Nozzle, nozzle assembly, rocket engine and rocket engine assembly

Abstract

The invention discloses a nozzle, a nozzle assembly, a rocket engine and a rocket engine assembly. The nozzle comprises a nozzle body. A unilateral expansion nozzle is obliquely arranged on the side wall of the nozzle end of the nozzle body and penetrates the nozzle body. The air inlet end of the unilateral expansion nozzle is located on the outer wall of the nozzle body, and the air outlet end of the unilateral expansion nozzle is located on the inner wall of the nozzle body. The air outlet end of the unilateral expansion nozzle faces the nozzle end of the nozzle body. The air inlet end and the air outlet end of the unilateral expansion nozzle are both in a flared shape. A throat is formed at the middle of the unilateral expansion nozzle. The angle between the axis of the throat and the axis of the nozzle body is θ1, and θ1 is 10-20°. By directly processing the unilateral expansion nozzle on the nozzle body, the processing cost is low. At the same time, the nozzle can accelerate the airflow to supersonic speed without increasing the weight. The supersonic airflow can directly blow away the separation boundary layer with the adverse pressure gradient in the nozzle body.

Description

Technical Field

[0001] The present invention belongs to the technical field of rocket engines, and in particular relates to a nozzle, a nozzle assembly, a rocket engine and a rocket engine assembly. Background Art

[0002] Liquid rocket engines are divided into open cycle and closed cycle according to the circulation mode. The open cycle includes gas generator cycle, open expansion cycle, exhaust cycle and other modes. Among them, the open cycle loses some performance because the working fluid driving the turbine is discharged into the environment at a lower speed. Generally, the specific impulse of the open engine with the same propellant is lower than that of the closed engine. In addition, the nozzle expansion ratio will also greatly affect the performance of the engine. A larger nozzle expansion ratio can make the fuel expand more fully in the nozzle, thereby obtaining a higher exhaust velocity and a higher specific impulse, but an excessively large expansion ratio will cause nozzle flow separation, resulting in a limitation of the expansion ratio, which will inhibit the improvement of nozzle performance. At present, there is a way to suppress separation by ejecting a high-speed airflow from the outside of the nozzle and using the ejection effect of the jet, but this method requires an external air source and an additional nozzle structure. Due to its large mass, it can only be applied to the test bench and the ignition stage of the rocket. Summary of the invention

[0003] In order to solve the above technical problems, the object of the present invention is to provide a nozzle with a simple structure and high integration, which can suppress nozzle flow separation and improve the expansion ratio.

[0004] In order to achieve the above-mentioned purpose, the technical scheme of the present invention is as follows: a nozzle, comprising a nozzle body, a unilateral expansion nozzle penetrating the side wall of the nozzle end of the nozzle body is obliquely arranged, the air inlet end of the unilateral expansion nozzle is located on the outer wall of the nozzle body, the air outlet end of the unilateral expansion nozzle is located on the inner wall of the nozzle body, and the air outlet end of the unilateral expansion nozzle faces the nozzle end of the nozzle body, the air inlet end and the air outlet end of the unilateral expansion nozzle are both flared, the middle part of the unilateral expansion nozzle forms a throat, and the angle between the axis of the throat and the axis of the nozzle body is θ 1 , and θ 1 It is 10-20°.

[0005] The beneficial effect of the above technical solution is that: by directly processing a unilateral expansion nozzle on the nozzle body, the processing cost is low, and at the same time, the nozzle accelerates the airflow to supersonic speed without increasing the weight. When the supersonic airflow is blown into the nozzle end inside the nozzle body, it can act on the separation flow field (the separation boundary layer with a reverse pressure gradient in the nozzle body is directly blown away), and the ability to suppress flow separation in the nozzle body is higher than passive injection outside the nozzle body. In addition, the supersonic airflow can also increase the thrust and specific impulse, and the angle between the axis of the throat and the axis of the nozzle body is set to 10-20°, so that the exhaust speed of the unilateral expansion nozzle can be converted into thrust to a large extent.

[0006] In the above technical solution, the distance between the gas outlet end of the unilateral expansion nozzle and the nozzle end of the nozzle body is L, wherein the maximum value of L is 0.1Lp;

[0007] When the nozzle body is a conical nozzle, the minimum value of L is 0;

[0008] When the nozzle body is a bell-shaped nozzle, the minimum value of L is 0.01Lp;

[0009] Wherein, Lp is the length of the expansion section of the nozzle body.

[0010] The beneficial effect of the above technical solution is that: this is because the outlet pressure of the single-sided expansion nozzle at the outlet end is smaller when it is closer to the nozzle end of the nozzle body, which is conducive to the single-sided expansion nozzle to obtain a larger jet velocity, so L cannot be too large; and in order to facilitate the processing of the single-sided expansion nozzle, for the conical nozzle, the minimum value of L is 0, and for the bell-shaped nozzle, the minimum value of L is 0.01Lp.

[0011] In the above technical solution, the cross-section of the unilateral expansion nozzle perpendicular to the axis at the throat is a rectangle; the height of the throat is Dc, and Dc is 3-10 mm.

[0012] The beneficial effects of the above technical solution are: it helps to simplify the nozzle processing process and can also ensure the reliability of the operation of the single-sided expansion nozzle.

[0013] In the above technical solution, the air inlet end and the air outlet end of the unilateral expansion nozzle are expanded toward the nozzle end of the nozzle body, and the unilateral expansion nozzle must meet at least one of the following conditions ae when expanding:

[0014] a: The width of the outlet end of the unilateral expansion nozzle is expanded to Do, and the expansion ratio of the outlet end of the unilateral expansion nozzle is ε 1 , ε 1 =Do / Dc,ε 1 The value of is 3-6;

[0015] b: The width of the inlet end of the unilateral expansion nozzle is expanded to Di, and the contraction ratio of the inlet end of the unilateral expansion nozzle is ε 2 , ε 2 =Di / Dc,ε 2 The value of is 1.5-3;

[0016] c: The angle between the flared surface at the outlet of the unilateral expansion nozzle and the center line at the throat is θ 2 , where θ 2 >θ 1 ;

[0017] d: The angle between the flared surface of the inlet end of the unilateral expansion nozzle and the axis at the throat is θ 3 , where θ 3 <90°-θ 1 ;

[0018] e: The throat is rounded on one side near the nozzle end of the nozzle body, and the radius of the rounded corner is R, Dc <R<2Dc。

[0019] The beneficial effects of the above technical solution are: it simplifies the processing flow, and enables the airflow to have a higher flow rate after being ejected through the unilateral expansion nozzle. At the same time, the airflow is smoother when ejected through the unilateral expansion nozzle without generating turbulence.

[0020] A second object of the present invention is to provide a nozzle assembly that has a simple structure, can suppress flow separation, has a large expansion ratio, and can improve the specific impulse of the rocket engine.

[0021] In order to achieve the above-mentioned purpose, the technical solution of the present invention is as follows: a nozzle assembly, comprising a ring sleeve and the nozzle as described above, the ring sleeve being coaxially sleeved on the outer wall of the nozzle end of the nozzle body, the ring sleeve and the nozzle body together enclose a ring-shaped air collecting chamber, the ring sleeve is provided with an air inlet connected to the air collecting chamber, a plurality of single-sided expansion nozzles are provided, and all are located in the air collecting chamber, and the plurality of single-sided expansion nozzles are distributed circumferentially at intervals.

[0022] The beneficial effect of the above technical solution is that multiple single-sided expansion nozzles are annularly spaced on the inner side of the nozzle end of the nozzle body to form numerous evenly distributed high-speed jets, which can improve the uniformity of the injection force distribution at the nozzle end of the nozzle and inhibit the reflux of the atmosphere into the nozzle body (inhibit jet separation).

[0023] A third object of the present invention is to provide a rocket engine that can suppress flow separation and has a higher specific impulse.

[0024] In order to achieve the above-mentioned purpose, the technical solution of the present invention is as follows: a rocket engine, comprising a combustion chamber and the nozzle assembly as described above, wherein the combustion chamber has a first oxidant inlet, a first fuel inlet and an injection port, and an end of the nozzle body away from its nozzle end is connected to the injection port of the combustion chamber.

[0025] The beneficial effect of the above technical solution is that by arranging the single-sided expansion nozzles at annular intervals at the nozzle end of the nozzle body, the rocket engine can suppress flow separation while having higher thrust and specific impulse.

[0026] A fourth object of the present invention is to provide a rocket engine assembly that has a simple structure, can suppress flow separation, and has a higher specific impulse.

[0027] In order to achieve the above-mentioned purpose, the technical solution of the present invention is as follows: A rocket engine assembly, comprising a turbine, an oxidizer pump, a fuel pump and the rocket engine as described above, the discharge port of the oxidizer pump is connected to the first oxidizer inlet, the discharge port of the fuel pump is connected to the first fuel inlet, the power input ends of the oxidizer pump and the fuel pump are both transmission-connected to the driving end of the turbine, the turbine has an air inlet and an exhaust gas outlet, the exhaust gas outlet of the turbine is connected to the air inlet, and the air inlet of the turbine is used to allow air flow to be driven.

[0028] The beneficial effect of the above technical solution is that the turbine can provide driving force for the oxidizer pump and the fuel pump, and at the same time discharge the exhaust gas of the turbine into the air collecting chamber, so that the exhaust gas of the turbine is sprayed into the nozzle body through the unilateral expansion nozzle to form a high-speed jet, thereby improving the thrust and specific impulse of the entire engine.

[0029] In the above technical solution, the combustion chamber is also provided with an air splitting port, and the air splitting port is communicated with the air inlet of the turbine.

[0030] The beneficial effect of the above technical solution is that part of the airflow generated during combustion in the combustion chamber can be introduced into the turbine to drive the turbine to operate, which makes its structure simple and has good operating reliability.

[0031] The above technical solution also includes a gas generator, which has a second oxidant inlet, a second fuel inlet and an exhaust port, the second oxidant inlet is connected to the discharge port of the oxidant pump, the second fuel inlet is connected to the discharge port of the fuel pump, and the exhaust port is connected to the air inlet of the turbine.

[0032] The beneficial effect of the above technical solution is that: by independently arranging a gas generator, part of the oxidant and fuel delivered by the oxidant pump and the fuel pump are burned to generate gas to drive the turbine to operate.

[0033] The above technical solution also includes a heat exchange jacket, which is arranged in an interlayer within the wall of the combustion chamber and / or the nozzle body, and has a third fuel inlet and an external discharge port. The third fuel inlet is connected to the discharge port of the fuel pump, and the external discharge port is connected to the air inlet of the turbine.

[0034] The beneficial effect of the above technical solution is that the distributed fuel can be introduced into the heat exchange jacket for heating and vaporization to serve as the airflow to drive the turbine. At the same time, the vaporized fuel is injected into the nozzle body through a unilateral expansion nozzle for combustion to further increase the thrust and specific impulse of the rocket engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 is a cross-sectional view of the nozzle described in Example 1 of the present invention;

[0036] Figure 2 Schematic diagram of processing of rectangular grooves in Example 1 of the present invention;

[0037] Figure 3 It is a schematic diagram of the expansion of both ends of the rectangular channel in Example 1 of the present invention;

[0038] Figure 4 Schematic diagram of the chamfered throat in Example 1 of the present invention;

[0039] Figure 5 is a schematic diagram of the nozzle assembly described in Example 2 of the present invention;

[0040] Figure 6 This is a schematic diagram of a rocket engine according to Embodiment 3 of the present invention;

[0041] Figure 7 Schematic diagram of the rocket engine assembly according to Embodiment 4 of the present invention;

[0042] Figure 8 This is a schematic diagram of a rocket engine assembly according to Embodiment 5 of the present invention;

[0043] Fig. 9 This is a schematic diagram of a rocket engine assembly according to Embodiment 6 of the present invention;

[0044] Fig.10 This is a schematic diagram of two turbines connected in series in Example 7 of the present invention;

[0045] Fig.11 This is a schematic diagram of two turbines connected in parallel in Example 7 of the present invention.

[0046] In the figure: 1. Rocket engine; 11. Nozzle assembly; 111. Nozzle; 1111. Nozzle body; 1112. Single-sided expansion nozzle; 11121. Throat; 112. Ring sleeve; 1121. Air inlet; 113. Gas collecting cavity; 12. Combustion chamber; 121. First oxidant inlet; 122. First fuel inlet; 123. Injection port; 124. Gas distributor; 2. Turbine; 3. Oxidant pump; 4. Fuel pump; 5. Gas generator; 51. Second oxidant inlet; 52. Second fuel inlet; 53. Exhaust port; 6. Heat exchange jacket; 61. Third fuel inlet; 62. External exhaust port. DETAILED DESCRIPTION

[0047] The principles and features of the present invention are described below in conjunction with the accompanying drawings. The examples given are only used to explain the present invention and are not used to limit the scope of the present invention. The present invention is described in more detail by way of example with reference to the accompanying drawings in the following paragraphs. The advantages and features of the present invention will become clearer according to the following description and claims. It should be noted that the drawings are all in a very simplified form and are not in precise proportions, and are only used to conveniently and clearly assist in explaining the purpose of the embodiments of the present invention.

[0048] Example 1

[0049] like Figure 1 and Figure 2 As shown, this embodiment provides a nozzle, including a nozzle body 1111, a unilateral expansion nozzle 1112 is obliquely arranged on the side wall of the nozzle end of the nozzle body 1111 and penetrates therethrough, the air inlet end of the unilateral expansion nozzle 1112 is located on the outer wall of the nozzle body 1111, the air outlet end of the unilateral expansion nozzle 1112 is located on the inner wall of the nozzle body 1111, and the air outlet end of the unilateral expansion nozzle 1112 faces the nozzle end of the nozzle body 1111, the air inlet end and the air outlet end of the unilateral expansion nozzle 1112 are both flared, and a throat 11121 is formed at the middle of the unilateral expansion nozzle 1112, and the angle between the axis of the throat 11121 and the axis of the nozzle body 1111 is θ 1 , and θ 1The angle of the nozzle body is 10-20° (it can be any value among 10°, 11°, 12°, 13°, 14°, 15°, 16°, 17°, 18°, 19° and 20° or a range corresponding to any two values). By directly processing a single-sided expansion nozzle on the nozzle body, the processing cost is low. At the same time, the nozzle accelerates the airflow to supersonic speed without increasing the weight. When the supersonic airflow is blown into the nozzle end inside the nozzle body, it can act on the separation flow field (the separation boundary layer with a reverse pressure gradient in the nozzle body is directly blown away). The ability to suppress flow separation in the nozzle body is higher than that of passively ejecting on the outside of the nozzle body. In addition, the supersonic airflow can also increase the thrust and specific impulse. The angle between the axis of the throat and the axis of the nozzle body is set to 10-20°, so that the exhaust speed of the single-sided expansion nozzle can be converted into thrust to the greatest extent.

[0050] In the above technical solution, the distance between the gas outlet end of the unilateral expansion nozzle 1112 and the nozzle end of the nozzle body 1111 is L, wherein the maximum value of L is 0.1Lp;

[0051] When the nozzle body 1111 is a conical nozzle, the minimum value of L is 0 (in this case, the value of L can be any value among 0.01Lp, 0.02Lp, 0.03Lp, 0.04Lp, 0.05Lp, 0.06Lp, 0.07Lp, 0.08Lp, 0.09Lp and 0.1Lp, or a range corresponding to any two values);

[0052] When the nozzle body 1111 is a bell-shaped nozzle, the minimum value of L is 0.01Lp (in this case, the value of L can be any value among 0.01Lp, 0.02Lp, 0.03Lp, 0.04Lp, 0.05Lp, 0.06Lp, 0.07Lp, 0.08Lp, 0.09Lp and 0.1Lp, or a range corresponding to any two values);

[0053] Among them, Lp is the length of the expansion section of the nozzle body 1111. This is because the outlet pressure of the single-sided expansion nozzle is smaller when it is closer to the nozzle end of the nozzle body. This is conducive to the single-sided expansion nozzle to obtain a larger jet velocity, so L cannot be too large; and in order to facilitate the processing of the single-sided expansion nozzle, for the conical nozzle, the minimum value of L is 0, and for the bell-shaped nozzle, the minimum value of L is 0.01Lp.

[0054] In the above technical solution, the cross-section of the unilateral expansion nozzle 1112 perpendicular to the axis at the throat 11121 is a rectangle, which helps to simplify the nozzle processing process.

[0055] The height of the throat 11121 described in the above technical solution is Dc, and Dc is 3-10mm (it can be any value among 3mm, 4mm, 5mm, 6mm, 7mm, 8mm, 9mm and 10mm or the corresponding range between any two values), which can also ensure the reliability of the operation of the unilateral expansion nozzle (the width of the throat in the nozzle annular direction in this embodiment can be set as needed, and is not specifically limited here).

[0056] In the above technical solution, the air inlet end and the air outlet end of the unilateral expansion nozzle 1112 are expanded toward the nozzle end of the nozzle body 1111, which makes its processing easier, and the angle between the jet ejected by the unilateral expansion nozzle and the axis of the nozzle body is smaller, and a higher jet velocity can be further obtained.

[0057] like Figure 3 As shown, in the above technical solution, the width of the outlet end of the unilateral expansion nozzle 1112 is expanded to Do, and the expansion ratio of the outlet end of the unilateral expansion nozzle 1112 is ε 1 , ε 1 =Do / Dc,ε 1 The value of is 3-6 (it can be any value among 3, 4, 5 and 6 or the corresponding range between any two values);

[0058] And / or the width of the inlet end of the unilateral expansion nozzle 1112 is expanded to Di, and the contraction ratio of the inlet end of the unilateral expansion nozzle 1112 is ε 2 , ε 2 =Di / Dc,ε 2 The value of is 1.5-3 (it can be any value among 1.5, 2, 2.5 and 3 or the corresponding range between any two values), which is helpful to ensure the reliable operation of the single-sided expansion nozzle at a higher efficiency.

[0059] like Figure 3 As shown, the angle between the expansion surface of the outlet end of the unilateral expansion nozzle 1112 and the axis of the throat 11121 in the above technical solution is θ 2 , where θ 2 >θ 1 ;

[0060] and / or the angle between the expansion surface of the inlet end of the unilateral expansion nozzle 1112 and the axis of the throat 11121 is θ 3 , where θ 3 <90°-θ 1 The jet ejected from the outlet end of the unilateral expansion nozzle is nearly parallel to the axial direction of the unilateral expansion nozzle (i.e., the axis of the throat), and the expansion at both ends of the throat is relatively gentle, thereby avoiding local jet turbulence.

[0061] like Figure 4As shown, in the above technical solution, the side of the throat 11121 close to the nozzle body 1111 nozzle end is chamfered, and the radius of the chamfered part is R, where Dc < R < 2Dc. This enables a smooth transition of the flared surfaces at the gas outlet end and the gas inlet end of the single-sided expansion nozzle 1112 at the throat, further avoiding local air flow turbulence.

[0062] In this embodiment, except for the side close to the nozzle body nozzle end which is a curved surface, the other three sides of the single-sided expansion nozzle are flat surfaces.

[0063] In this embodiment, the steps of machining the single-sided expansion nozzle on the side wall of the nozzle body are mainly divided into three steps:

[0064] Step 1: Mill a rectangular channel on the wall surface at the nozzle end of the nozzle body, with an angle θ with the axis of the nozzle body. 1 The width of the rectangular channel is Dc (as Figure 2 shown);

[0065] Step 2: Use a milling cutter to cut and flare the sides of the rectangular channel close to the nozzle end of the nozzle body at both ends (as Figure 3 shown);

[0066] Step 3: Chamfer the throat (as Figure 4 shown).

[0067] Embodiment 2

[0068] As Figure 5 shown, this embodiment provides a nozzle assembly, including a collar 112 and the nozzle 111 as described in Embodiment 1. The collar 112 is coaxially sleeved on the outer wall of the nozzle body 1111 nozzle end. The collar 112 and the nozzle body 1111 jointly enclose an annular gas collection chamber 113. An air inlet 1121 communicating with the inside of the gas collection chamber 113 is provided on the collar 112. A plurality of single-sided expansion nozzles 1112 are provided and are all located in the gas collection chamber 113. The plurality of single-sided expansion nozzles 1112 are circumferentially spaced apart, and the distance between adjacent two single-sided expansion nozzles 1112 does not exceed 0.5 cm (any value or the corresponding range between any two values among 0.1 cm, 0.2 cm, 0.3 cm, 0.4 cm, and 0.5 cm). In this way, a large number of uniformly distributed high-speed jets are formed circumferentially and spaced apart on the inner side of the nozzle body nozzle end by the plurality of single-sided expansion nozzles, which can improve the uniformity of the jet force distribution at the nozzle end of the nozzle and inhibit the backflow of the atmosphere into the nozzle body (inhibit jet separation). In addition, the distance between adjacent two single-sided expansion nozzles does not exceed 0.5 cm, which can prevent the nozzle from experiencing lateral loads.

[0069] Embodiment 3

[0070] As Figure 6 As shown, this embodiment provides a rocket engine, including a combustion chamber 12 and a nozzle assembly 11 as described in Example 2, wherein the combustion chamber 12 has a first oxidant inlet 121, a first fuel inlet 122 and an injection port 123, and the end of the nozzle body 1111 away from its nozzle end is connected to the injection port 123 of the combustion chamber 12, and a single-sided expansion nozzle is arranged at an annular interval at the nozzle end of the nozzle body, so that the rocket engine can suppress flow separation while having higher thrust and specific impulse.

[0071] Example 4

[0072] like Figure 7 As shown, this embodiment provides a rocket engine assembly, including a turbine 2, an oxidizer pump 3, a fuel pump 4 and a rocket engine 1 as described in Example 3, the discharge port of the oxidizer pump 3 is connected to the first oxidizer inlet 121, the discharge port of the fuel pump 4 is connected to the first fuel inlet 122, the power input ends of the oxidizer pump 3 and the fuel pump 4 are both transmission-connected to the driving end of the turbine 2, the turbine 2 has an air inlet and an exhaust gas outlet, the exhaust gas outlet of the turbine 2 is connected to the air inlet 1121, the air inlet of the turbine 2 is used to pass airflow for driving, so that the turbine can provide driving force for the oxidizer pump and the fuel pump, and at the same time, the exhaust gas of the turbine is discharged into the air collecting chamber, so that the exhaust gas of the turbine is sprayed into the nozzle body through the unilateral expansion nozzle to form a high-speed jet, thereby improving the thrust and specific impulse of the entire engine.

[0073] In this embodiment, the driving shaft of the turbine, the pump shaft of the oxidant pump and the pump shaft of the fuel pump are coaxially distributed and connected to each other, so that the turbine can drive the oxidant pump and the fuel pump to operate simultaneously.

[0074] In the above technical solution, the combustion chamber 12 is also provided with an air distribution port 124, and the air distribution port 124 is connected to the air inlet of the turbine 2, so that part of the airflow generated during combustion in the combustion chamber can be introduced into the turbine to drive the turbine to operate, which makes its structure simple and has good operating reliability.

[0075] Example 5

[0076] like Figure 8As shown, the same as Example 4, the difference is that the above technical solution also includes a gas generator 5, the gas generator 5 has a second oxidant inlet 51, a second fuel inlet 52 and an exhaust port 53, the second oxidant inlet 51 is connected to the discharge port of the oxidant pump 3, the second fuel inlet 52 is connected to the discharge port of the fuel pump 4, and the exhaust port 53 is connected to the air inlet of the turbine 2. Through the independent arrangement of the gas generator, part of the oxidant and fuel delivered by the oxidant pump and the fuel pump are burned to generate fuel gas to drive the turbine to operate.

[0077] Example 6

[0078] like Fig. 9 As shown, the same as Example 4, the difference is that the above technical solution also includes a heat exchange jacket 6, which is arranged in the interlayer in the wall of the combustion chamber 12 and / or the nozzle body 1111 (the combustion chamber and the nozzle body are prevented from having an interlayer cavity for cooling, which is the prior art in the field and will not be repeated here), and the heat exchange jacket 6 has a third fuel inlet 61 and an external discharge port 62, the third fuel inlet 61 is connected to the discharge port of the fuel pump 4, and the external discharge port 62 is connected to the air inlet of the turbine 2, so that the distributed fuel can be introduced into the heat exchange jacket for heating and gasification, so as to serve as the airflow to drive the turbine to operate, and at the same time, the gasified fuel can be burned through the unilateral expansion nozzle to the injection port of the nozzle body to further increase the thrust and specific impulse of the rocket engine.

[0079] Example 7

[0080] like Fig.10 and Fig.11 As shown, the same as Example 4, Example 5 or Example 6, the difference is that in the above technical solution, two turbines 2 are provided, the fuel pump 4 and the oxidant pump 3 are each driven by one of the turbines 2, and the two turbines 2 are connected in series or in parallel (series or parallel means that the air inlets and exhaust gas outlets of the two turbines are connected in series or in parallel), so that the fuel pump and the oxidant pump are respectively driven by independent turbines, and the reliability of their operation is better.

[0081] Among them, Fig.11 As shown, the two turbines are connected in parallel, which means that the air inlets of the two turbines converge, and the exhaust outlets of the two turbines also converge. At this time, the air flows pass through the two turbines respectively and then enter the air collecting chamber, as shown in FIG. Fig.10 As shown, the two turbines are connected in series, which means that the air inlet of one turbine is connected to the exhaust outlet of the other turbine. At this time, the air flow passes through the two turbines in sequence and then enters the air collecting chamber (wherein, Fig.10 and Fig.11The embodiment 4 is taken as an example, and turbines are added to be connected in parallel and in series. The principle of adding turbines in parallel and in series in the embodiments 5 and 6 is the same as Fig.10 and Fig.11 are consistent, so will not be elaborated here).

[0082] The above description is only a preferred embodiment of the present invention and does not limit the present invention in any form. Any ordinary technician in the industry can smoothly implement the present invention as shown in the drawings and described above. However, any equivalent changes, modifications and evolutions made by technicians familiar with the profession without departing from the scope of the technical solution of the present invention using the technical content disclosed above are all equivalent embodiments of the present invention. At the same time, any equivalent changes, modifications and evolutions made to the above embodiments based on the essential technology of the present invention are still within the protection scope of the technical solution of the present invention.

Claims

1. A nozzle, characterized in that: The invention comprises a nozzle body (1111), wherein a unilateral expansion nozzle (1112) is obliquely arranged on the side wall of the nozzle end of the nozzle body (1111) and penetrates the nozzle body (1111), an air inlet end of the unilateral expansion nozzle (1112) is located on the outer wall of the nozzle body (1111), an air outlet end of the unilateral expansion nozzle (1112) is located on the inner wall of the nozzle body (1111), and an air outlet end of the unilateral expansion nozzle (1112) is located on the inner wall of the nozzle body (1111). The gas end faces the nozzle end of the nozzle body (1111), the gas inlet end of the unilateral expansion nozzle (1112) is in a contracted shape, the gas outlet end of the unilateral expansion nozzle (1112) is in a flared shape, and the middle part of the unilateral expansion nozzle (1112) forms a throat (11121), and the angle between the axis of the throat (11121) and the axis of the nozzle body (1111) is θ1, and θ1 is 10-20°.

2. The nozzle according to claim 1, characterized in that The distance between the gas outlet end of the unilateral expansion nozzle (1112) and the nozzle end of the nozzle body (1111) is L, wherein the maximum value of L is 0.1Lp; When the nozzle body (1111) is a conical nozzle, the minimum value of L is 0; When the nozzle body (1111) is a bell-shaped nozzle, the minimum value of L is 0.01Lp; Wherein, Lp is the length of the expansion section of the nozzle body (1111).

3. The nozzle according to claim 1, characterized in that The cross section of the unilateral expansion nozzle (1112) perpendicular to the axis at the throat (11121) is rectangular; the height of the throat (11121) is Dc, and Dc is 3-10 mm.

4. The nozzle according to claim 3, characterized in that The air inlet end of the unilateral expansion nozzle (1112) is contracted toward the nozzle end of the nozzle body (1111), and the air outlet end of the unilateral expansion nozzle (1112) is expanded toward the nozzle end of the nozzle body (1111), and at least one of the following conditions ae must be met when the unilateral expansion nozzle (1112) is contracted or expanded: a: the width of the gas outlet end of the unilateral expansion nozzle (1112) is expanded to Do, the expansion ratio of the gas outlet end of the unilateral expansion nozzle (1112) is ε1, ε1=Do / Dc, and the value of ε1 is 3-6; b: the width of the air inlet end of the unilateral expansion nozzle (1112) is Di, the contraction ratio of the air inlet end of the unilateral expansion nozzle (1112) is ε2, ε2=Di / Dc, and the value of ε2 is 1.5-3; c: the included angle between the expanded surface of the outlet end of the unilateral expansion nozzle (1112) and the center line of the throat (11121) is θ2, where θ2>θ1; d: the angle between the end surface of the air inlet end of the unilateral expansion nozzle (1112) and the axis at the throat (11121) is θ3, where θ3<90°-θ1; e: The throat (11121) is rounded on one side close to the nozzle end of the nozzle body (1111), and the radius of the rounded corner is R, Dc <R<2Dc。 5. A nozzle assembly, characterized in that: It comprises a ring sleeve (112) and a nozzle (111) as claimed in any one of claims 1 to 4, wherein the ring sleeve (112) is coaxially sleeved on the outer wall of the nozzle end of the nozzle body (1111), the ring sleeve (112) and the nozzle body (1111) together enclose an annular gas collecting chamber (113), and the ring sleeve (112) is provided with an air inlet (1121) connected to the gas collecting chamber (113), and a plurality of the single-sided expansion nozzles (1112) are provided and are all located in the gas collecting chamber (113), and the plurality of the single-sided expansion nozzles (1112) are distributed circumferentially at intervals.

6. A rocket engine, characterized in that: The invention comprises a combustion chamber (12) and a nozzle assembly (11) as claimed in claim 5, wherein the combustion chamber (12) has a first oxidant inlet (121), a first fuel inlet (122) and an injection port (123), and an end of the nozzle body (1111) away from the nozzle end thereof is connected to the injection port (123) of the combustion chamber (12).

7. A rocket engine assembly, characterized in that: It comprises a turbine (2), an oxidizer pump (3), a fuel pump (4) and a rocket engine (1) as claimed in claim 6, wherein the discharge port of the oxidizer pump (3) is connected to the first oxidizer inlet (121), the discharge port of the fuel pump (4) is connected to the first fuel inlet (122), the power input ends of the oxidizer pump (3) and the fuel pump (4) are both drivingly connected to the driving end of the turbine (2), the turbine (2) has an air inlet and an exhaust gas outlet, the exhaust gas outlet of the turbine (2) is connected to the air inlet (1121), and the air inlet of the turbine (2) is used to allow air flow to be introduced for driving.

8. The rocket engine assembly according to claim 7, characterized in that: The combustion chamber (12) is also provided with an air splitting port (124), and the air splitting port (124) is in communication with an air inlet of the turbine (2).

9. The rocket engine assembly according to claim 7, characterized in that: The invention also comprises a gas generator (5), wherein the gas generator (5) has a second oxidant inlet (51), a second fuel inlet (52) and an exhaust port (53), wherein the second oxidant inlet (51) is connected to the discharge port of the oxidant pump (3), the second fuel inlet (52) is connected to the discharge port of the fuel pump (4), and the exhaust port (53) is connected to the air inlet of the turbine (2).

10. The rocket engine assembly according to claim 7, characterized in that: It also comprises a heat exchange jacket (6), the heat exchange jacket (6) being arranged in a layer within the wall surface of the combustion chamber (12) and / or the nozzle body (1111), the heat exchange jacket (6) having a third fuel inlet (61) and an external discharge port (62), the third fuel inlet (61) being connected to the discharge port of the fuel pump (4), and the external discharge port (62) being connected to the air inlet of the turbine (2).

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