• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    Device for preventing "hum" during supersonic air intake to reaction engine, especially with solid fuel driven frame engine without control of the amount of air taken in, amount of fuel and outflow and with a non-fully rotationally symmetrical cross-section of the air intake, by freeing the air intake plane (E) on the central body (2) by means of salmon-tail slits (7) or a perforated channel (9) covering outlet channels (10) in the central body (Fig. 1).
    公开号:SE7900140A1
    申请号:SE7900140
    申请日:1979-01-08
    公开日:2012-02-24
    发明作者:G Jungclaus；E-O Krohn
    申请人:
    IPC主号:
    专利说明:

    15 20 25 35 40 o ut no oo IIO O IU II the sound diffuser is further reduced with a simultaneous increase in pressure.
    No matter how striking the phenomenon of pressure turnover at minimum volume through compression shocks and the utilization of this phenomenon in supersonic flow, it is just as difficult today to stabilize the final compression shock in certain operating areas of the supersonic diffuser for maintaining the desired motor effects.
    In the case of supersonic diffusers, a distinction can be made between two different operating states, namely the "supercritical" operating state, in which the final perpendicular compression stroke lies in the interior of the diffuser, and the "subcritical" operating state in which the final compression stroke travels in front of the intake duct. The most favorable operating point, namely the maximum of the product pressure times air flow where in an operating diagram the first-mentioned operating quantity is plotted on the ordinate and the second operating quantity on the abscissa, lies precisely at the transition point between the "subcritical" and the "super-critical" operating state.
    As a special disturbance of the air inflow in supersonic diffusers, a phenomenon which can be called "hum" or shorter "hum" has shown itself. This disturbance occurs in the "subcritical" operating state. In this case, the perpendicular compression shock, which has already moved to the position in front of the front impact edge of the intake channel, finds no stable position and runs back and forth on the central body. This not only leads to strong pressure fluctuations in the flow and thus to a considerable reduction of the average pressure and air flow (extinguishing of the combustion chamber in jet fuel-powered jet engines; extinguishing flames in solid-fuel-driven frame engines), J but can also cause mechanical destruction of air intake and fuselage.
    On these grounds, it is necessary to unconditionally prevent the most unfavorable "hum".
    In conventional supersonic air intakes, this is achieved by not placing the operating point in the above-mentioned most favorable operating point, but displacing it slightly into the supercritical area. Thereby and with additional measures and devices, such as a variable geometry of the diffuser and control of the fuel dosage to the combustion chamber of the associated engine and by propulsion nozzle adjustment, 10 FJ UI (_11 (DO]) has been achieved. 0 o 0 o 0 00 noocu 0 o Uno o 0 o that the deviation from the operating point, in any 0000 0000 šåan "bàoëådé Išå" special flight maneuvers, never enters the subcritical operating state, which is a prerequisite for generating the "hum" (pulsation forward and back of the perpendicular compression stroke on the central body). However, the above-mentioned measures and devices cause a loss of power due to the deviation from the most favorable operating point and result in more expensive construction and weight gain.
    The object of the invention is to provide, by means of a relatively simple device, a safe elimination of the extremely unfavorable "hum" in the case of supersonic diffusers.
    The invention achieves this by arranging air outflow openings in the central body in the direction of flow in front of the inlet plane of the air intake pipe.
    In an embodiment of the invention, at an supersonic intake, in which between the inside of the air intake resp. The central body and the adjacent outer contour (s) of the fuselage and an air gap permitting its boundary layer flow are arranged which are bridged by a boundary layer plow, several air outlets arranged in the central body in the flow direction one after the other and extending towards the boundary layer gap. In this case, the boundary layer plow is designed so that the pressure in the boundary layer gap is lower than that on the surface of the central body, at least lower than the pressure behind the extending perpendicular compression shock. ' In this case, these air outflow slots in cross section can be salmontail-shaped, whereby uniformity of the outflow radially from the outside and inwards is ensured.
    Another preferred embodiment of the invention consists of one or more air outlet channels arranged in the central body in the flow direction one after the other and extending in the circumferential direction, which are closed with a many rows of inflow holes having a lid, a heel plate, radially outwards and opening towards the boundary layer gap.
    The proposed measures have proved to be effective in numerous attempts to prevent “hum.” Furthermore, the device according to the invention is simple to build, inexpensive and also weight-saving.
    The drawings show design examples according to the invention, in which Fig. 1 and Fig. 2 show an acoustic intake with air flow devices of different natures.
    The semi-rotating body-shaped supersonic air intake arranged on a fuselage 1 consists essentially of a central body 2, air intake pipe 3 and a boundary layer plow 4, which bridges the free air gap 5 between the inwardly facing flat surface 6 of the supersonic intake and the adjacent surface area 1a of it is unsuitable for air intake, ie. heated and less energy-rich boundary layer flow. The boundary layer plow 4 is then designed so that the boundary flow has a lower pressure than the flow on the surface of the central body 2.
    The airflow arriving at supersonic velocity is denoted by L. It is slowed down in a known manner in the supersonic diffuser and its kinetic energy is converted into several, here for example three consecutive oblique compression shocks a, b and c and finally in a perpendicular compression shock, to pressure energy. After the last compression shock, there is an ultrasonic speed.
    In the operating area of the engine in question resp. at the supersonic intake, the compression shocks a, b, c and d stably assume the plotted directions.
    When the fuselage, e.g. by special flight maneuvers, reduces its speed or the flow conditions in any other way deteriorate, the pressure and thus the air flow in the diffuser decreases.
    The engine uses less air during this flight condition, and the final compression shock d then travels forward in front of the impact edge Sa resp. in front of the intake plane E of the air intake pipe 3 as indicated by d '. The air intake then operates in the subcritical state, whereby it is important to keep the perpendicular impact d 'stable on the central body 2.
    According to Fig. 1, in the subcritical state, a part L 'of the air L flowing over the central body 2 flows via the S in cross-section salmontail-shaped air outflow slots 7 to the boundary layer flow L ". The same according to Fig. 2 via many small holes 8 in a lid 9 (perforated plate ) which closes two air outlet ducts 10 radially outwards.
    Through the outflow of the boundary layer in the critical area of the intake diffuser resp. at the central body 2 before the air intake plane E, the outwardly displaced compression shock d 'is stabilized and thus hum is prevented.
    权利要求:
    Claims (4)
    [1]
    1. intake to air-fed reaction fuel-driven frame-engine out air, fuel quantity and scatter only fuselages with a non-intake cross-section, in particular the characteristic of the intake plane (E) of air-entrained (2) air outlets
    [2]
    2. the outer contour of the air gap transmitting air gap is a layer plow, the body (2) in the flow direction and in the circumferential direction extending (5) opens air flow plow (4) is designed so that (5) is lower than that at the surface lower than the pressure behind the release shock (d '), which then lies
    [3]
    Device according to claim of air outflow slot shaped.
    [4]
    4. Device according to a crane of one or more, in one direction successive air outlet ducts or several rows of inflow 122369 / ST, E g i f ií. '- WR lx- N _motor, especially for with fixed n control of the amount of intake taken for propulsion of control- fully rotationally symmetrical air- Ivrotationally symmetrical cross-section, Ltt in the flow direction in front of the gas pipe (3) are arranged in a central body (7; 8). Device according to claim 1, wherein between the inside of the air intake resp. the central body1 and the adjacent one its boundary layer flow arranged bridged by a boundary knot none successively arranged by several in central opening and towards the boundary layer gap (7), the boundary layer t the pressure in the boundary layer gap of the central body (Z), at least to offset perpendicular compresses in front of the outflow. v Z, the characteristic (7) in cross section is the salmon tail y 1 or Z, the characteristic central body (2) in the flow direction and in the circumferential direction (10), which are closed with a hole (8) having lid (9) radially outwards and which open on both sides towards the boundary layer gap (5). E
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    同族专利:
    公开号 | 公开日
    GB2012370A|1979-07-25|
    US4502651A|1985-03-05|
    DE2801119C2|1982-12-02|
    DE2801119A1|1981-04-30|
    GB2012370B|1982-05-26|
    FR2496765B1|1987-07-31|
    FR2496765A1|1982-06-25|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3062484A|1953-04-10|1962-11-06|Curtiss Wright Corp|Supersonic air inlet construction|
    US3030770A|1957-07-30|1962-04-24|United Aircraft Corp|Variable supersonic inlet|
    US3046733A|1959-05-29|1962-07-31|Marquardt Corp|Acoustic buzz suppressor|
    US3477455A|1965-10-15|1969-11-11|Lockheed Aircraft Corp|Supersonic inlet for jet engines|
    DE1926553C3|1969-05-23|1973-10-18|Messerschmitt-Boelkow-Blohm Gmbh, 8000 Muenchen|
    US4007891A|1975-09-12|1977-02-15|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Jet engine air intake system|
    US4000869A|1975-10-21|1977-01-04|Northrop Corporation|Strong shock boundary layer interaction control system|DE3011796C2|1980-03-27|1982-11-11|Messerschmitt-Bölkow-Blohm GmbH, 8000 München|Supersonic air inlet for jet engines, in particular for ramjet engines for propelling missiles|
    DE3142463C1|1981-10-27|1983-07-07|Deutsche Forschungs- und Versuchsanstalt für Luft- und Raumfahrt e.V., 5000 Köln|Half rotationally symmetrical supersonic air inlet for reaction jet engines, especially for ram jet rocket engines powered by solid fuel|
    DE3142464C1|1981-10-27|1983-07-07|Deutsche Forschungs- und Versuchsanstalt für Luft- und Raumfahrt e.V., 5000 Köln|Half rotationally symmetrical supersonic air inlet for reaction jet engines, especially for ram jet rocket engines powered by solid fuel|
    DE3142465C2|1981-10-27|1984-02-23|Deutsche Forschungs- und Versuchsanstalt für Luft- und Raumfahrt e.V., 5000 Köln|Semi-rotationally symmetrical supersonic air intake for recoil engines, especially for ramjet rocket engines operated with solid fuels|
    DE3236487C2|1982-10-01|1984-07-19|Deutsche Forschungs- und Versuchsanstalt für Luft- und Raumfahrt e.V., 5000 Köln|Rotationally symmetrical supersonic air intake for jet engines|
    US4611616A|1984-01-10|1986-09-16|Messerschmitt-Bolkow-Blohm Gmbh|Axially semisymmetrical supersonic air intake for reaction engines, particularly solid fuel ram jet rocket engines|
    US4957242A|1988-04-12|1990-09-18|The United States Of America As Represented By The Secretary Of The Navy|Fluid mixing device having a conical inlet and a noncircular outlet|
    GB9424495D0|1994-12-05|1995-01-25|Short Brothers Plc|Aerodynamic low drag structure|
    US5749542A|1996-05-28|1998-05-12|Lockheed Martin Corporation|Transition shoulder system and method for diverting boundary layer air|
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    US20150315966A1|2014-05-01|2015-11-05|The Boeing Company|Hypersonic Vehicle Base Drag Reduction and Improved Inlet Performance Through Venting Forebody Bleed Air to Base Area Using Open Core Ceramic Composites|
    RU2672825C2|2017-04-17|2018-11-19|Федеральное государственное бюджетное учреждение науки Институт теоретической и прикладной механики им. С.А. Христиановича Сибирского отделения Российской академии наук |Supersonic air intake |
    CN109944701B|2019-03-19|2021-06-18|南京航空航天大学|External pressure type supersonic air inlet channel|
    CN113236424A|2021-06-22|2021-08-10|西安航天动力研究所|Double-lower-side rear supersonic air inlet|
    法律状态:
    2012-11-13| NAV| Patent application has lapsed|
    优先权:
    申请号 | 申请日 | 专利标题
    DE2801119A|DE2801119C2|1978-01-12|1978-01-12|Supersonic air inlet of fixed geometry for air-breathing recoil engines of steerable missiles or aircraft, in particular for stagnation drives operated with solid fuels|
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