cavity with the thermal and chemical effects of laser induced plasma.The differences of flame temperature at a=1.0b and 0.6 were 503 K and 523 K with LIPCE,which were 20.07%and42.64%lower than those without LIPCE.The flame emissivity with methane assisted ignition was 80.46%lower than that without methane assisted ignition,due to the carbon-hydrogen ratio of kerosene was higher than that of methane.The spectral emissivities at 600 nm with LIPCE were 1.25%,22.2%,and 4.22%lower than those without LIPCE at a=1.0a(with methane assisted ignition),1.0b(without methane assisted ignition)and 0.6.The effect of concentration in the emissivity was removed by normalization to analyze the flame radiative properties in the RBCC combustor chamber.The maximum differences of flame normalized emissivity were 50.91%without LIPCE and 27.53%with LIPCE.The flame radiative properties were stabilized under the thermal and chemical effects of laser induced plasma at a=0.6.
基金supported by the National Natural Science Foundation of China(Nos.11925207 and 92252206)the Hunan Province Graduate Innovation Project,China(No.XJCX2023059)。
文摘To reduce the drag generated by the recirculation flow at the rocket base in a RocketBased Combined Cycle(RBCC)engine operating in the ramjet/scramjet mode,a novel annular rocket RBCC engine based on a central plug cone was proposed.The performance loss mechanism caused by the recirculation flow at the rocket base and the influence of the plug cone configuration on the thrust performance were studied.Results indicated that the recirculation flow at the rocket base extended through the entire combustor,which creates an extensive range of the"low-kineticenergy zone"at the center and leads to an engine thrust loss.The plug cone serving as a surface structure had a restrictive effect on the internal flow of the engine,making it smoothly transit at the position of the large separation zone.The model RBCC engine could achieve a maximum thrust augmentation of 37.6%with a long plug cone that was twice diameter of the inner isolator.However,a shorter plug cone that was half diameter of the inner isolator proved less effective at reducing the recirculation flow for a supersonic flow and induced an undesirable flow fraction that diminished the thrust performance.Furthermore,the effectiveness of the plug cone increased with the flight Mach number,indicating that it could further broaden the operating speed range of the scramjet mode.
基金supported by the National Natural Science Foundation of China (Grant No.52076094).
文摘Ejector mode is a unique and critical phase of wide-range rocket-based combined cycle(RBCC)engine.In this paper,a quasi-one-dimensional thermodynamic performance modeling method,with more detailed model treatments for the inlet-diffuser system,pri-mary/secondaryflow interaction,and pressure feedback matching,was developed for operating characteristics studies and multi-objective optimization analysis of the ejector mode of an actual RBCC engine.A series of three-dimensional simulations of separate inlet and fullflowpath was completed to validate the modeling study and provide further insight into the operating charac-teristics.The primary/secondary equilibrium pressure ratio functions a significant effect on ejector mode performance,a higher performance augmentation can be obtained by lower rocket pressure ratio,larger mixing section area ratio,smaller throttling throat and higher equivalence ratio,within an appropriate range.The positive performance augmentation can be realized at lowflight Mach conditions,the coordination and trade-off relationships between specific im-pulse,performance augmentation ratio and thrust-to-area ratio during ejector mode are present by the Pareto-front from MOP analysis.It is further verified by CFD simulation that,the operating back-pressure at the exit of inlet-diffuser system functions a decisive influence on the airbreathing characteristics,the pressure feedback and matching should be well-controlled for secondaryflowrate and performance augmentation.The thermodynamic modeling analysis re-sults are basically consistent with those of numerical simulation,to validate the rationality and effectiveness of the modeling method.
基金This work was supported by the National Natural Science Foundation of China(NSFC No.50776045,51076064)China Scholarship Council's International Students Scholarship(CSC No.2011YXS867)from the Minister of Education,China and NUAA.
文摘Combined-cycle pulse detonation engines are promising contenders for hypersonic propulsion systems.In the present study,design and propulsive performance analysis of combined-cycle pulse detonation turbofan engines(PDTEs)is presented.Analysis is done with respect to Mach number at two consecutive modes of operation:(1)Combined-cycle PDTE using a pulse detonation afterburner mode(PDA-mode)and(2)combined-cycle PDTE in pulse detonation ramjet engine mode(PDRE-mode).The performance of combined-cycle PDTEs is compared with baseline afterbuming turbofan and ramjet engines.The comparison of afterburning modes is done for Mach numbers from 0 to 3 at 15.24 km altitude conditions,while that of pulse detonation ramjet engine(PDRE)is done for Mach 1.5 to Mach 6 at 18.3 km altitude conditions.The analysis shows that the propulsive performance of a tubine engine can be greatly improved by replacing the conventional afterbumer with a pulse detonation afterburner(PDA).The PDRE also outperforms its ramjet counterpart at all flight conditions considered herein.The gains obtained are outstanding for both the combined-cycle PDTE modes compared to baseline turbofan and ramjet engines.
