Hypersonic airbreathing propulsion is one of the top techniques for future aerospace flight, but there are still no practical engines after seventy years’ development. Two critical issues are identified to be the bar...Hypersonic airbreathing propulsion is one of the top techniques for future aerospace flight, but there are still no practical engines after seventy years’ development. Two critical issues are identified to be the barriers for the ramjet-based engine that has been taken as the most potential concept of the hypersonic propulsion for decades. One issue is the upstream-traveling shock wave that develops from spontaneous waves resulting from continuous heat releases in combustors and can induce unsteady combustion that may lead to engine surging during scramjet engine operation.The other is the scramjet combustion mode that cannot satisfy thrust needs of hypersonic vehicles since its thermos-efficiency decreases as the flight Mach number increases. The two criteria are proposed for the ramjet-based hypersonic propulsion to identify combustion modes and avoid thermal choking. A standing oblique detonation ramjet(Sodramjet) engine concept is proposed based on the criteria by replacing diffusive combustion with an oblique detonation that is a unique pressure-gain phenomenon in nature. The Sodramjet engine model is developed with several flow control techniques, and tested successfully with the hypersonic flight-duplicated shock tunnel.The experimental data show that the Sodramjet engine model works steadily, and an oblique detonation can be made stationary in the engine combustor and is controllable. This research demonstrates the Sodramjet engine is a promising concept and can be operated stably with high thermal efficiency at hypersonic flow conditions.展开更多
The hysteresis during the throat regulation process of a supersonic variable inlet is unconducive to restart.Hence,detailed experimental studies of such a hysteresis and its control are necessary.A throat variable sup...The hysteresis during the throat regulation process of a supersonic variable inlet is unconducive to restart.Hence,detailed experimental studies of such a hysteresis and its control are necessary.A throat variable supersonic inlet was designed at a shock-on-lip Mach number of 4.0 and an Internal Contraction Ratio(ICR)ranging over 1.21–2.94.Meanwhile,a distributed bleed system was proposed to suppress the hysteresis.The wind tunnel tests were conducted at Mach number 2.9.The throat regulation processes were recorded using a high-speed schlieren and dynamic pressure acquisition system.The results indicate that the unstart and restart ICRs during the uncontrolled inlet’s throat regulation process were 1.95 and 1.48,respectively,demonstrating an unstart-restart hysteresis.Four typical flowfields were summarized during the uncontrolled inlet’s restart process.The proposed bleed control increased the unstart and restart ICRs to 2.06 and 1.75,respectively,and the inlet realized the designed state as the ICR was further decreased to 1.67.The controlled inlet’s hysteresis loop was decreased compared to the uncontrolled inlet.Finally,the mechanism of the hysteresis,dominated by the entrance separation-induced wave system,was clarified.The mechanisms of the bleed control to broaden the unstart and restart boundaries and suppress the hysteresis were elucidated.展开更多
To simulate the actual flowfield at the exit of the supersonic/hypersonic inlet, a wind tunnel is designed to study the flow in the scramjet isolator under the asymmetric incoming flow. And compression fields in the i...To simulate the actual flowfield at the exit of the supersonic/hypersonic inlet, a wind tunnel is designed to study the flow in the scramjet isolator under the asymmetric incoming flow. And compression fields in the isolator are investigated using wall static and pitot pressure measurements. Three incoming Mach numbers are considered as 1.5, 1.8 and 2. Results show that the increase of the asymmetry of the flow at the isolator entrance leads to the increase of the shock train length in the isolator for a given pressure ratio. Based on the analysis of the flow asymmetry effect at the isolator entrance on the shock train length, a modified correlation is proposed to calculate the length of the shock train. Predicted results of the proposed correlation are in good agreement with the experimental data.展开更多
Despite over fifty years of research on shock wave boundary layer effects and interactions,many related technical issues continue to be controversial and debated.The present survey provides an overview of the present ...Despite over fifty years of research on shock wave boundary layer effects and interactions,many related technical issues continue to be controversial and debated.The present survey provides an overview of the present state of knowledge on such effects and interactions,including discussions of:(i)general features of shock wave interactions,(ii)test section configurations for investigation of shock wave boundary layer interactions,(iii)origins and sources of unsteadiness associated with the interaction region,(iv)interactions which included thermal transport and convective heat transfer,and(v)shock wave interaction control investigations.Of particular interest are origins and sources of low-frequency,large-scale shock wave unsteadiness,flow physics of shock wave boundary layer interactions,and overall structure of different types of interactions.Information is also provided in regard to shock wave investigations,where heat transfer and thermal transport were important.Also considered are investigations of shock wave interaction control strategies,which overall,indicate that no single shock wave control strategy is available,which may be successfully applied to different shock wave arrangements,over a wide range of Mach numbers.Overall,the survey highlights the need for additional understanding of fundamental transport mechanisms,as related to shock waves,which are applicable to turbomachinery,aerospace,and aeronautical academic disciplines.展开更多
基金supported by the National Natural Science Foundation of China (No. 11532014)。
文摘Hypersonic airbreathing propulsion is one of the top techniques for future aerospace flight, but there are still no practical engines after seventy years’ development. Two critical issues are identified to be the barriers for the ramjet-based engine that has been taken as the most potential concept of the hypersonic propulsion for decades. One issue is the upstream-traveling shock wave that develops from spontaneous waves resulting from continuous heat releases in combustors and can induce unsteady combustion that may lead to engine surging during scramjet engine operation.The other is the scramjet combustion mode that cannot satisfy thrust needs of hypersonic vehicles since its thermos-efficiency decreases as the flight Mach number increases. The two criteria are proposed for the ramjet-based hypersonic propulsion to identify combustion modes and avoid thermal choking. A standing oblique detonation ramjet(Sodramjet) engine concept is proposed based on the criteria by replacing diffusive combustion with an oblique detonation that is a unique pressure-gain phenomenon in nature. The Sodramjet engine model is developed with several flow control techniques, and tested successfully with the hypersonic flight-duplicated shock tunnel.The experimental data show that the Sodramjet engine model works steadily, and an oblique detonation can be made stationary in the engine combustor and is controllable. This research demonstrates the Sodramjet engine is a promising concept and can be operated stably with high thermal efficiency at hypersonic flow conditions.
基金This work was co-funded by the National Natural Science Foundation of China(Nos.U20A2070,12025202,and 12172175)the National Science and Technology Major Project,China(No.J2019-II-0014-0035).
文摘The hysteresis during the throat regulation process of a supersonic variable inlet is unconducive to restart.Hence,detailed experimental studies of such a hysteresis and its control are necessary.A throat variable supersonic inlet was designed at a shock-on-lip Mach number of 4.0 and an Internal Contraction Ratio(ICR)ranging over 1.21–2.94.Meanwhile,a distributed bleed system was proposed to suppress the hysteresis.The wind tunnel tests were conducted at Mach number 2.9.The throat regulation processes were recorded using a high-speed schlieren and dynamic pressure acquisition system.The results indicate that the unstart and restart ICRs during the uncontrolled inlet’s throat regulation process were 1.95 and 1.48,respectively,demonstrating an unstart-restart hysteresis.Four typical flowfields were summarized during the uncontrolled inlet’s restart process.The proposed bleed control increased the unstart and restart ICRs to 2.06 and 1.75,respectively,and the inlet realized the designed state as the ICR was further decreased to 1.67.The controlled inlet’s hysteresis loop was decreased compared to the uncontrolled inlet.Finally,the mechanism of the hysteresis,dominated by the entrance separation-induced wave system,was clarified.The mechanisms of the bleed control to broaden the unstart and restart boundaries and suppress the hysteresis were elucidated.
文摘To simulate the actual flowfield at the exit of the supersonic/hypersonic inlet, a wind tunnel is designed to study the flow in the scramjet isolator under the asymmetric incoming flow. And compression fields in the isolator are investigated using wall static and pitot pressure measurements. Three incoming Mach numbers are considered as 1.5, 1.8 and 2. Results show that the increase of the asymmetry of the flow at the isolator entrance leads to the increase of the shock train length in the isolator for a given pressure ratio. Based on the analysis of the flow asymmetry effect at the isolator entrance on the shock train length, a modified correlation is proposed to calculate the length of the shock train. Predicted results of the proposed correlation are in good agreement with the experimental data.
文摘Despite over fifty years of research on shock wave boundary layer effects and interactions,many related technical issues continue to be controversial and debated.The present survey provides an overview of the present state of knowledge on such effects and interactions,including discussions of:(i)general features of shock wave interactions,(ii)test section configurations for investigation of shock wave boundary layer interactions,(iii)origins and sources of unsteadiness associated with the interaction region,(iv)interactions which included thermal transport and convective heat transfer,and(v)shock wave interaction control investigations.Of particular interest are origins and sources of low-frequency,large-scale shock wave unsteadiness,flow physics of shock wave boundary layer interactions,and overall structure of different types of interactions.Information is also provided in regard to shock wave investigations,where heat transfer and thermal transport were important.Also considered are investigations of shock wave interaction control strategies,which overall,indicate that no single shock wave control strategy is available,which may be successfully applied to different shock wave arrangements,over a wide range of Mach numbers.Overall,the survey highlights the need for additional understanding of fundamental transport mechanisms,as related to shock waves,which are applicable to turbomachinery,aerospace,and aeronautical academic disciplines.
