CN108894893B - Liquid film cooling ejection rocket engine thrust chamber for rocket stamping combined engine - Google Patents
Liquid film cooling ejection rocket engine thrust chamber for rocket stamping combined engine Download PDFInfo
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- CN108894893B CN108894893B CN201810671792.3A CN201810671792A CN108894893B CN 108894893 B CN108894893 B CN 108894893B CN 201810671792 A CN201810671792 A CN 201810671792A CN 108894893 B CN108894893 B CN 108894893B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/52—Injectors
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
- F02K9/64—Combustion or thrust chambers having cooling arrangements
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Abstract
The invention discloses a liquid film cooling ejection rocket engine thrust chamber for a rocket stamping combined engine, which comprises a head, wherein the head comprises a head shell, and one end of the head shell is provided with an ignition device; the head shell is also provided with an oxygen direct-connection joint and a kerosene direct-connection joint which are both communicated to the gas-liquid coaxial nozzle; the gas-liquid coaxial nozzle comprises an oxygen direct-current jet orifice which is arranged in the head shell and is coaxial with the head shell, and also comprises a sectional direct-current slit nozzle for kerosene circulation, wherein the sectional direct-current slit nozzle is a three-section slit which is arranged around the outer side of the oxygen direct-current jet orifice; the other end of the head shell is coaxially communicated with the liquid film flange plate, the thrust chamber shell and the graphite spray pipe through a graphite gasket. The problem of among the prior art traditional passive cooling engine can't work for a long time, jacket cooling thrust chamber body structure complicacy is solved.
Description
[ technical field ] A method for producing a semiconductor device
The invention belongs to the technical field of liquid rocket engines, and particularly relates to a liquid film cooling ejection rocket engine thrust chamber for a rocket ramjet combined engine.
[ background of the invention ]
The thrust chamber is the primary site for propellant combustion heat release in liquid rocket engines. In the thrust chamber, the propellant finishes the processes of atomization, evaporation, mixing and combustion after being sprayed out by the nozzle, converts chemical energy into heat energy of high-temperature gas, and converts the heat energy of the gas into kinetic energy through the Laval nozzle with the contraction-expansion molded surface to generate thrust.
Generally, in order to improve the combustion efficiency and the specific impulse performance of an engine, on one hand, a high-performance injector needs to be designed to improve the atomization and mixing quality of a propellant; on the other hand, the thrust chamber is required to work at higher room temperature and room pressure, so that the expansion work-doing capacity of the fuel gas is improved. The severe thermal environment of the thrust chamber exceeds the heat resistance limit of many existing metal materials, and therefore necessary thermal protection measures need to be carried out on the thrust chamber to ensure stable and reliable operation of the engine.
At present, the common cooling modes of the thrust chamber mainly comprise external cooling, internal cooling, ablation cooling, heat-containing cooling and the like. In general, for small thrust engines, it is possible to operate for a short period of time without cooling the thrust chamber, and the heat from the thrust chamber can be absorbed by the thicker chamber walls as a heat sink, or passively thermally protected in the form of a body fitted with a thermally insulating lining. In the prior publication of RBCC ejection rocket combustion chamber design and experimental research (propulsion technology, 2014, 35(10)), a red copper heat-containing structure integrating a thrust chamber body and a nozzle is adopted, and in the patent CN104329187A, a high silica phenolic insulation lining is additionally arranged on the thrust chamber body.
For long-time working conditions, the thrust chamber needs to be designed with active thermal protection. The water-cooling thrust chamber body structure adopted in the research on the main rocket engine gas oxygen/kerosene thrust chamber of the RBCC (rocket propulsion, 2009, 35(6)) is commonly used in a ground verification test of the thrust chamber, but an additional cooling water supply system is required, so that the complexity of the system is increased, and meanwhile, the thrust chamber body needs to be designed into a jacket structure, so that the forming process is complex and the cost is high. The liquid film cooling is used as a mode of cooling in the thrust chamber, has the advantages of simple structure and convenient introduction, and is widely used for cooling the side area of the thrust chamber of a large-scale engine. The cooling agent is injected in a liquid state, spreads on the wall surface and absorbs heat through phase change evaporation, and meanwhile, the low-temperature fuel-rich gas film formed on the wall surface protects the wall surface from being washed by high-temperature hot air flow, so that the temperature of the wall surface can be reduced.
In addition, the injector, as a core component of the liquid rocket engine head, controls the atomization and mixing of the propellant components, thereby affecting the combustion efficiency, combustion stability and cooling of the thrust chamber of the propellant. The gas-liquid coaxial straight-flow nozzle is a common gas-liquid nozzle form, has the characteristics of simple structure and stable work, and can also have the problems of nonuniform atomization and mixing of a propellant, long combustion interval and the like.
[ summary of the invention ]
The invention aims to provide a liquid film cooling ejection rocket engine thrust chamber for a rocket stamping combined engine, and aims to solve the problems that a traditional passive cooling engine cannot work for a long time in the prior art, and the structure of a jacket cooling thrust chamber body is complex.
The invention adopts the following technical scheme: the liquid film cooling ejection rocket engine thrust chamber for the rocket stamping combined engine is characterized in that the thrust chamber comprises a head, the head comprises a head shell, the head shell is two cylinders which are coaxially arranged and have different diameters, and one end of the head shell is provided with an ignition device;
the head shell is also provided with an oxygen direct-connection joint and a kerosene direct-connection joint which are both communicated to the gas-liquid coaxial nozzle; the gas-liquid coaxial nozzle comprises an oxygen direct-current jet orifice which is arranged in the head shell and is coaxial with the head shell, and also comprises a sectional direct-current slit nozzle for kerosene circulation, wherein the sectional direct-current slit nozzle is a three-section slit which is arranged around the outer side of the oxygen direct-current jet orifice;
the other end of the head shell is coaxially communicated with the liquid film flange plate, the thrust chamber shell and the graphite spray pipe through a graphite gasket, the inner contour of the graphite spray pipe forms a Laval spray pipe molded surface for accelerating the expansion of high-temperature gas, and the spray pipe gland covers the outer part of the graphite spray pipe to enable the graphite spray pipe to be tightly attached to the thrust chamber shell.
Furthermore, the oxygen straight-through joint is communicated to the oxygen straight-flow jet hole through the oxygen cavity, and the kerosene straight-through joint is communicated to the slit through the coal oil cavity.
Furthermore, the oxygen chamber and the coal oil chamber are both annular chambers which are coaxial with the shell, and the coal oil chamber is closer to the lower panel than the oxygen chamber.
Further, the flange disc with the liquid membrane method comprises a flange disc body, the side surface of the flange disc body is communicated with a kerosene through joint, the kerosene through joint is communicated with a liquid membrane cavity arranged in the flange disc body,
the liquid film cavity is an annular cavity which is arranged around the axis of the flange plate body, the liquid film cavity is communicated with the thrust chamber and is provided with a plurality of liquid film cooling spray holes, and the included angle between the axis of each liquid film cooling spray hole and the axis of the inner wall surface of the thrust chamber shell is 15 degrees.
Furthermore, the oxygen straight joint is arranged on the upper panel, and the kerosene straight joint is arranged on the side surface of the head shell.
Further, ignition includes that the center department of last panel is fixed to be provided with the spark plug seat, installs spark plug ignition in the spark plug base, and panel 2mm under the top surface distance of spark plug ignition.
Furthermore, the upper panel of the head, the side surface of the head and the side surface of the liquid membrane flange body are connected with pressure measuring seats.
Furthermore, five temperature measuring seats and three pressure measuring seats are uniformly distributed on the thrust chamber shell along the axial direction of the thrust chamber along the way, and the five temperature measuring seats are circumferentially arranged in a staggered manner.
Compared with the prior art, the invention has at least the following beneficial effects: the head nozzle is simple and efficient in structure, and the external kerosene is pneumatically atomized by the oxygen jet ejected from the center of the nozzle at a high speed, so that a good atomization and mixing effect can be achieved; because the liquid film flange is independent of the design of the head and the body, the liquid film introducing mode can be controlled only by replacing the liquid film flange with different liquid film spray hole apertures, numbers and incidence angles.
[ description of the drawings ]
FIG. 1 is a schematic structural diagram of a liquid film cooling ejection rocket engine thrust chamber for a rocket ramjet combined engine according to the present invention;
FIGS. 2 a-2 b are front and left views of a liquid-film-cooled ejector rocket engine thrust chamber gas-liquid coaxial nozzle for a rocket ramjet combined engine according to the present invention;
FIGS. 3 a-3 b are front and top views, respectively, of a liquid film flange plate of a liquid film cooling ejection rocket engine thrust chamber for a rocket ramjet combined engine according to the present invention;
FIG. 3c is an enlarged view of FIG. 3b at c;
FIG. 4 is a perspective view of a liquid film cooling ejection rocket engine thrust chamber for the rocket-ramjet combined engine according to the present invention;
FIG. 5 is a diagram of experimental data of a liquid film-free thrust chamber ignition experiment of a liquid film-cooled ejector rocket engine thrust chamber for a rocket ramjet combined engine according to the present invention;
FIG. 6 is a diagram of experimental data of liquid film-based thrust chamber ignition in a liquid film-cooled ejector rocket engine thrust chamber for a rocket ramjet combined engine according to the present invention;
wherein: 1. a spark plug igniter; 2. an oxygen straight-through joint; 3. a pressure measuring seat; 4. a head housing; 5. an oxygen chamber; 6. a coal oil cavity; 7. a gas-liquid coaxial nozzle; 8. a kerosene straight-through joint; 9. a liquid film flange plate; 10. a liquid film cavity; 11. a thrust chamber housing; 12. a spark plug seat; 13. a temperature measuring seat; 14. a graphite nozzle; 15. a spray pipe gland; 16. a graphite gasket; 17. oxygen direct current jet orifice; 18. a slit; 19. an upper panel; 20. and a lower panel 21. the liquid film cools the spray hole.
[ detailed description ] embodiments
The technical solution of the present invention is further described in detail by the accompanying drawings and embodiments.
The invention provides a liquid film cooling ejection rocket engine thrust chamber for a rocket stamping combined engine, as shown in figure 1, the thrust chamber comprises a head part, the head part comprises a head part shell 4, the head part shell 4 is two cylinders which are coaxially arranged and have different diameters, one end surface is an upper panel 19, the other end surface is a lower panel 20, and an ignition device is arranged at the position of the upper panel 19; the head shell 4 is provided with an oxygen through joint 2 and a kerosene through joint 8, and the oxygen through joint 2 and the kerosene through joint 8 are both communicated to a gas-liquid coaxial nozzle 7; the gas-liquid coaxial nozzle 7 comprises an oxygen direct current jet orifice 17 which is arranged in the head shell 4 and is coaxial with the head shell, and also comprises a sectional direct current slit nozzle for kerosene circulation, wherein the sectional direct current slit nozzle is a three-section slit 18 which is arranged around the outer side of the oxygen direct current jet orifice 17; the lower panel 20 is coaxially connected to the membrane flange 9, the thrust housing 11 and the graphite nozzle 14 through a graphite gasket 16, the inner contour of the graphite nozzle 14 forms a laval nozzle profile for accelerating the expansion of the high-temperature gas, and the nozzle cover 15 covers the graphite nozzle 14 so as to be tightly attached to the thrust housing 11.
The gas-liquid coaxial nozzles 7 are symmetrically distributed on two sides of the circumference with the center point of the head as the center of a circle, the number of the nozzles of the gas-liquid coaxial nozzles 7 is two, a central oxygen direct-current spray hole and an outer kerosene sectional direct-current slit nozzle form are adopted, and in consideration of the processing precision problem, if the kerosene nozzle adopts an integral circular seam form, the circular seam is too small to be processed easily, so that three slit channels are formed. The upper and lower outer walls of the coal chamber are respectively welded with the oxygen chamber 5 and the baffle plate of the coal chamber 6 in a seamless manner, the oxygen chamber 5 is communicated with the thrust chamber through a central hole to form a channel for leading oxygen to enter the thrust chamber, and the coal chamber 6 is communicated with the thrust chamber through an outer sectional slit to form a channel for leading kerosene to enter the thrust chamber, as shown in fig. 2 b.
The gas-liquid coaxial nozzle 7 is simple in structure and convenient to process, and the problem that the coaxiality of the annular seam nozzle is difficult to control is solved by the design of the kerosene sectional type direct current slit. The advantages of atomization mixing, stable combustion and the like are considered, and the problems of complex structure and the like of a conventionally used gas-liquid swirl nozzle are solved. The high-speed oxygen jet flow sprayed out from the center of the nozzle realizes pneumatic atomization of the external kerosene, so that a good atomization and mixing effect can be achieved.
The oxygen straight-through joint 2 is communicated with the oxygen straight-through spray hole 17 through the oxygen chamber 5, and the kerosene straight-through joint 8 is communicated with the slit 18 through the coal oil chamber 6. The head shell 4 is welded with the oxygen through joint 2, the pressure measuring seat 3, the kerosene through joint 8, the spark plug seat 12 and the gas-liquid coaxial nozzle 7 into a whole to form the head part of the thrust chamber. The oxygen chamber 5 and the coal oil chamber 6 are both positioned in the head, the oxygen chamber 5 and the coal oil chamber 6 are both annular chambers coaxially arranged with the shell, the coal oil chamber 6 is arranged to be closer to the lower panel than the oxygen chamber 5, and the lower panel of the head shell 4 is cooled by utilizing the liquid cooling effect of the kerosene in the coal oil chamber 6.
The liquid film method flange 9 comprises a flange body, the side surface of the flange body is communicated with a kerosene through joint 8, the kerosene through joint 8 is communicated with a liquid film cavity 10 arranged in the flange body, the liquid film cavity is an annular cavity arranged around the axis of the flange body, the liquid film cavity is communicated with a plurality of liquid film cooling spray holes 21 towards the direction of a thrust chamber, and the included angle between the axis of each liquid film cooling spray hole 21 and the axis of the inner wall surface of a thrust chamber shell 11 is 15 degrees. If the incidence angle of the liquid film cooling spray hole 21 is too large, liquid drop sputtering can occur after the jet flow collides the wall, so that part of the cooling liquid can not participate in the cooling process; and the free stroke of the jet flow with the small incident angle is too large, so that the jet flow is easier to generate physical and chemical reaction with main stream gas in a combustion chamber, and the cooling efficiency is reduced.
In practice, the liquid membrane method flange plate 9 is welded with the pressure measuring seat 3 and the kerosene through joint 8 into a whole; the liquid film cavity 10 is positioned in the liquid film flange 9 and is communicated with the thrust chamber through liquid film cooling spray holes, as shown in figure 3. The number of the liquid film cooling spray holes 21 can be 32, and the liquid film cooling spray holes are symmetrically and uniformly distributed on the circumference with the center point of the liquid film flange as the center. This quantity is mainly decided by the indoor wall diameter of thrust that will cool off, and this quantity need guarantee that the indoor wall circumference of thrust all will have the cover of liquid film for the indoor wall circumference of thrust all is in under the cooling action of liquid film.
The formation of liquid film usually has the head to set up limit district nozzle and body and sets up the cooling clitellum dual mode, and the design of liquid film cooling orifice is similar to setting up limit district direct current nozzle at the head, compares in setting up the cooling clitellum in body, and this structure is simpler, and can not influence thrust chamber body structure. The design of the liquid film flange plate enables a liquid film to be conveniently and controllably introduced, and the liquid film flange plate is independent of the design of the head part and the body part, so that on one hand, the introduction mode of the liquid film can be controlled by only replacing the liquid film flange plates with different liquid film spray hole apertures, numbers and incidence angles under the condition of not changing the structures of the head part and the body part, and the influence of the factors on the cooling effect of the liquid film is conveniently researched through experiments; on the other hand, the liquid film path kerosene and the head main path kerosene are mutually independent, and the kerosene flow for cooling the liquid film can be controlled by changing the parameters of the liquid film path kerosene supply system.
The oxygen through-connection 2 is provided on the upper panel 19, and the kerosene through-connection 8 is provided on the side of the head housing 4. The ignition device comprises a spark plug seat 12 fixedly arranged at the center of the upper panel 19, a spark plug type igniter 1 is arranged in the spark plug seat 12, and the top surface of the spark plug type igniter 1 is 2mm away from the lower panel so as to prevent the spark plug from being ablated by high-temperature gas.
The thrust chamber body part consists of a thrust chamber shell 11, a graphite nozzle 14 and a nozzle gland 15. Wherein, thrust chamber casing 11 has evenly distributed 5 temperature measurement seats 13 and 3 pressure measurement seats 3 along the journey along thrust chamber axis direction for measure thrust chamber during operation along axial pressure and temperature variation. Due to the limitation of the structural size of the body part, the 5 temperature measuring seats 13 are arranged in a staggered mode along the circumferential direction, as shown in figure 4. The inner contour of the graphite lance 14 forms the laval lance profile necessary for accelerating the expansion of the hot gas and is pressed into close contact with the thrust chamber housing 11 by the lance cover 15. The head 4, the liquid membrane flange 9, the thrust housing 11, and the nozzle cover 15 are sealed with graphite gaskets 16, and are connected by bolts.
The upper panel 19 of the head, the side surface of the head and the side surface of the liquid membrane flange 9 body are connected with a pressure measuring seat 3. Five temperature measuring seats 13 and three pressure measuring seats 3 are uniformly distributed on the thrust chamber shell 11 along the axial direction of the thrust chamber along the way, and the five temperature measuring seats 13 are arranged in a staggered mode along the circumferential direction. Here three pressure measuring seats 3 are used to measure the distribution of the chamber pressure in the thrust chamber along the way; the five temperature measuring seats 13 are circumferentially staggered, so that the thrust chamber cannot be arranged on one axis due to the limitation of the structural size of the thrust chamber, and the five temperature measuring seats 13 uniformly distributed along the way are used for measuring the axial distribution of the wall temperature of the thrust chamber, so that the spreading length and the cooling effect of a cooling liquid film can be deduced.
Example (b):
in the experiment i, the engine is assembled as shown in fig. 1, wherein the spark plug type igniter 1, the oxygen through joint 2, the pressure measuring seat 3, the head 4, the gas-liquid coaxial nozzle 7, the kerosene through joint 8, the liquid membrane flange 9, the thrust chamber shell 11, the temperature measuring seat 13 and the spray pipe gland 15 are all made of stainless steel 1Cr18Ni9Ti, and the spray pipe is made of high-temperature resistant graphite in consideration of the severe thermal environment of the throat. The spark plug type igniter 1 and the head spark plug seat are in threaded connection by adopting a liquid film cooling injection rocket engine thrust chamber 1 for an M16 rocket punching combined engine; the length of the thrust chamber shell 11 is 210mm, the wall thickness is 15mm, the outer diameter is 102mm, and the inner diameter is 72 mm; the number of holes is 8 in equal distance of the flange of the head 4, the flange disc 9 of the liquid membrane method and the upper flange and the lower flange of the thrust chamber shell 11, the aperture is 13mm, the wall thickness of the flange of the head is 20mm, and the wall thickness of the other flanges is 15 mm; the length of the head gas-liquid coaxial nozzle is 29mm, the aperture of the oxygen is 5mm, the inner diameter of the kerosene slit is 6mm, the outer diameter of the kerosene slit is 7mm, the number of the kerosene slits is 3, and the included angle of the circle center occupied by each slit is 40 degrees; the aperture of the liquid film spray holes is 0.3mm, and the number of the liquid film spray holes is 32; the throat diameter of the graphite nozzle 14 is determined by experimental conditions.
In the engine, each connecting part adopts end face sealing, and the bolt and the nut are composed of GB/T5781-2000 hexagon head bolt-full bolt C-grade M12 l-55 and GB/T41-2000 hexagon head nut-C-grade M12.
ii, the invention discloses a thrust chamber of an injection rocket engine with a novel gas-liquid coaxial nozzle 7 and a liquid film flange plate, and the thrust chamber is tested through experiments. Firstly, cold debugging is carried out, and the method mainly comprises oxygen and kerosene flow calibration, filling time measurement and the like. After cold debugging is finished, counting down to ignite hot test run after all detection is correct.
In the working process of the liquid film cooling ejection rocket engine thrust chamber for the rocket stamping combined engine, a thrust chamber shell 11 is fixed on a test bed, a head oxygen straight-through joint 2 and a kerosene straight-through joint 8 are respectively connected with oxygen and kerosene supply pipelines, wherein a kerosene supply system is divided into two paths, one path is communicated with a head coal oil cavity 6, and the other path is communicated with a liquid film cavity 10 in a liquid film flange plate 9. The two paths of kerosene are mutually independent, and the flow rates of the two paths of kerosene can be respectively controlled according to the requirements of working conditions so as to achieve the purpose of controlling the flow rates of the main path kerosene and the liquid film kerosene.
Oxygen and kerosene are respectively introduced into an oxygen cavity 5 and a coal oil cavity 6 in the head 4 and atomized and mixed under the action of certain pressure drop through a gas-liquid coaxial nozzle 7; liquid film cooling kerosene enters a liquid film cavity 10 in a liquid film flange 9 through a liquid film kerosene straight-through joint and is sprayed into the wall surface of the thrust chamber through a liquid film spray hole to realize liquid film cooling.
Two paths of kerosene are controlled to simultaneously enter the thrust chamber through a time sequence, and oxygen is controlled to enter the thrust chamber before the kerosene enters the thrust chamber for 400ms, so that an oxygen-rich atmosphere is formed before ignition, and the successful ignition is facilitated. When oxygen enters the thrust chamber, the spark plug type igniter 1 is started after being excited by the high-energy igniter and is closed after the engine works for 1s, so that the kerosene accumulation danger caused by long-time misfire is prevented. After entering the thrust chamber, the main path kerosene and the liquid film kerosene are atomized and combusted in the thrust chamber to form high-temperature and high-pressure fuel gas, and the fuel gas is sprayed out through the graphite spray pipe 14. When the engine is shut down, the oxygen is shut down for 200ms before the kerosene to ensure that the engine can be immediately stopped, and when the kerosene is shut down, the nitrogen is blown off to prevent tempering by blowing off the kerosene in the pipeline, the head and the liquid film flange.
In the experiment, the key working parameters are head oxygen and kerosene flow, liquid film kerosene flow, combustion chamber pressure, body near-wall temperature and the like.
The oxidant of the thrust room experiment adopts oxygen, the fuel adopts JP-10 kerosene, and the experiment comprises a liquid film-free experiment and a liquid film experiment. The experimental working condition is that the total flow is 400g/s, the oxygen-fuel ratio is 1.3, the room pressure is 2MPa, and the throat diameter of the spray pipe is 18 mm. All kerosene is sprayed out through the head nozzle in the liquid-film-free experiment; in the liquid film test, the main path kerosene flow is 139g/s, and the liquid film kerosene flow is 35 g/s.
After the test run is finished, the collected pressure and temperature data are processed as shown in fig. 5, and the performance and wall surface thermal environment of the engine in the working process with or without a liquid film can be obtained. As can be seen from FIG. 5, the maximum temperature of the body of the thrust chamber is 1100 ℃ when the engine works without a liquid film, and the maximum temperature of the body of the engine works with the liquid film is reduced to 900 ℃, which shows that the liquid film cooling has a good effect. In two experiments, the engine works normally, the chamber pressure is close to the design value, and the nozzle design achieves better combustion efficiency.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like, indicate orientations or positional relationships based on those shown in the drawings, and are used only for convenience in describing the present invention and for simplicity in description, and do not indicate or imply that the referenced devices or elements must have a particular orientation, be constructed and operated in a particular orientation, and thus, are not to be construed as limiting the present invention. Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of the present invention, "a plurality" means two or more unless otherwise specified.
In the description of the present invention, it should be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
The above-mentioned contents are only for illustrating the technical idea of the present invention, and the protection scope of the present invention is not limited thereby, and any modification made on the basis of the technical idea of the present invention falls within the protection scope of the claims of the present invention.
Claims (6)
1. The liquid film cooling ejection rocket engine thrust chamber for the rocket stamping combined engine is characterized in that the thrust chamber comprises a head, the head comprises a head shell (4), the head shell (4) is two cylinders which are coaxially arranged and have different diameters, and one end of the head shell (4) is provided with an ignition device;
the head shell (4) is also provided with an oxygen through joint (2) and a kerosene through joint (8), and the oxygen through joint (2) and the kerosene through joint (8) are communicated to a gas-liquid coaxial nozzle (7); the gas-liquid coaxial nozzle (7) comprises an oxygen direct-current jet hole (17) which is arranged in the head shell (4) and is coaxial with the head shell, and further comprises a sectional direct-current slit nozzle for kerosene circulation, wherein the sectional direct-current slit nozzle is a three-section slit (18) which is arranged around the outer side of the oxygen direct-current jet hole (17);
the other end of the head shell (4) is coaxially communicated with the liquid film flange plate (9), the thrust chamber shell (11) and the graphite spray pipe (14) through a graphite gasket (16), the inner contour of the graphite spray pipe (14) forms a Laval spray pipe profile which enables high-temperature gas to expand and accelerate, and a spray pipe gland (15) covers the outer part of the graphite spray pipe (14) to enable the graphite spray pipe (14) to be tightly attached to the thrust chamber shell (11);
the oxygen straight-through joint (2) is communicated to the oxygen straight-through spray hole (17) through an oxygen cavity (5), and the kerosene straight-through joint (8) is communicated to the slit (18) through a coal oil cavity (6);
the oxygen chamber (5) and the coal oil chamber (6) are both annular chambers coaxially arranged with the shell, and the coal oil chamber (6) is closer to the lower panel than the oxygen chamber (5).
2. The liquid film cooling ejector rocket engine thrust chamber for the rocket stamping combination engine as recited in claim 1, wherein said liquid film flange (9) comprises a flange body, the side surface of said flange body is communicated with kerosene through joint (8), said kerosene through joint (8) is communicated with liquid film cavity (10) arranged in the flange body,
the liquid film cavity is an annular cavity arranged around the axis of the flange plate body, the liquid film cavity is communicated with a plurality of liquid film cooling spray holes (21) towards the thrust chamber, and the included angle between the axis of each liquid film cooling spray hole (21) and the axis of the inner wall surface of the thrust chamber shell (11) is 15 degrees.
3. The liquid film-cooled ejector rocket engine thrust chamber for a rocket ramjet combined engine according to claim 1 or 2, wherein said oxygen through joint (2) is provided on an upper panel (19), and said kerosene through joint (8) is provided on a side surface of a head casing (4).
4. The liquid film cooling ejector rocket engine thrust chamber for the rocket stamping combined engine as recited in claim 1 or 2, characterized in that said ignition device comprises a spark plug seat (12) fixedly arranged at the center of the upper panel (19), a spark plug igniter (1) is installed in the spark plug seat (12), and the top surface of said spark plug igniter (1) is 2mm away from the lower panel.
5. The liquid film cooling ejector rocket engine thrust chamber for the rocket stamping combination engine as recited in claim 1 or 2, characterized in that the upper panel (19) of the head, the side surface of the head and the side surface of the liquid film flange plate (9) body are all connected with a pressure measuring seat (3).
6. The liquid film cooling ejector rocket engine thrust chamber for the rocket stamping combined engine as recited in claim 1 or 2, characterized in that five temperature measuring seats (13) and three pressure measuring seats (3) are uniformly distributed on the thrust chamber shell (11) along the axial direction of the thrust chamber along the way, and the five temperature measuring seats (13) are arranged in a staggered way along the circumferential direction.
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