초록

A supersonic combustion apparatus and method of using the same including a side wall cavity having an enhanced mixing system with ground-based oxygen injection for hypersonic material and engine testing.

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1. A supersonic combustion heater apparatus capable of withstanding high enthalpy flow for operating at high Mach numbers comprising:a means for providing a high-pressure flow;a first nozzle having a throat to withstand said high pressure flow, whereby a boundary layer flow is created downstream of

1. A supersonic combustion heater apparatus capable of withstanding high enthalpy flow for operating at high Mach numbers comprising:a means for providing a high-pressure flow;a first nozzle having a throat to withstand said high pressure flow, whereby a boundary layer flow is created downstream of said first nozzle;a supersonic combustion region is located adjacent to said first nozzle, said region including a fuel injection means for ignition and an oxygen injection means for maintaining flame stabilization; andan expansion zone dimensioned and configured for withstanding high enthalpy and a supersonic combustion flow, said expansion zone is adjacent to said supersonic combustion region, said expansion zone including a second expansion nozzle, and a divergent area dimensioned and configured to withstand high enthalpy flow and a supersonic combustion flow, said divergent area is adjacent to said supersonic combustion region whereby increasing high Mach speeds are achieved as said supersonic combustion flow reaches downstream of said divergent area.2. A supersonic combustion heater apparatus capable of withstanding high enthalpy flow for operating at high Mach numbers comprising:an upstream air heater to provide heated high-pressure flow;a first nozzle having a throat to withstand said heated high-pressure flow, whereby a boundary layer flow is created downstream of said first nozzle;a supersonic combustion region including at least one acoustic cavity having a downstream lip to cause shedding of periodic coherent vortices downstream, a fuel injection means for ignition and rapid mixing with said vortices, and an oxygen injection means for maintaining flame stabilization; andan expansion zone dimensioned and configured for withstanding high enthalpy and a supersonic combustion flow, said expansion zone including a second expansion nozzle, and a divergent area dimensioned and configured to withstand high enthalpy flow and a supersonic combustion flow, wherein increasing high Mach speeds are achieved as said supersonic combustion flow reaches downstream of said divergent area.3. The supersonic combustion apparatus according to claim 1, wherein said air heater is a vitiator which supplies oxygen into said heated high-pressure flow.4. The supersonic combustion apparatus according to claim 1, wherein said first nozzle is constructed to withstand a partial expansion beyond Mach 1.0.5. The supersonic combustion apparatus according to claim 1, wherein said cavity having a side wall cavity of a length to depth ratio of about four to one.6. The supersonic combustion apparatus according to claim 1, wherein said cavity is dimensioned and configured for desired acoustic resonance to aid in driving coherent vorticity within said boundary layer flow.7. The supersonic combustion apparatus according to claim 1, wherein said fuel injection means supplies a combustible fuel into the wake of said cavity.8. The supersonic combustion apparatus according to claim 6, wherein said combustible fuel is selected from the group consisting of hydrogen and hydrocarbons or the like, or any combination thereof.9. The supersonic combustion apparatus according to claim 1, wherein said combustible fuel is hydrogen.10. The supersonic combustion apparatus according to claim 1, wherein said oxygen injection means is introduced adjacent of said fuel injection means and said cavity for maintaining supersonic flow and combustion.11. The supersonic combustion apparatus according to claim 1, wherein said second nozzle is constructed to withstand a partial expansion of Mach 3.0 or greater.12. The supersonic combustion apparatus according to claim 1, wherein said increasing high Mach speeds are achieved as said supersonic combustion flow reaches downstream of said divergent area is between about Mach 1.0 to about Mach 6.0.13. The supersonic combustion apparatus according to claim 1, wherein said high Mach speeds are from approximately Mach 3 to about Mach 6.5.14. The supersonic combustion apparatus according to claim 1, wherein said high pressure flow is between approximately 100 psi to about 2000 psi.15. The supersonic combustion apparatus according to claim 1, wherein said high enthalpy is from approximately 500 Kelvin to about 2400 Kelvin at about Mach 3.0 to about 6.5.16. A supersonic combustion apparatus and heater capable of withstanding high enthalpy flow for operating at high Mach numbers comprising:air heater to provide heated high-pressure flow;a subsonic combustion region including:a first combustor chamber for subsonic combustion;a first moderate temperature first nozzle having a throat to withstand subsonic combustion flow, said heated high-pressure flow is expanded through said first nozzle creating a boundary layer flow downstream of said nozzle; anda supersonic combustion region including:at least one side wall cavity having a length to depth ratio dimensioned and configured for desired acoustic resonance, said cavity having a downstream lip, whereby said boundary layer flow flaps over said cavity to impinge on said downstream lip, thereby causing period shedding of vortices downstream of said boundary layer flow;at least one fuel injection means for supplying a combustible fuel for ignition and rapid mixing with said vortices to enhance supersonic combustion;at least one oxygen injection means adjacent of said cavity and said vortices for maintaining flame stabilization; said fuel injection means and said oxygen injection means are for maintaining supersonic flow and combustion;an expansion zone being downstream of said cavity, said expansion zone having an expansion angle dimensioned and configured for withstanding high enthalpy and a supersonic combustion flow, said expansion zone sustaining a significant portion of said high enthalpy;a second expansion nozzle having a throat downstream of said expansion zone, said second nozzle to withstand the remaining portion of the total enthalpy and supersonic combustion flow; anda divergent area having an expansion angle dimensioned and configured to withstand high enthalpy flow and a supersonic combustion flow, said divergent area is downstream of said second nozzle, wherein increasing high Mach speeds are achieved while the supersonic combustions flow reaches downstream of said divergent area.17. The supersonic combustion apparatus according to claim 16, wherein said air heater is an axissymmetric burner.18. The supersonic combustion apparatus according to claim 16, wherein said air heater is a vitiator which supplies oxygen into said heated high pressure flow.19. The supersonic combustion apparatus according to claim 16, wherein said first nozzle is constructed to withstand a partial expansion beyond Mach 1.0.20. The supersonic combustion apparatus according to claim 16, wherein said first nozzle is constructed to withstand a partial expansion to accelerate said flow to supersonic velocities.21. The supersonic combustion apparatus according to claim 16, wherein said cavity is a side wall cavity having a length to depth ratio of about four to one.22. The supersonic combustion apparatus according to claim 16, wherein said cavity is dimensioned and configured for desired acoustic resonance to aid in driving coherent vorticity within said boundary layer flow.23. The supersonic combustion apparatus according to claim 16, wherein said downstream lip of said cavity causes shedding of periodic coherent lateral vortices downstream.24. The supersonic combustion apparatus according to claim 16, wherein said fuel injection means supplies a combustible fuel into the wake of said cavity.25. The supersonic combustion apparatus according to claim 24, wherein said combustible fuel is selected from the group consisting of hydrogen and hydrocarbons, or the like.26. The supersonic combustion apparatus according to claim 16, wherein said combustible fuel is hydrogen.27. The supersonic combustion apparatus according to claim 16, wherein said oxygen injection means is introduced downstream to said fuel injection means.28. The supersonic combustion apparatus according to claim 16, wherein said second nozzle is constructed to withstand a partial expansion of Mach 3.0 or greater.29. The supersonic combustion apparatus according to claim 16, wherein said increasing high Mach speeds are achieved as said supersonic combustion flow reaches downstream of said divergent area is between about Mach 1.0 to about Mach 8.0.30. The supersonic combustion apparatus according to claim 16, wherein said increasing high Mach speeds are achieved as said supersonic combustion flow reaches downstream of said divergent area is between about Mach 1.0 to about Mach 6.0.31. A method of using supersonic combustion to create a high enthalpy flow source for application in scramjets comprising the steps of:providing a high-pressure flow which is expanded through a first nozzle creating a duct flow having a boundary layer flow;injecting three fluid streams for rapid mixing including the duct flow, a fuel, and auxiliary oxygen; andpartitioning a significant portion of the total enthalpy to an expansion zone and directing the remaining enthalpy via supersonic combustion downstream of a second expansion nozzle.32. A method of using supersonic combustion to create a high enthalpy flow source for application in scramjets comprising the steps of:providing advanced active combustion control by controlling input enthalpy with a preheater;providing a heated high-pressure flow which is expanded through a first nozzle creating a duct flow having a boundary layer flow;generating coherent vortices using a resonant acoustic side wall cavity having a downstream lip;flapping of said boundary layer flow over said side wall cavity with periodical impinging on its downstream lip causes shedding of periodic coherent vortices downstream to enhance supersonic mixing rates and shorten mixing times while increasing combustion efficiency;injecting fuel downstream of the vortex shedding point;entraining of fuel into the supersonic vortex and rapid mixing with the duct flow;injecting oxygen for enhancing kinetics, increasing enthalpy, and enhancing flame stability; andpartitioning a significant portion of the total enthalpy to an expansion zone and directing the remaining enthalpy via supersonic combustion downstream of a second expansion nozzle, wherein increasing high Mach speeds are achieved while the supersonic combustions flow reaches downstream of a divergent area.33. The method according to claim 32, wherein the step of providing said heated high-pressure flow utilizes a vitiator.34. The method according to claim 32, wherein the step of providing said heated high-pressure flow utilizes an air heater.35. The method according to claim 32, wherein the step of injecting fuel downstream of said vortex shedding point is carried out with at least one combustible propellant.36. The method according to claim 32, wherein the combustible propellant is selected from the group consisting of hydrogen and hydrocarbons, or the like.37. The method according to claim 32, wherein the combustible propellant is hydrogen.38. The method according to claim 32, further comprises the step of preheating the fuel.39. The method according to claim 32, further comprising the step of optimizing local fuel to air/oxidizer ratios and temperature to insure ignition.40. The method according to claim 32, wherein said high Mach speeds are from approximately Mach 3 to about Mach 6.5.41. The method according to claim 32, wherein said high pressure flow is between approximately 100 psi to about 2000 psi.42. The method according to claim 32, wherein said high enthalpy is from approximately 500 Kelvin to about 2400 Kelvin at about Mach 3.0 to about 6.5.