[특허]Isentropic compression inlet for supersonic aircraft

Gulfstream Aerospace Corporation

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Isentropic compression inlet for supersonic aircraft

국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F02C-007/04
미국특허분류(USC)	060/767; 060/768; 060/204; 137/015.1; 244/053B
출원번호	US-0338010 (2011-12-27)
등록번호	US-8286434 (2012-10-16)
발명자 / 주소	Henne, Preston A. Conners, Timothy R. Howe, Donald C.
출원인 / 주소	Gulfstream Aerospace Corporation
대리인 / 주소	Ingrassia Fisher & Lorenz, P.C.
인용정보	피인용 횟수 : 3 인용 특허 : 82

초록

A supersonic inlet employs relaxed isentropic compression to improve net propulsive force by shaping the compression surface of the inlet. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce cowl lip surface angles, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining performance.

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1. A supersonic propulsion system, said propulsion system designed for flight at a specific and pre-determined Mach number, comprising: an engine comprising an air intake and an exhaust system;a subsonic diffuser section coupled to the air intake of the engine, and configured to diffuse a flow and to admit the diffused flow to said air intake of said engine at a predetermined subsonic condition suitable for the engine; anda supersonic compression section coupled to the subsonic diffuser section by a throat, the supersonic compression section comprising a compression ramp and cowl;said cowl having an upstream lip;said compression ramp having an upstream straight compression ramp having a leading edge or an apex, connected downstream with a concave surface relative to the flow, said concave surface connected downstream with a straight surface;said leading edge or apex having an angle, and said cowl lip is positioned such that an inclined shock wave generated at said leading edge or apex intercepts said cowl lip;said cowl lip operable to produce a terminal shock wave extending to the compression surface;said concave surface having a radius of concavity operable to produce successive shocklets;said radius of concavity being larger than a radius that would be operable to cause said shocklets to focus on said cowl lip;said concavity operable to generate each of a plurality of said shocklets such that, at said specific and pre-determined flight Mach number, each shocklet of the plurality of said shocklets intercepts said terminal shock wave at a different location between said cowl lip and said compression surface. 2. The supersonic propulsion system of claim 1, wherein, during operation of the engine at the specific and pre-determined Mach number, none of the successive shocklets focus on a point substantially adjacent to the cowl lip. 3. The supersonic propulsion system of claim 1, wherein, the compression ramp is further configured to cause the terminal shock to have a bowed region having a configuration such that as the bowed region approaches a point substantially adjacent to the cowl lip, a tangent of the bowed region approaches a direction orthogonal to a supersonic flow at a free-stream condition. 4. The supersonic propulsion system of claim 3, wherein the compression ramp is further configured to cause a variation in a Mach number along a length of the terminal shock, and wherein a first Mach number adjacent to the compression surface is substantially less than a second Mach number adjacent to the cowl lip. 5. The supersonic propulsion system of claim 3, wherein the compression ramp is further configured to cause a Mach number to vary along a length of the terminal shock such that a first gradient of a first Mach number across the bowed region of the terminal shock is greater than a second gradient of a second Mach number along the terminal shock from the compression ramp to the bowed region. 6. The supersonic propulsion system of claim 2, wherein the cowl lip is substantially aligned with a flow angle adjacent to the cow lip. 7. A supersonic aircraft comprising: an airframe configured for supersonic flight;at least one engine mounted to the airframe and comprising an air intake and an exhaust system;said air intake designed for flight at a specific and pre-determined Mach number, comprising:a subsonic diffuser section coupled to the air intake of the engine, and configured to diffuse a flow and to admit the diffused flow to said air intake of said engine at a predetermined subsonic condition suitable for the engine; anda supersonic compression section coupled to the subsonic diffuser section by a throat, the supersonic compression section comprising a compression ramp and cowl;said cowl having an upstream lip;said compression ramp having an upstream straight compression ramp having a leading edge or an apex, connected downstream with a concave surface relative to the flow, said concave surface connected downstream with a straight surface;said leading edge or apex having an angle, and said cowl lip is positioned such that an inclined shock wave generated at said leading edge intercepts said cowl lip;said cowl lip operable to produce a terminal shock wave extending to the compression surface;said concave surface having a radius of concavity operable to produce successive shocklets;said radius of concavity being larger than a radius that would be operable to cause said shocklets to focus on said cowl lip;said concavity operable to generate each of a plurality of said shocklets such that, at said specific and pre-determined flight Mach number, each shocklet of the plurality of said shocklets intercepts said terminal shock wave at a different location between said cowl lip and said compression surface. 8. The supersonic aircraft of claim 7, wherein, during operation of the at least one engine at the specific and pre-determined Mach number, none of the successive shocklets focus on a point substantially adjacent to the cowl lip. 9. The supersonic propulsion system of claim 7, wherein, the compression ramp is further configured to cause the terminal shock to have a bowed region having a configuration such that as the bowed region approaches a point substantially adjacent to the cowl lip, a tangent of the bowed region approaches a direction orthogonal to a supersonic flow at a free-stream condition. 10. The supersonic aircraft of claim 9, wherein the compression ramp is further configured to cause a variation in a Mach number along a length of the terminal shock, and wherein a first Mach number adjacent to the compression surface is substantially less than a second Mach number adjacent to the cowl lip. 11. The supersonic aircraft of claim 9, wherein the compression ramp is further configured to cause a Mach number to vary along a length of the terminal shock such that a first gradient of a first Mach number across the bowed region of the terminal shock is greater than a second gradient of a second Mach number along the terminal shock from the compression ramp to the bowed region. 12. The supersonic aircraft of claim 7, wherein the cowl lip is substantially aligned with a flow angle adjacent to the cowl lip.

이 특허에 인용된 특허 (82)

이 특허를 인용한 특허 피인용횟수: 3

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Isentropic compression inlet for supersonic aircraft

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Isentropic compression inlet for supersonic aircraft

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이 특허에 인용된 특허 (82)

이 특허를 인용한 특허 피인용횟수: 3

IPC 상위 출원인

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IPC 상세정보(version : 201201)

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