A solid rocket motor structure with adjustable radial nozzle position

Abstract

This application discloses a solid rocket motor structure with adjustable radial nozzle position, relating to the field of solid rocket motors. It includes a front combustion chamber shell, a base assembly, and a rear combustion chamber shell connected sequentially. The base assembly is provided with a gas passage connecting the front and rear combustion chamber shells, allowing the gas generated within the front and rear combustion chamber shells to flow through the gas passage. The base assembly is equipped with multiple radial nozzles, one end of which is connected to either the front or rear combustion chamber, and the other end is connected to the outside. This design enables the base assembly to be adjustable within a certain range between the front and rear combustion chamber shells, solving the problem of optimizing the balance between the position of conventional single-chamber orbit control devices and the overall rocket's center of mass.

Description

Technical Field

[0001] This invention relates to the field of solid rocket motors, and more specifically to a solid rocket motor structure with an adjustable radial nozzle position. Background Technology

[0002] To intercept high-altitude, high-speed, and ultra-high-speed incoming targets, air defense missiles need to provide radially controllable thrust for trajectory changes in the thin atmosphere of high altitudes; and the trajectory control radial thrust needs to be matched with the center of mass of the entire missile to maximize its effectiveness.

[0003] Conventional orbit control devices typically employ a combination of a nozzle base and a single combustion chamber. The radial nozzle is located at either the foremost or rearmost end of the orbit control device. The center of gravity of the entire missile is primarily balanced through various onboard equipment (such as the warhead and integrated electronic equipment). This involves numerous related devices, and achieving matching is time-consuming and labor-intensive. Optimal matching design between the missile orbit control device and the center of gravity is challenging, thus hindering the improvement of missile performance. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a solid rocket motor structure with adjustable radial nozzle position, which realizes the adjustable design of the intermediate position within a certain range and solves the problem of optimizing the balance between the position of the conventional single-chamber track control device and the center of mass of the entire rocket.

[0005] The technical solution of this invention is:

[0006] A solid propellant engine structure with adjustable radial nozzle position includes a front combustion chamber housing, a base assembly, and a rear combustion chamber housing connected in sequence. Solid propellant is disposed inside both the front and rear combustion chamber housings. The base assembly is provided with a gas passage connecting the front and rear combustion chamber housings so that the gas generated inside the front and rear combustion chamber housings can flow through the gas passage. The base assembly is provided with multiple radial nozzles, one end of which is connected to the front or rear combustion chamber and the other end is connected to the outside.

[0007] In one specific implementation, the base assembly includes a nozzle base, a front cover plate assembly connected to both ends of the nozzle base, and a rear thermal protection structure. The radial nozzle is located inside the nozzle base, with one end connected to the nozzle base and communicating with the outside, and the other end closed. The rear thermal protection structure is provided with multiple gas inlets, each of which is connected to a radial nozzle.

[0008] In one specific implementation, the gas inlet position of the rear thermal protection structure is provided with an inlet pipe, which is inserted into a radial nozzle, and each gas inlet is connected to a radial nozzle through the inlet pipe.

[0009] In one specific implementation, the rear thermal protection structure is provided with a gas pipeline, the other end of which passes through the front cover assembly to connect the front combustion chamber housing and the rear combustion chamber housing.

[0010] In one specific implementable embodiment, an integrated device is arranged within the nozzle base, and a through-hole is provided on the non-pressure-bearing sidewall of the nozzle base, with the electrical connector of the integrated device installed in the through-hole.

[0011] In one specific implementation, the inner wall of the nozzle base is provided with a front annular step and a rear annular step. The front cover plate assembly abuts against the side of the front annular step away from the rear annular step, and the rear thermal protection structure abuts against the side of the rear annular step away from the front annular step. Pin holes are provided at both ends of the nozzle base. The end of the front combustion chamber housing is inserted into the interior of one end of the nozzle base, and the end of the rear combustion chamber housing is inserted into the interior of the other end of the nozzle base. The exhaust pin passes through the pin hole and connects to the front combustion chamber housing or the rear combustion chamber housing to realize the connection between the front combustion chamber housing and the nozzle base, and the connection between the rear combustion chamber housing and the nozzle base.

[0012] In one specific implementation, the front cover assembly includes a front insulation panel and a front cover panel bonded together by a high-temperature adhesive, with the front insulation panel located on the side of the front cover panel facing away from the integrated device.

[0013] In one specific implementation, the rear thermal protection structure is a protective layer that fully covers the profile of the nozzle base end for heat insulation and ablation prevention; a thermal protection structure is provided outside the gas pipeline; and a thermal protection structure is provided outside the radial nozzle.

[0014] In one specific implementation, the front combustion chamber and the rear combustion chamber adopt an end-burning charge configuration, and simultaneously burn to generate gas.

[0015] In one specific implementation, the meltable igniter is fixed in the gas passage of the base assembly and matches the passage profile.

[0016] The nozzle base, situated between the front and rear combustion chamber shells, is an irregularly shaped structural component. The radially arranged nozzle structure is a non-metallic structure for mounting the radial nozzles. The center of the radial nozzles is used to install the drive control equipment. One angled area of ​​the nozzle serves as a gas passage, while the remaining portion provides installation space for the drive control equipment and energy equipment. The base assembly consists of the nozzle base and front cover plate, a rear insulation structure (including the gas passage), and a sealing structure, collectively forming a closed protection for the servo drive device. The electrical interface of the drive control device can be connected to the cable network through the non-pressure-bearing sidewall. The base assembly is structurally and thermally protected to the front and rear combustion chambers via a drain pin. Ignition is performed using a meltable igniter adapted to the gas passage. Both the front and rear combustion chambers employ an end-burning propellant configuration, with combustion matched to internal ballistic performance design. The resulting high-temperature, high-pressure gas is ejected through the gas passage of the nozzle base, along with the propellant gas from the rear combustion chamber, through the radial nozzles, achieving radial thrust release of the engine.

[0017] In summary, this application includes at least the following beneficial technical effects:

[0018] 1) Adjustable radial nozzle position

[0019] The radial nozzle base is located between the front and rear combustion chambers. It is connected to the combustion chamber through the offset gas passage on the base to form an integrated space, realizing the energy conversion of high temperature and high pressure gas. The cross-sectional mass of the front and rear combustion chambers is basically the same, which makes it easier to adjust the position within a certain range according to the overall mass matching requirements of the projectile.

[0020] 2) Internal ballistic performance optimization

[0021] The two combustion chambers, front and rear, can be designed with end faces or other propellant configurations. Compared to a single propellant charge, it is easier to use propellants with suitable burning rates. The internal ballistic performance can be optimized through the joint design of the combustion surfaces.

[0022] 3) Increase the amount of propellant loaded in the engine

[0023] The engine of this invention employs two combustion chambers, front and rear. Under the premise of adjusting the radial nozzle position to match the center of mass of the entire projectile, the space in the front compartment can be effectively utilized to increase the amount of propellant in the engine. Attached Figure Description

[0024] Figure 1 The diagram shows a schematic of the structure of a solid rocket motor with an adjustable radial nozzle position.

[0025] Figure 2 The diagram shows the flow path of the combustion gas inside a solid rocket motor.

[0026] Figure 3 The diagram shown is a schematic of the radial base assembly.

[0027] Figure 4 The diagram shown is an exploded view of the base assembly, specifically the front cover plate assembly, drive control, nozzle base diagram, and rear thermal protection structure diagram.

[0028] Figure 5 The diagram shown is a schematic of the nozzle base.

[0029] Figure 6 The diagram shown is of a meltable igniter.

[0030] In the diagram: 1. Front combustion chamber shell;

[0031] 2. Nozzle base; 21. Radial nozzle; 22. Through-hole; 23. Drive control equipment; 24. Energy equipment; 25. Electrical connector; 26. Front annular step; 27. Rear annular step;

[0032] 3. Front cover assembly; 31. Heat protection plate; 32. Front cover;

[0033] 4. Rear thermal protection structure; 41. Gas inlet; 42. Gas pipeline; 43. Fusible igniter; 5. Rear combustion chamber shell; 6. Base assembly. Detailed Implementation

[0034] Referring to the accompanying drawings illustrating embodiments of the invention, the invention will be described in more detail below. However, the invention can be implemented in many different forms and should not be construed as being limited to the embodiments set forth herein. Rather, these embodiments are presented to achieve a full and complete disclosure and to enable those skilled in the art to fully understand the scope of the invention. In the drawings, the dimensions and relative dimensions of layers and regions may be enlarged for clarity.

[0035] This embodiment discloses a solid rocket motor structure with adjustable radial nozzle position, such as... Figure 1 As shown, the nozzle base 2 includes a front combustion chamber, a base assembly 6, and a rear combustion chamber connected in sequence. The front combustion chamber includes a front combustion chamber housing 1 and a solid propellant located inside the front combustion chamber housing 1. The rear combustion chamber includes a rear combustion chamber housing 5 and a solid propellant located inside the rear combustion chamber housing 5. The ends of the front combustion chamber housing 1 and the rear combustion chamber housing 5 are inserted into the inner side of the end of the base assembly 6 and are connected to the base assembly 6 by a drain pin (not limited to) . This arrangement allows the position of the nozzle base 2 to be adjusted back and forth within a certain range.

[0036] like Figure 2As shown, both the front combustion chamber shell 1 and the rear combustion chamber shell 5 have an open end facing the base assembly 6. The front and rear combustion chamber shells 1 and 5 are connected in a stepped manner using an insulation layer structure, sealed with a sealing ring. The front and rear combustion chambers adopt an end-burning configuration, or are jointly optimized according to internal ballistic performance requirements. The propellant loading in the front and rear combustion chambers is matched; the dual-burning-face design, adjusted according to the nozzle seat position, is more conducive to propellant burning rate adaptation, fully utilizing the shell volume and maximizing the loading coefficient and energy level. The solid propellant in the front and rear combustion chamber shells 1 and 5 has a burning surface facing the base assembly 6, forming a dual-burning-face structure, allowing for a wider range of thrust performance adjustment.

[0037] like Figure 2 As shown, the base assembly 6 includes a nozzle base 2, a front cover plate assembly 3 connected to both ends of the nozzle base 2, a rear thermal protection structure 4, and an integrated device arranged inside the nozzle base 2. The front cover plate assembly 3 is connected to one end of the nozzle base 2 and sealed with a sealing ring. The rear thermal protection structure 4 is connected to the other end of the nozzle base 2 and sealed with a sealing ring. The integrated device is encapsulated between the front cover plate assembly 3 and the rear thermal protection structure 4 and protected.

[0038] like Figure 3 , Figure 4 and Figure 5 As shown, multiple radial nozzles 21 are connected to the inner side of the nozzle base 2. The radial nozzles 21 are evenly distributed along the axial direction of the solid rocket motor. The ends of the radial nozzles 21 pass through the side wall of the nozzle base 2 and communicate with the outside. Each radial nozzle 21 is provided with an adjustment device at the other end. The adjustment device is used to seal the end of the radial nozzle 21 and to adjust the gas flow rate inside the radial nozzle 21.

[0039] The nozzle base 2 near the rear thermal protection structure 4 uses a non-metallic thermal protection structure for heat insulation. This structure can be integrally molded or RTM-formed based on the nozzle base 2, or manufactured in sections and bonded to the nozzle base 2 integrally. The rear thermal protection structure 4 has multiple gas inlets 41, each with an inlet pipe. Each gas inlet 41 is connected to a radial nozzle 21 via an inlet pipe. The thermal protection structure uses stepped joints and is bonded with a high-temperature resistant adhesive. The inlet pipes and radial nozzles 21, as well as the gas pipe 42 and the front cover plate assembly 3, are sealed using a high-temperature resistant adhesive. The rear thermal protection structure 4 includes a rear cover plate and a gas pipe 42. One end of the gas pipe 42 is connected to the rear cover plate, and the other end passes through the front cover plate assembly 3. The gas pipe 42 passes through a through-hole in the front cover plate assembly 3 and is sealed to it. A thermal protection structure is installed outside the gas pipe 42, and a gas passage is formed inside the gas pipe 42. The two combustion chambers are connected so that after ignition, the gas produced by the combustion of solid propellant in the front combustion chamber can enter the rear combustion chamber through the gas passage. Then, the gas is ejected from the rear combustion chamber through the gas inlet 41 and the radial nozzle 21, ultimately achieving the required radial thrust output.

[0040] The non-pressure-bearing sidewall of the nozzle base 2 is provided with a through-hole 22. The electrical connector 25 of the integrated equipment is installed in the through-hole 22. The radial nozzle 21 is protected by thermal protection material for all parts in contact with the high-temperature gas.

[0041] The front cover assembly 3 includes a front insulation plate and a front cover plate 32 bonded together by high-temperature adhesive. The front insulation plate is located on the side of the front cover plate 32 away from the integrated equipment.

[0042] Specifically, in this embodiment, four radial nozzles 21 are provided, forming four fan-shaped angled regions and a central portion located on the axis of the nozzle base 2. The angled regions and the central portion are used to arrange integrated equipment and gas pipelines 42. The integrated equipment includes a drive control device 23, an energy device 24, and an electrical connector 25. The energy device 24 supplies power to the drive control device 23, which drives the movement of the nozzles and receives control commands from the projectile via the electrical connector 25. In this embodiment, the drive control device 23 is located in the central portion, while the energy device 24 and gas pipeline 42 are located in the angled regions. Figure 4 As shown, this embodiment provides an installation method for a drive control device 23, with electrical connectors 25 provided on both sides of the nozzle base 2.

[0043] like Figure 3 , Figure 4 and Figure 5As shown, the base assembly 6 is designed with a large-opening flat end cap structure. One area of ​​the radial nozzle 21 is used to connect the gas passage between the front and rear combustion chambers, and the rest is used to install integrated equipment. Topology optimization design and lightweight alloy 3D additive manufacturing can be used to improve the efficiency of structural space utilization.

[0044] Figure 6 The meltable igniter is designed to match the profile of the gas pipeline 42 on the base 2 of the radial nozzle 21, and features good structural fixation and space utilization efficiency. The meltable igniter is installed inside the gas pipeline 42 and aligned with its profile. The ignition wire is connected to the cable network via the pressure-resistant electrical connector 25 interface in the equipment area on the nozzle base 2, and controlled by the electrical interface of the drive control device 23.

[0045] The nozzle base 2 is cylindrical, and its inner wall is provided with a front annular step 26 and a rear annular step 27. The front cover plate assembly 3 abuts against the front annular step 26 on the side opposite to the rear annular step 27, and the rear thermal protection structure 4 abuts against the rear annular step 27 on the side opposite to the front annular step 26. The non-metallic protective structures are sealed with a heat-resistant adhesive, and the metallic structures are sealed with sealing rings. Specifically, the front cover plate 32 is sealed to the front annular step 26 of the nozzle base 2 by a matching sealing ring, and the rear thermal protection structure 4 is sealed to the rear annular step 27 of the nozzle base 2 by a heat-resistant adhesive. Pin holes are provided at both ends of the nozzle base 2. The end of the front combustion chamber housing 1 is inserted into the interior of one end of the nozzle base 2, and the end of the rear combustion chamber housing 5 is inserted into the interior of the other end of the nozzle base 2. The front combustion chamber housing 1 presses the front cover plate assembly 3 onto the surface of the front annular step 26, and the rear combustion chamber housing 5 presses the rear thermal protection structure 4 onto the surface of the rear annular step 27. The exhaust pin passes through the pin hole and connects with the front combustion chamber housing 1 or the rear combustion chamber housing 5, thereby realizing the connection between the front combustion chamber housing 1 and the nozzle base 2, and the connection between the rear combustion chamber housing 5 and the nozzle base 2.

[0046] The base with the radial nozzle 21 is placed in the middle and connected to the front and rear combustion chambers through radial exhaust pins, forming a complete combustion chamber shell and propellant loading structure. Servo and other drive control equipment 23 is arranged in the included area of ​​the radial nozzle 21 on the base, and connected to the control cable network through the side wall penetration hole 22. A gas passage is set in the included area of ​​one of the radial nozzles 21, which is encapsulated by the body head and the installable head. The front and rear heads and the gas passage are all protected by a high-temperature resistant thermal protection structure. The propellant loading in the front and rear combustion chambers is optimized according to the internal ballistic performance requirements. It is connected to the base through the gas passage, forming an integrated flow space. During operation, the high-temperature gas flows into the radial nozzle 21 through the gas inlet 41, thereby generating thrust. The radial nozzle 21 base 2 is located in the middle of the front and rear combustion chambers, with good position adjustment capability to adapt to the matching requirements of different missile centers of mass for the radial position of the nozzle. The space utilization of the propellant loading structure is higher, improving the engine energy density.

[0047] The implementation principle of this invention is as follows:

[0048] The installation steps are as follows: The integrated equipment is installed inside the nozzle base 2. Then, the front cover plate assembly 3 and the rear thermal protection structure 4 are sealed and connected to the inner sides of both ends of the nozzle base 2. The inlet pipe is inserted into the radial nozzle 21 and sealed to the radial nozzle 21. The gas pipeline 42 passes through the front cover plate assembly 3 and is sealed to the front cover plate assembly 3. Finally, the front combustion chamber shell 1 and the rear combustion chamber shell 5 are connected to both ends of the nozzle base 2, respectively. 1. The radial nozzle 21 base 2 is located between the front and rear combustion chambers, possessing excellent position adjustment design capabilities to adapt to the matching requirements of different missile centers of mass on the radial position of the nozzle. This results in higher space utilization of the propellant structure and improved engine energy density.

[0049] When the solid propellant engine is in use, the solid propellant in the front combustion chamber housing 1 and the rear combustion chamber housing 5 burns to produce gas. The gas produced in the front combustion chamber housing 1 passes through the gas passage into the rear combustion chamber housing 5. Then, the gas located in the rear combustion chamber enters the radial nozzle 21 from the gas inlet 41 and is finally ejected from the radial nozzle 21, thereby generating thrust.

[0050] The present invention has been described in detail above with reference to specific embodiments and exemplary examples; however, these descriptions should not be construed as limiting the present invention. Those skilled in the art will understand that various equivalent substitutions, modifications, or improvements can be made to the technical solutions and embodiments of the present invention without departing from the spirit and scope of the invention, and all such modifications and improvements fall within the scope of the present invention. The scope of protection of the present invention is defined by the appended claims.

[0051] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A solid rocket motor structure with adjustable radial nozzle position, characterized in that: The system includes a front combustion chamber housing (1), a base assembly (6), and a rear combustion chamber housing (5) connected in sequence. Solid propellant is provided in both the front combustion chamber housing (1) and the rear combustion chamber housing (5). The base assembly (6) is provided with a gas passage connecting the front combustion chamber housing (1) and the rear combustion chamber housing (5) so that the gas generated in the front combustion chamber housing (1) and the rear combustion chamber housing (5) can flow through the gas passage. The base assembly (6) is provided with multiple radial nozzles (21). One end of the radial nozzle (21) is connected to the front combustion chamber or the rear combustion chamber, and the other end is connected to the outside.

2. The solid rocket motor structure with adjustable radial nozzle position according to claim 1, characterized in that: The base assembly (6) includes a nozzle base (2), a front cover plate assembly (3) connected to both ends of the nozzle base (2), and a rear thermal protection structure (4). The radial nozzle (21) is located inside the nozzle base (2). One end of the radial nozzle (21) is connected to the nozzle base (2) and communicates with the outside, while the other end is closed. The rear thermal protection structure (4) is provided with multiple gas inlets (41), and each gas inlet (41) is connected to a radial nozzle (21).

3. The solid rocket motor structure with adjustable radial nozzle position according to claim 2, characterized in that: The gas inlet (41) of the rear thermal protection structure (4) is provided with an inlet pipe, which is inserted into the radial nozzle (21). Each gas inlet (41) is connected to a radial nozzle (21) through the inlet pipe.

4. A solid rocket motor structure with adjustable radial nozzle position according to claim 2, characterized in that: the rear thermal protection structure (4) is provided with a gas pipeline (42), the other end of which passes through the front cover assembly (3) to connect the front combustion chamber housing (1) and the rear combustion chamber housing (5).

5. The solid rocket motor structure with adjustable radial nozzle position according to claim 2, characterized in that: An integrated device is arranged inside the nozzle base (2). The non-pressure-bearing sidewall of the nozzle base (2) is provided with a through-hole (22). The electrical connector (25) of the integrated device is installed in the through-hole (22).

6. The solid rocket motor structure with adjustable radial nozzle position according to claim 2, characterized in that: The inner wall of the nozzle base (2) is provided with a front annular step (26) and a rear annular step (27). The front cover plate assembly (3) abuts against the side of the front annular step (26) away from the rear annular step (27). The rear thermal protection structure (4) abuts against the side of the rear annular step (27) away from the front annular step (26). Pin holes are provided at both ends of the nozzle base (2). The end of the front combustion chamber housing (1) is inserted into the interior of one end of the nozzle base (2), and the end of the rear combustion chamber housing (5) is inserted into the interior of the other end of the nozzle base (2). The exhaust pin passes through the pin hole and connects with the front combustion chamber housing (1) or the rear combustion chamber housing (5) to realize the connection between the front combustion chamber housing (1) and the nozzle base (2) and the connection between the rear combustion chamber housing (5) and the nozzle base (2).

7. The solid rocket motor structure with adjustable radial nozzle position according to claim 5, characterized in that: The front cover assembly (3) includes a front insulation plate and a front cover plate (32) bonded together by high-temperature adhesive. The front insulation plate is located on the side of the front cover plate (32) away from the integrated equipment.

8. The solid rocket motor structure with adjustable radial nozzle position according to claim 4, characterized in that: The rear thermal protection structure (4) is a protective layer that fully covers the surface away from the nozzle base (2) for heat insulation and ablation prevention; a thermal protection structure is provided outside the gas pipeline (42); a thermal protection structure is provided outside the radial nozzle (21).

9. The solid rocket motor structure with adjustable radial nozzle position according to claim 1, characterized in that: The front and rear combustion chambers adopt an end-burning propellant configuration, and simultaneously generate combustion gases.

10. A solid rocket motor structure with adjustable radial nozzle position according to claim 1, characterized in that: The meltable igniter (43) is fixed in the gas passage of the base assembly (6) and matches the passage profile.

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