Integrated gas-generator cycle liquid rocket engine

Abstract

The present invention discloses an integrated staged combustion cycle liquid rocket engine, which adopts a two-stage combustion mode and includes: a combustion chamber at the upper end, a staged combustion chamber at the lower end, and a turbopump rotor located between the combustion chamber and the staged combustion chamber; the turbopump rotor is an integrated structure printed by 3D printing, which includes a turbine and an integrated pump, and the turbine is coaxially arranged on the outer periphery of the integrated pump. The present invention abandons the independent turbopump and gas generator, integrates the turbine and the double pump impeller into the engine, simplifies the system architecture of the engine, no longer requires a cumbersome assembly process, and achieves a significant weight reduction effect; in the present invention, the integrated pump and the turbine are coaxially arranged, the traditional transmission shaft is abandoned, and it is manufactured by 3D printing technology, integrating complex structures into one body, reducing the overall weight and the manufacturing cost.

Description

Technical Field

[0001] The present invention relates to an integrated staged combustion cycle liquid rocket engine. Background Art

[0002] At present, the pump-fed engines used in launch vehicles and space vehicles usually consist of an independent turbopump and a staged combustion and afterburning chamber, with a complex system and high manufacturing cost.

[0003] For example, the invention patent with the publication number CN114060170A relates to an open staged combustion and gas extraction cycle integrated staged combustion cycle liquid rocket engine, which consists of a set of turbopumps, a staged combustion and afterburning chamber, and corresponding control valves, etc. The staged combustion and afterburning chamber adopts a two-stage combustion mode of an upper chamber and a lower chamber. Part of the fuel and all the oxidizer organize rich combustion in the upper chamber of the afterburning chamber, forming a rich combustion zone with a relatively low temperature. The generated rich combustion gas is then introduced into the lower chamber of the afterburning chamber to perform afterburning combustion with most of the fuel introduced after the oxygen pump.

[0004] The above-mentioned invention patent still retains an independent turbopump. Therefore, this open cycle integrated staged combustion cycle liquid rocket engine not only has a reduction in engine efficiency due to the external exhaust of the gas driving the turbine, but also has a complex system and high manufacturing cost.

[0005] The present invention proposes an integrated staged combustion cycle liquid rocket engine without an independent turbopump, completely abandoning the gas generator, and integrating a turbine and a double-pump impeller in the combustion chamber. This new type of integrated staged combustion cycle liquid rocket engine has the characteristics of a simple and reliable system and extremely low manufacturing cost. Summary of the Invention

[0006] The present invention aims to provide a technical solution to solve the above problems in order to overcome the above deficiencies.

[0007] An integrated staged combustion cycle liquid rocket engine, adopting a two-stage combustion mode, includes: a combustion chamber at the upper end, an afterburning chamber at the lower end, and a turbopump rotor located between the combustion chamber and the afterburning chamber;

[0008] The turbopump rotor is an integrated structure printed by 3D printing, which includes a turbine and an integrated pump. The turbine is coaxially arranged on the outer periphery of the integrated pump, and the outer side of the turbine is rotationally connected to the inner wall of the bottom of the combustion chamber through a first bearing ring;

[0009] The integrated pump consists of an oxygen impeller arranged on the central axis, a fuel impeller sleeved on the outer periphery of the oxygen impeller, and a pump casing sleeved on the outer periphery of the fuel impeller;

[0010] A plurality of oxygen blades on the outer periphery of the oxygen impeller are connected to the fuel impeller, and an oxidant flow channel is formed between any two oxygen blades. The oxidant flow channel communicates with the combustion chamber through a plurality of oxidant injection pipes, and the oxidant flow channel communicates with the afterburning chamber through a plurality of afterburning nozzles;

[0011] A plurality of fuel blades on the outer periphery of the fuel impeller are connected to the pump housing, and a fuel flow channel is formed between any two fuel blades. A plurality of fuel nozzles communicating with the combustion chamber are provided at the lower end of the pump housing, and a plurality of cooling nozzles for cooling the turbine are provided at the end of the pump housing, and the ejection end of each cooling nozzle is located between any two turbine blades;

[0012] Preferably, a plurality of guide vanes are uniformly arranged on the inner peripheral wall of the upper end of the afterburning chamber, and a fixing frame is connected at the concentric position of the plurality of guide vanes;

[0013] Preferably, the central shaft of the oxygen impeller is rotatably installed at the center of the fixing frame, and a second bearing ring is further installed between the center of the oxygen impeller and the fixing frame;

[0014] Preferably, the upper end of the pump housing is rotatably connected to the head of the combustion chamber through a third bearing ring, and the head end of the pump housing passes through the head of the combustion chamber and is rotatably connected to the fuel conduit;

[0015] Preferably, the head end of the fuel impeller passes through the fuel conduit and is rotatably connected to the oxidant conduit.

[0016] Compared with the prior art, the advantages of the present invention are:

[0017] The present invention abandons the independent turbo pump and gas generator, integrates the turbine and the double pump impeller into the engine, simplifies the system architecture of the engine, no longer requires a cumbersome assembly process, and achieves a significant weight reduction effect;

[0018] In the present invention, the integrated pump and the turbine are arranged coaxially, abandoning the traditional transmission between the two by using a transmission shaft, and manufactured by 3D printing technology. The complex structures are integrated into one body, reducing the overall weight and the manufacturing cost.

[0019] The additional aspects and advantages of the present invention will be partially given in the following description, partially become obvious from the following description, or be understood through the practice of the present invention. Description of the Drawings

[0020] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0021] Figure 1 It is a three-dimensional view of the structure of the turbine pump rotor of the present invention.

[0022] Figure 2 It is another three-dimensional view of the turbine pump rotor of the present invention from a different angle.

[0023] Figure 3 It is a schematic structural diagram of the present invention.

[0024] Figure 4 It is Figure 3 The enlarged view within circle A in

[0025] Figure 5 It is a partial cross-sectional view of the afterburning chamber of the present invention.

[0026] Figure 6 It is the front view structural diagram of the turbine pump rotor of the present invention.

[0027] Figure 7 It is Figure 6 The cross-sectional view along line B-B in

[0028] Figure 8 It is Figure 6 The cross-sectional view along line R-R in Specific embodiments

[0029] Next, the technical solutions in the embodiments of the present invention will be described clearly and completely. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of them. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the scope of protection of the present invention.

[0030] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus should not be construed as a limitation of the present invention.

[0031] In addition, in the description of the present invention, unless otherwise clearly defined and limited, the terms "installation", "connection", "connection", etc. should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can also be the communication inside two elements. It can be a wireless connection or a wired connection. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific situations.

[0032] In addition, the technical features involved in different embodiments of the present invention described hereinafter can be combined with each other as long as they do not conflict with each other.

[0033] Please refer to Figures 1 to 8 , in an embodiment of the present invention, an integrated staged combustion cycle liquid rocket engine adopts a two-stage combustion mode, and includes: a combustion chamber 1 at the upper end, a staged combustion chamber 2 at the lower end, and a turbopump rotor 3 located between the combustion chamber 1 and the staged combustion chamber 2. In the present invention, the functions of the turbine 30 pump are integrated and designed in the combustion chamber 1, the layout is optimized, the connections between components are made more compact, and the functions of the turbine 30 pump and the burner are still retained in the combustion chamber 1, reducing a large number of external pipeline connections, which helps to reduce the weight of the engine and improve its performance and carrying capacity.

[0034] The present invention uses a bipropellant. The fuel can be liquid hydrogen or methane, etc. In order to reduce costs, the present invention can also use kerosene or alcohol as the fuel, and the oxidizer is selected as common liquid oxygen.

[0035] In this embodiment, the present invention uses 3D printing technology to manufacture the integrated structure of the turbopump rotor 3, integrating the turbine 30 and the integrated pump 300 together, reducing the complexity brought by the connection of multiple components in traditional manufacturing, and abandoning the cumbersome assembly process, reducing the connection error between components, and improving the accuracy and reliability of the overall structure.

[0036] In an embodiment of the present invention, the turbine 30 is coaxially arranged on the outer periphery of the integrated pump 300, completely abandoning the traditional technical solution in the prior art that realizes power transmission between the turbine 30 and the pump through a transmission shaft, reducing the number of components of the engine system, making the structure simpler and more compact, and avoiding the risks of system vibration and failure caused by problems such as insufficient installation accuracy, wear, and looseness of the transmission shaft, enhancing the stability and reliability of the entire engine structure.

[0037] In an embodiment of the present invention, the turbopump rotor 3 is integrally formed by 3D printing technology. The integrated pump 300 is composed of three coaxially nested parts: an oxygen impeller 310 located on the central axis, a fuel impeller 320 sleeved on the outer periphery of the oxygen impeller 310, and a pump casing 330 wrapping the fuel impeller 320. The three are precisely connected through 3D printing to form a compact and efficient propellant delivery structure.

[0038] In an embodiment of the present invention, a plurality of oxygen vanes 311 are distributed on the outer periphery of the oxygen impeller 310. The oxygen vanes 311 not only play a role in guiding the flow of the oxidizer, but also connect to the outer fuel impeller 320 to form a stable structural connection. An oxidizer flow channel 340 is formed between any two oxygen vanes 311. Among them, the oxidizer flow channel 340 is communicated with the combustion chamber 1 through a plurality of oxidizer injection pipes 341.

[0039] In an embodiment of the present invention, a plurality of fuel vanes 321 on the outer periphery of the fuel impeller 320 are connected to the pump housing 330, forming a reliable support and connection structure, and a fuel flow passage 350 is formed between any two fuel vanes 321. A plurality of fuel nozzles 351 communicating with the combustion chamber 1 and the fuel flow passage 350 are provided at the lower end of the pump housing 330.

[0040] In an embodiment of the present invention, a plurality of cooling nozzles 352 for cooling the turbine 30 are provided at the end of the pump housing 330. The ejection end of each cooling nozzle 352 is located between any two turbine vanes 31, effectively avoiding a cooling blind area and ensuring that all parts of the turbine 30 and the turbine vanes 31 can be evenly and sufficiently cooled under high-temperature and high-speed operating conditions.

[0041] In an embodiment of the present invention, the oxidant flow passage 340 can also deliver an appropriate amount of oxidant to the afterburning chamber 2 through a plurality of afterburning nozzles 342, providing sufficient oxidant for the afterburning chamber 2.

[0042] In an embodiment of the present invention, the outer ring of the turbine 30 is rotatably connected to the bottom side wall of the combustion chamber 1 through a first bearing ring 400. The rotation of the first bearing ring 400 provides a stable and reliable rotation basis for the turbine pump rotor 3. The first bearing ring 400 can accurately bear the huge radial force generated during the high-speed rotation of the turbine pump rotor 3, ensuring that the turbine pump rotor 3 maintains an accurate rotation axis during operation and effectively reducing vibration and wear.

[0043] As Figures 1 - 5 shown, in an embodiment of the present invention, a plurality of guide vanes 21 are evenly arranged on the inner peripheral wall at the upper end of the afterburning chamber 2. A fixing frame 22 is connected at the concentric position of the plurality of guide vanes 21. The central axis of the oxygen impeller 310 is rotatably installed at the center of the fixing frame 22, and a second bearing ring 500 is also installed between the center of the oxygen impeller 310 and the center of the fixing frame 22. The setting of the fixing frame 22 and the second bearing ring 500 provides a stable support and rotation condition for the central axis of the oxygen impeller 310, ensuring the stability of the turbine pump rotor 3 during high-speed rotation and further ensuring the stability and reliability of oxidant delivery.

[0044] In an embodiment of the present invention, the upper end of the pump housing 330 is rotatably connected to the head of the combustion chamber 1 through a third bearing ring 600, ensuring the stability of the turbo pump rotor 3 during high-speed operation. The third bearing ring 600 effectively bears various acting forces generated during the rotation of the pump housing 330, reducing the friction and wear between the pump housing 330 and the head of the combustion chamber 1, ensuring the smooth rotation of the pump housing 330. Moreover, the head end of the pump housing 330 passes through the head of the combustion chamber 1 and is rotatably connected to the fuel conduit, realizing the smooth transmission of fuel from the fuel conduit to the pump housing 330. It should be emphasized here that the head end of the fuel impeller 320 passes through the fuel conduit and is rotatably connected to the oxidizer conduit. During the high-speed rotation of the fuel impeller 320, it can not only maintain a stable connection with the fuel conduit to ensure fuel input, but also provide an inlet for the oxidizer on its inner circumference.

[0045] The arrangements of the first bearing ring 400, the second bearing ring 500, and the third bearing ring 600 not only improve the working stability and reliability of the turbo pump rotor 3, but also make the rotation of the turbo pump rotor 3 smoother, reduce the energy loss during power transmission, and thus ensure that the turbo pump rotor 3 can operate at a higher speed and in a more stable state.

[0046] Working principle of the present invention:

[0047] When the rocket is in operation, the two propellant tanks of the rocket are pressurized, so that the oxidizer and fuel are respectively transported to the rocket engine through the oxidizer conduit and the fuel conduit. The turbo pump rotor 3 rotates at a high speed to generate a strong centrifugal force, causing the oxidizer to enter the oxidizer flow channel 340 from the inner circumferential annular space of the fuel impeller 320. Under the action of the centrifugal force, the oxidizer quickly diffuses towards the end of the oxygen impeller 310, and the pressure of the oxidizer is significantly increased. A part of the oxidizer is accurately injected into the combustion chamber 1 through the oxidizer injection pipe 341;

[0048] Similarly, under the action of the strong centrifugal force generated by the high-speed rotation of the turbo pump rotor 3, the fuel enters from the annular space between the outer circumference of the fuel impeller 320 and the inner circumference of the pump housing 330. The pressurized and accelerated fuel flows along the fuel flow channel 350. Under the action of the centrifugal force, the fuel is thrown towards the outer edge of the fuel impeller 320, and its pressure increases significantly. A part of the fuel is evenly injected into the combustion chamber 1 through the fuel nozzle 351;

[0049] When the oxidizer and fuel are fully mixed in the combustion chamber 1, the igniter in the combustion chamber 1 ignites the mixed propellant, and the instantaneous combustion forms high-temperature and high-pressure gas. The gas is directed to the turbine 30 and impacts the turbine blades 31 at a high speed. At this time, the kinetic energy of the gas is converted into the mechanical energy of the turbine 30, driving the turbine 30 to rotate at a high speed. The rotation of the turbine 30 drives the integrally connected pump 300, the oxygen impeller 310, and the fuel impeller 320 to work synchronously, forming a power cycle and continuously providing power for the transportation of the propellant.

[0050] Another part of the fuel is ejected through the cooling nozzle 352. This part of the fuel acts as a cooling medium and covers the surface of the turbine blade 31, quickly removing heat through convective heat transfer. After the high-temperature gas in the combustion chamber 1 is mixed with the fuel acting as the cooling medium, it forms a rich gas and passes through the turbine 30 at high speed, and is ejected into the afterburner 2.

[0051] Another part of the oxidizer is transported to the afterburner 2 through the afterburning nozzle 342 under pressure. These oxidizers are fully mixed with the rich gas flowing out of the combustion chamber 1, and a violent combustion reaction occurs again in the afterburner 2, releasing more energy. This not only improves the overall efficiency of the engine but also significantly increases the thrust of the engine, enabling the rocket to obtain more powerful power.

[0052] For those skilled in the art, it is obvious that the present invention is not limited to the details of the above-described exemplary embodiments, and the present invention can be implemented in other specific forms without departing from the spirit or basic characteristics of the present invention. Therefore, from any point of view, the embodiments should be regarded as exemplary and non-limiting. The scope of the present invention is defined by the appended claims rather than the above description. Therefore, all changes falling within the meaning and scope of the equivalent elements of the claims are intended to be embraced within the present invention.

Claims

1. An integrated staged combustion cycle liquid rocket engine, adopting a two-stage combustion mode, comprising: A combustion chamber at the upper end, a afterburning chamber at the lower end, and a turbopump rotor located between the combustion chamber and the afterburning chamber; characterized in that: The turbopump rotor is an integrally structured 3D printed component, which includes a turbine and an integrated pump. The turbine is coaxially arranged on the outer periphery of the integrated pump, and the outer side of the turbine is rotationally connected to the inner wall of the bottom of the combustion chamber through a first bearing ring; The integrated pump is composed of an oxygen impeller arranged on the central axis, a fuel impeller sleeved on the outer periphery of the oxygen impeller, and a pump casing sleeved on the outer periphery of the fuel impeller; Multiple oxygen blades on the outer periphery of the oxygen impeller are connected to the fuel impeller, and an oxidant flow channel is formed between any two oxygen blades. The oxidant flow channel is communicated with the combustion chamber through multiple oxidant injection pipes, and the oxidant flow channel is communicated with the afterburning chamber through multiple afterburning nozzles; Multiple fuel blades on the outer periphery of the fuel impeller are connected to the pump casing, and a fuel flow channel is formed between any two fuel blades. Multiple fuel nozzles communicating with the combustion chamber are arranged at the lower end of the pump casing, and multiple cooling nozzles for cooling the turbine are arranged at the end of the pump casing, and the ejection end of each cooling nozzle is located between any two turbine blades; The central shaft of the oxygen impeller is rotationally installed at the center of the fixed frame, and a second bearing ring is also installed between the center of the oxygen impeller and the fixed frame; The upper end of the pump casing is rotationally connected to the head of the combustion chamber through a third bearing ring, and the head end of the pump casing passes through the head of the combustion chamber and is rotationally connected to the fuel conduit; 2. The integrated topping combustion cycle liquid rocket engine according to claim 1, wherein A plurality of guide vanes are evenly arranged on the inner peripheral wall of the upper end of the afterburning chamber, and a fixed frame is connected at the concentric position of the plurality of guide vanes; 3. The integrated topping cycle liquid rocket engine according to claim 1, characterized in that, The head end of the fuel impeller passes through the fuel conduit and is rotationally connected to the oxidant conduit.

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