摘要 :
Heat transfer characteristics near the walls of the Rocket Based Combined Cycle (RBCC) whole flow path engine at typical operation conditions (rocket-ejector mode at Ma1.5, ramjet mode at Ma4.0, scramjet mode at Ma6.0) were numeri...
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Heat transfer characteristics near the walls of the Rocket Based Combined Cycle (RBCC) whole flow path engine at typical operation conditions (rocket-ejector mode at Ma1.5, ramjet mode at Ma4.0, scramjet mode at Ma6.0) were numerically studied. The numerical model employed the SST k-ω turbulent model in the flow field and a three-step quasi-global chemical kinetics model for combustion simulation of C_(12)H_(23) which was selected as kerosene surrogate. It is found that the distribution surface heat flux in RBCC engine which serves a vital role in the process of TPS design has its unique characteristic at rocket-ejector mode, ramjet mode and scramjet mode, especially on the combustor wall, the reasons for the non-uniform thermal environment are that the mass flow rate of primary rocket and positions of combustion organization vary with operating modes. Among the three typical operation conditions, Ma6.0 has the highest surface heat flux, the averaged heat flux is 2MW/m2. The heat flux distribution on the side wall is lower than on the top and bottom wall because the influence area of the primary rocket is limited which mainly effect the non-uniform distribution on the top and bottom wall. Another notable characteristic is that the cavity's thermal environment is quite different with the combustor wall, and the maximum heat flux is often located at the trailing edge of the cavity which approximate to 5MW/m2, therefore, the TPS design of the cavity must consider this unique phenomenon. As concluding remarks, RBCC engine has a more complex surface thermal environment than ordinary scramjet engine, the TPS design must depend on a comprehensive and detailed analysis to obtain a high efficient of closed-loop cooling cycle.
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摘要 :
Heat transfer characteristics near the walls of the Rocket Based Combined Cycle (RBCC) whole flow path engine at typical operation conditions (rocket-ejector mode at Ma1.5, ramjet mode at Ma4.0, scramjet mode at Ma6.0) were numeri...
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Heat transfer characteristics near the walls of the Rocket Based Combined Cycle (RBCC) whole flow path engine at typical operation conditions (rocket-ejector mode at Ma1.5, ramjet mode at Ma4.0, scramjet mode at Ma6.0) were numerically studied. The numerical model employed the SST k-ω turbulent model in the flow field and a three-step quasi-global chemical kinetics model for combustion simulation of C_(12)H_(23) which was selected as kerosene surrogate. It is found that the distribution surface heat flux in RBCC engine which serves a vital role in the process of TPS design has its unique characteristic at rocket-ejector mode, ramjet mode and scramjet mode, especially on the combustor wall, the reasons for the non-uniform thermal environment are that the mass flow rate of primary rocket and positions of combustion organization vary with operating modes. Among the three typical operation conditions, Ma6.0 has the highest surface heat flux, the averaged heat flux is 2MW/m2. The heat flux distribution on the side wall is lower than on the top and bottom wall because the influence area of the primary rocket is limited which mainly effect the non-uniform distribution on the top and bottom wall. Another notable characteristic is that the cavity's thermal environment is quite different with the combustor wall, and the maximum heat flux is often located at the trailing edge of the cavity which approximate to 5MW/m2, therefore, the TPS design of the cavity must consider this unique phenomenon. As concluding remarks, RBCC engine has a more complex surface thermal environment than ordinary scramjet engine, the TPS design must depend on a comprehensive and detailed analysis to obtain a high efficient of closed-loop cooling cycle.
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摘要 :
Rocket-Based Combined Cycle (RBCC) engine, which is formed out of rockets and ramjet/scramjet by structurally integration and thermodynamic cycle combination, can be designed into various type and configuration to adapt to differe...
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Rocket-Based Combined Cycle (RBCC) engine, which is formed out of rockets and ramjet/scramjet by structurally integration and thermodynamic cycle combination, can be designed into various type and configuration to adapt to different mission requirements. According to both of acceleration and cruise performance requirements, a wide operating range RBCC engine scheme was presented in this paper, based on variable geometry inlet of contraction-ratio adjustment, fixed-geometry combustor of thermal-throat regulation and variable geometry nozzle of expansion-ratio adjustment for adaptation to wide Mach Number range operating requirements. Numerical simulations of the full flowpath reactive flow-field at typical flight conditions were carried out for engine performance studies. Furthermore, the major influence factor on engine performance and thrust regulating characteristics during wide Mach Number range were analyzed. The results show that: 1) superior specific impulse at high Mach Number cruise were achieved while quite large thrust generated for low Mach number acceleration; 2 ) large thrust regulating range and agile, convenient thrust regulating ability were found during the wide Mach Number operating envelope, the thrust regulating ratio (the maximum/the minimum) is obviously increased with flight Mach Number rising; 3 ) the basic performance and characteristics obtained from the present efforts have preliminarily validated the great advantages of acceleration ability and maneuverability of RBCC engine, the flow-path design optimization and combined cycle performance optimum control are needed according to the specific mission requirements.
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摘要 :
Rocket-Based Combined Cycle (RBCC) engine, which is formed out of rockets and ramjet/scramjet by structurally integration and thermodynamic cycle combination, can be designed into various type and configuration to adapt to differe...
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Rocket-Based Combined Cycle (RBCC) engine, which is formed out of rockets and ramjet/scramjet by structurally integration and thermodynamic cycle combination, can be designed into various type and configuration to adapt to different mission requirements. According to both of acceleration and cruise performance requirements, a wide operating range RBCC engine scheme was presented in this paper, based on variable geometry inlet of contraction-ratio adjustment, fixed-geometry combustor of thermal-throat regulation and variable geometry nozzle of expansion-ratio adjustment for adaptation to wide Mach Number range operating requirements. Numerical simulations of the full flowpath reactive flow-field at typical flight conditions were carried out for engine performance studies. Furthermore, the major influence factor on engine performance and thrust regulating characteristics during wide Mach Number range were analyzed. The results show that: 1) superior specific impulse at high Mach Number cruise were achieved while quite large thrust generated for low Mach number acceleration; 2 ) large thrust regulating range and agile, convenient thrust regulating ability were found during the wide Mach Number operating envelope, the thrust regulating ratio (the maximum/the minimum) is obviously increased with flight Mach Number rising; 3 ) the basic performance and characteristics obtained from the present efforts have preliminarily validated the great advantages of acceleration ability and maneuverability of RBCC engine, the flow-path design optimization and combined cycle performance optimum control are needed according to the specific mission requirements.
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摘要 :
As a highly integrated air breathing combined cycle engine, the TRRE engine combines the turbine, rocket and ramjet through the high structural integration and the organic combination of the thermal cycle and operation process. Th...
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As a highly integrated air breathing combined cycle engine, the TRRE engine combines the turbine, rocket and ramjet through the high structural integration and the organic combination of the thermal cycle and operation process. The present paper performed a preliminary analysis on the key technical difficulties and technical approaches of the TRRE. A brief introduction was conducted on the develop progress of the principle prototype from the Beijing Power Machinery Research Institute. The performance data from the flow channel design of the principle prototype and numerical simulations were utilized to assess the propulsion performance of the TRRE engine, which primitively validated its comprehensive advantages on the acceleration, cruise, mobility and other aspects. The results show that the TRRE engine can reconcile the demands of high thrust at lower Mach numbers and high specific impulse at a Mach number of 6.0 (pure scramjet). In a wide Mach number operating range, the TRRE engine provides the relatively wide, flexible and convenient adjustment ability in the thrust. It can satisfy the smooth relay between the high and low speed channels, illustrating the extremely advantageous adaptability in the acceleration, maneuver flight and high Mach number cruise.
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摘要 :
As a highly integrated air breathing combined cycle engine, the TRRE engine combines the turbine, rocket and ramjet through the high structural integration and the organic combination of the thermal cycle and operation process. Th...
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As a highly integrated air breathing combined cycle engine, the TRRE engine combines the turbine, rocket and ramjet through the high structural integration and the organic combination of the thermal cycle and operation process. The present paper performed a preliminary analysis on the key technical difficulties and technical approaches of the TRRE. A brief introduction was conducted on the develop progress of the principle prototype from the Beijing Power Machinery Research Institute. The performance data from the flow channel design of the principle prototype and numerical simulations were utilized to assess the propulsion performance of the TRRE engine, which primitively validated its comprehensive advantages on the acceleration, cruise, mobility and other aspects. The results show that the TRRE engine can reconcile the demands of high thrust at lower Mach numbers and high specific impulse at a Mach number of 6.0 (pure scramjet). In a wide Mach number operating range, the TRRE engine provides the relatively wide, flexible and convenient adjustment ability in the thrust. It can satisfy the smooth relay between the high and low speed channels, illustrating the extremely advantageous adaptability in the acceleration, maneuver flight and high Mach number cruise.
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摘要 :
Three-dimensional two-phase numerical method and ground direct-connect test system were used to investigate the influence of primary rocket mass flow rate with different inflow velocity on RBCC performance. Numerical results indic...
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Three-dimensional two-phase numerical method and ground direct-connect test system were used to investigate the influence of primary rocket mass flow rate with different inflow velocity on RBCC performance. Numerical results indicated that: under different inflow condition, the increment of primary flow had different contribution to net thrust and specific impulse; under low mach inflow condition, the increase of primary flow had positive effect on secondary combustion and greatly boosted thrust; under high mach condition, increasing primary flow would chock flow path, and weakened the influence of primary mass flow rate changing on RBCC thrust. Experimental results validated the rule obtained by numerical investigation between primary flow and RBCC performance. Especially under high mach inflow condition, with the Mach number increasing, the contribution of increment of primary flow to RBCC performance was reduced. Both the numerical and experimental investigation results could show that whether in low mach or high mach inflow condition, an optimized primary mass flow rate existed, which was benefit for improving RBCC performance.
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摘要 :
Ramjets exhibit undesirable combustion instabilities under certain operating conditions. In this paper, large-eddy simuIation(LES) and premixed combustion model have been efficiently implemented to study combustion instabilities i...
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Ramjets exhibit undesirable combustion instabilities under certain operating conditions. In this paper, large-eddy simuIation(LES) and premixed combustion model have been efficiently implemented to study combustion instabilities in ramjet combustors. The main objective is to predict dominant frequencies and various flow features of the combustors. Combustion instabilities in the ramjet dump combustor have been numerically simulated. Low-frequency, large-amplitude instabilities are observed, and instability frequencies and flame spread during various operating conditions are in good agreement with experimental observations. The results show large vortex structures dominate the flame propagation and vortex convection coupled unsteady heat release excites pressure oscillation in ramjet combustor.
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摘要 :
Accurate information on heat transfer data of combustion products in the solid rocket motor chamber is a crucial prerequisite for the engine thermal protection. A measurement technique was well developed to acquire steady-state he...
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Accurate information on heat transfer data of combustion products in the solid rocket motor chamber is a crucial prerequisite for the engine thermal protection. A measurement technique was well developed to acquire steady-state heat flux data of two-phase flow and was used successfully in the hostile environment. Experimental heat flux measurement has been obtained with an innovative designed instrument by simulating the flow field of complex charging configuration. The total heat flux of combustion products in the chamber was brought away by the coolant and calculated by its enthalpy rise in this device. The data could be used to analyze the heat transfer phenomena in SRMs and provide boundary condition for establishing insulation erosion model.
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摘要 :
Accurate information on heat transfer data of combustion products in the solid rocket motor chamber is a crucial prerequisite for the engine thermal protection. A measurement technique was well developed to acquire steady-state he...
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Accurate information on heat transfer data of combustion products in the solid rocket motor chamber is a crucial prerequisite for the engine thermal protection. A measurement technique was well developed to acquire steady-state heat flux data of two-phase flow and was used successfully in the hostile environment. Experimental heat flux measurement has been obtained with an innovative designed instrument by simulating the flow field of complex charging configuration. The total heat flux of combustion products in the chamber was brought away by the coolant and calculated by its enthalpy rise in this device. The data could be used to analyze the heat transfer phenomena in SRMs and provide boundary condition for establishing insulation erosion model.
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