Engine suitable for extremely cold environment

Abstract

The utility model discloses an engine suitable for an extremely cold environment, which relates to the technical field of rocket engines and comprises a first-stage rocket engine, an interstage separation device and a second-stage rocket engine. The first-stage rocket engine comprises a second combustion chamber, one end of the second combustion chamber is provided with an ignition device, and the other end of the second combustion chamber is connected with a second spray pipe; the interstage separation device comprises an interstage containing cavity, a plurality of gas permeation holes are formed in the interstage containing cavity, and the first-stage rocket engine and the second-stage rocket engine are connected through an elastic clamping base. The two-stage rocket engine comprises a fuel storage tank, an oxidizing agent storage tank, an electric pump, an injector, a first combustion chamber and a first spray pipe. Local materials are used in an extremely cold environment, the raw materials are simple in source and low in cost, and the burden of operators is relieved; and moreover, the solid rocket engine and the gel rocket engine are subjected to optimized layout, the advantage combination of the solid rocket engine and the gel rocket engine is simple, reliable, flexible and adjustable, and diversified tasks can be executed under the extremely cold condition.

Description

Technical Field

[0001] This utility model relates to the field of rocket engine technology, and in particular to an engine suitable for extremely cold environments. Background Technology

[0002] Against the backdrop of global climate change, extreme weather events are becoming increasingly frequent, especially in high-latitude or high-altitude regions where extreme low temperatures have become the norm, often dropping below -30°C and, in extreme cases, below -50°C. Such environments greatly increase the difficulty of military operations and mobile maneuvers. With the increasing frequency of extreme weather events globally, the growing complexity of international security situations, and the increasing number of missions operating in extremely cold environments, finding a balance between cost-effectiveness and operational efficiency in extreme conditions has become a pressing issue in the military field.

[0003] Low temperatures can lead to fuel freezing, mechanical equipment malfunctions, and increased maneuverability and reduced speed due to snow and ice, significantly decreasing combat efficiency. This necessitates more portable or simplified propulsion devices to alleviate the burden. Against this backdrop, it is necessary to consider the unique characteristics of combat operations in extremely cold environments and the operational limitations of military equipment under extreme conditions. The aim is to provide the PLA with a way to enhance combat capabilities in harsh environments while maintaining high performance and low cost, thereby extending endurance and effectively reducing operational costs.

[0004] The existing engine structure is incompatible with cryogenic solid propellants and gel propellants, making it unsuitable for extremely cold environments and failing to achieve the expected results.

[0005] Therefore, this invention provides an engine suitable for extremely cold environments to solve the problems existing in the prior art. Utility Model Content

[0006] To achieve the above objectives, the present invention provides the following solution: The present invention provides an engine suitable for extremely cold environments, including a first-stage rocket engine, an interstage separation device, and a second-stage rocket engine;

[0007] The first-stage rocket engine includes a second combustion chamber, one end of which is equipped with an ignition device, and the other end of which is connected to a second nozzle. The first-stage rocket engine uses aluminum ice as a propellant, and the second-stage rocket engine uses hydrogen peroxide gel as a propellant.

[0008] The interstage separation device includes an interstage cavity with several gas permeation holes, and the first-stage rocket engine and the second-stage rocket engine are connected by a flexible bracket.

[0009] The second-stage rocket engine includes a fuel tank, an oxidizer tank at one end of the fuel tank, and an electric pump at the other end of the fuel tank. The oxidizer tank is connected to the electric pump through an oxidizer pipe, and the electric pump is connected to the fuel tank through a fuel pipe. An injector is located at the end of the electric pump away from the fuel tank, and the injector is connected to a first combustion chamber. The first combustion chamber is connected to a first nozzle.

[0010] Preferably, the second combustion chamber is provided with a front end cap and a rear end cap at both ends, the front end cap being located at one end of the ignition device and the rear end cap being located at one end of the second nozzle.

[0011] Preferably, the first-stage rocket engine is provided with a tail fin on its outer side.

[0012] Preferably, the second-stage rocket engine is provided with wings on its outer side.

[0013] Preferably, the second-stage rocket engine is equipped with a warhead at its end, the warhead comprising a warhead and a guidance system.

[0014] Preferably, the front end cap is ellipsoidal, and the rear end cap and the second nozzle are sealed with an O-ring.

[0015] This utility model discloses the following technical effects: it can use local materials in extremely cold environments, the raw material sources are simple and the cost is low, reducing the burden on operators and protecting the environment at the same time; and it optimizes the layout of solid rocket motors and gel rocket motors, the combination of the advantages of solid motors and gel motors is simple, reliable, flexible and adjustable, and can perform a variety of tasks under extremely cold conditions. Attached Figure Description

[0016] The accompanying drawings, which form part of this application, are used to provide a further understanding of this application. The illustrative embodiments and descriptions of this application are used to explain this application and do not constitute an undue limitation of this application. In the drawings:

[0017] Figure 1 This is a schematic diagram of the overall structure of the engine of this utility model;

[0018] Figure 2 This is a schematic diagram of the structure of the first-stage rocket engine of this utility model;

[0019] Figure 3 This is a schematic diagram of the structure of the second-stage rocket engine of this utility model;

[0020] In the diagram: 1. Warhead; 2. Guidance system; 3. Oxidizer tank; 4. Fuel tank; 5. Wings; 6. Electric pump; 7. First combustion chamber; 8. Gas infiltration port; 9. Second combustion chamber; 10. Tail fins; 11. Injector; 12. First nozzle; 13. Flexible mounting bracket; 14. Interstage cavity; 15. Ignition device; 16. Second nozzle; 17. Front end cap; 18. Rear end cap; 19. Oxidizer conduit; 20. Fuel conduit. Detailed Implementation

[0021] The technical solutions of the present utility model will be clearly and completely described below with reference to the accompanying drawings of the embodiments. Obviously, the described embodiments are only some embodiments of the present utility model, and not all embodiments. Based on the embodiments of the present utility model, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the protection scope of the present utility model.

[0022] To make the above-mentioned objectives, features and advantages of this utility model more apparent and understandable, the utility model will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0023] Reference Figures 1-3 As shown, this embodiment provides an engine suitable for extremely cold environments, including a first-stage rocket engine, an interstage separation device, and a second-stage rocket engine;

[0024] The first-stage rocket engine includes a second combustion chamber 9, with an ignition device 15 installed at one end of the second combustion chamber 9 and a second nozzle 16 connected to the other end of the second combustion chamber 9;

[0025] The interstage separation device includes an interstage cavity 14, which is provided with several gas infiltration holes 8. The first-stage rocket engine and the second-stage rocket engine are connected by an elastic bracket 13.

[0026] The second-stage rocket engine includes a fuel tank 4, an oxidizer tank 3 at one end of the fuel tank 4, and an electric pump 6 at the other end of the fuel tank 4. The oxidizer tank 3 is connected to the electric pump 6 through an oxidizer pipe 19. The electric pump 6 is connected to the fuel tank 4 through a fuel pipe 20. An injector 11 is located at the end of the electric pump 6 away from the fuel tank 4. The injector 11 is connected to a first combustion chamber 7. The first combustion chamber 7 is connected to a first nozzle 12.

[0027] The first-stage rocket engine uses aluminum ice as the propellant, while the second-stage rocket engine uses hydrogen peroxide gel as the propellant. This combination of aluminum ice and hydrogen peroxide gel propellants leverages the advantages of each, fully utilizing the abundant raw materials and high specific impulse of aluminum ice, while cleverly utilizing the controllability advantages of hydrogen peroxide gel. While ensuring adaptability to extremely cold environments, it features clean, low-carbon, high-energy, and excellent combustion characteristics. In terms of preparation and use: it turns the disadvantages of extremely cold environments into advantages, using locally sourced, simple, and low-cost raw materials, reducing the burden on operators, and also protecting the environment. In terms of engine design: a multi-stage rocket engine is designed based on the characteristics of both propellants, optimizing the layout of solid rocket engines and gel rocket engines. The first-stage engine uses high-energy-density aluminum ice propellant to provide a powerful boost in the initial stage of missile launch, rapidly accelerating the second-stage engine to a specific Mach number. The second-stage engine uses hydrogen peroxide gel propellant to provide adjustable thrust, enabling more precise trajectory adjustments or attitude control, thereby improving strike accuracy. The combination of solid propellant and gel propellant engines offers advantages such as simplicity, reliability, flexibility, and adjustability, enabling diverse tasks to be performed under extremely cold conditions.

[0028] Further optimizing the design, the second combustion chamber 9 is equipped with a front end cap 17 and a rear end cap 18 at both ends. The front end cap 17 is located at one end of the ignition device 15, and the rear end cap 18 is located at one end of the second nozzle 16. The front end cap 17, the second combustion chamber 9, and the rear end cap 18 are all made of ultra-high strength alloy steel 30GrMnSiA, which has good mechanical properties, weldability, and high-temperature performance, and is suitable for the operating conditions of a first-stage engine with high thrust and short operating time.

[0029] The design was further optimized by adding a tail fin 10 to the outside of the first-stage rocket engine. The tail fin 10 ensures stability during flight.

[0030] The design was further optimized, with fins 5 installed on the outside of the second-stage rocket engine.

[0031] The design was further optimized so that a warhead was installed at the end of the second-stage rocket engine. The warhead includes a warhead 1 and a guidance system 2.

[0032] The design was further optimized by making the front end cap 17 ellipsoidal and using an O-ring to seal the rear end cap 18 and the second nozzle 16.

[0033] Work process:

[0034] (1) Propellant preparation and loading

[0035] The preparation method for aluminum ice propellant is as follows: ice water is collected on-site under extremely cold conditions, mixed with nano-aluminum powder, stirred and frozen into shape, and then added to the second combustion chamber 9 of the first-stage engine. The preparation method for hydrogen peroxide gel propellant is as follows: a certain amount of sodium alginate is added to an H2O2 solution and stirred thoroughly until completely dissolved. Calcium ion compounds are added to the mixture, and mechanically stirred until homogeneous, resulting in an H2O2 sol. This sol is placed in an environment of approximately 263K for 1-6 hours for initial gelation. Finally, the sample is stored at a low temperature of 243K for 48 hours to obtain a fully cured HTP gel, which is then filled into the oxidizer tank 3 of the second-stage engine for long-term stable storage. Boron powder can be selected as the fuel for the hydrogen peroxide gel propellant, and it is filled into the fuel tank 4.

[0036] (2) Launch system activation and preparation

[0037] After the two-stage engine is assembled and placed on the launch pad, the power system is immediately activated once the target information is confirmed, providing a stable power supply to the entire launch system.

[0038] (2) First-stage rocket engine booster

[0039] After the launch command is officially issued, the ignition device 15 of the first-stage rocket engine is detonated, triggering the combustion of the propellant and generating tremendous thrust to rapidly lift the surface-to-air missile off the launch pad. During this phase, the first-stage rocket engine's role is to provide the main missile with the necessary initial velocity and altitude to reach the predetermined Mach number.

[0040] (3) Interstage separation mechanism activated

[0041] When the missile reaches the predetermined Mach number, the high-pressure gas accumulated in the interstage cavity 14 activates the interstage separation device. Under the action of the high-pressure gas, the first-stage rocket engine successfully separates from the missile body, creating conditions for the subsequent start-up of the second-stage rocket engine and further acceleration of the missile.

[0042] (4) Second stage rocket engine activation

[0043] With the separation of the first-stage rocket engine, the electric pump 6 and igniter (using an electric detonator and solid propellant ignition) of the second-stage rocket engine immediately activate, marking the missile's entry into the second stage of propulsion. The electric pump 6 draws oxidizer from the oxidizer tank 3 and fuel from the fuel tank 4, delivering these two key components through the oxidizer line 19 and fuel line 20 to the injector 11 located at the front of the combustion chamber. The injector 11 atomizes the fuel and injects it into the combustion chamber for efficient mixing. The igniter then ignites the oxidizer and fuel, resulting in intense combustion that continues to provide thrust to the missile.

[0044] (5) Precision strike adjustment

[0045] Operators precisely adjust the flow rate of electric pump 6 according to the target's dynamic state, thereby adjusting the thrust, optimizing the missile's flight trajectory and speed, and ensuring high-precision air strikes.

[0046] In the description of this utility model, it should be understood that the terms "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this utility model and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this utility model.

[0047] The embodiments described above are merely preferred embodiments of the present utility model and are not intended to limit the scope of the present utility model. Various modifications and improvements made to the technical solutions of the present utility model by those skilled in the art without departing from the spirit of the present utility model should fall within the protection scope defined by the claims of the present utility model.

Claims

1. An engine suitable for extremely cold environments, characterized in that: This includes the first-stage rocket engine, the interstage separation device, and the second-stage rocket engine; The first-stage rocket engine includes a second combustion chamber (9), one end of which is provided with an ignition device (15), and the other end of which is connected to a second nozzle (16); The interstage separation device includes an interstage cavity (14), which is provided with a plurality of gas infiltration holes (8). The first-stage rocket engine and the second-stage rocket engine are connected by an elastic bracket (13). The second-stage rocket engine includes a fuel tank (4), an oxidizer tank (3) at one end of the fuel tank (4), and an electric pump (6) at the other end of the fuel tank (4). The oxidizer tank (3) is connected to the electric pump (6) through an oxidizer pipe (19). The electric pump (6) is connected to the fuel tank (4) through a fuel pipe (20). An injector (11) is provided at the end of the electric pump (6) away from the fuel tank (4). The injector (11) is connected to a first combustion chamber (7), and the first combustion chamber (7) is connected to a first nozzle (12).

2. The engine suitable for extremely cold environments according to claim 1, characterized in that: The second combustion chamber (9) is provided with a front end cap (17) and a rear end cap (18) at both ends. The front end cap (17) is located at one end of the ignition device (15), and the rear end cap (18) is located at one end of the second nozzle (16).

3. The engine suitable for extremely cold environments according to claim 1, characterized in that: The first-stage rocket engine is equipped with a tail fin (10) on its outer side.

4. The engine suitable for extremely cold environments according to claim 1, characterized in that: The second-stage rocket engine is equipped with wings (5) on its outer side.

5. The engine suitable for extremely cold environments according to claim 1, characterized in that: The second-stage rocket engine is equipped with a warhead at its end, which includes a warhead (1) and a guidance system (2).

6. The engine suitable for extremely cold environments according to claim 2, characterized in that: The front end cap (17) is ellipsoidal, and the rear end cap (18) and the second nozzle (16) are sealed with an O-ring.

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