Single-chamber double-thrust solid rocket engine charge structure and process

Abstract

The application discloses a single-chamber double-thrust solid rocket engine charging structure and process, which comprises a grain, a heat insulation layer and a combustion chamber shell. The inner surface of the combustion chamber shell is subjected to sand blasting treatment and uniformly sprayed with adhesive, and then is bonded with the heat insulation layer. The inner surface of the heat insulation layer is roughened and uniformly sprayed with a lining layer. The grain is integrally solidified on the inner surface of the heat insulation layer by adopting a vacuum wall pasting pouring process. The solid rocket engine combustion chamber charging structure of the application adopts a variable cross-section integrated design, improves the complexity of the traditional series combination type single-chamber double-thrust star hole grain forming process, avoids quality problems such as debonding, slag inclusion, air bubbles and cracks of the combination type star hole grain in the forming process of the transition section, guarantees the charging structure integrity and interior ballistic performance, realizes the minimum initial thrust required for the missile to leave the rack and the characteristic requirement of the front position of the missile mass center, effectively solves the problem of missile flight instability, improves the missile flight reliability, and is successfully applied in a type.

Description

TECHNICAL FIELD

[0001] The application relates to a solid rocket engine combustion chamber charge structure, in particular to a single-chamber double-thrust solid rocket engine charge structure and process. BACKGROUND

[0002] The solid rocket engine combustion chamber is one of important components of a missile engine, is a place for solid propellant filling and thrust generation, is a part of a missile body structure, bears internal pressure and external load during missile flight, and provides support and connection for external interfaces of the missile.

[0003] A certain type of missile adopts an internal sliding rail launching overall design scheme, and it is required that the missile must fly away from the carrier at a high speed after leaving the carrier to avoid affecting the carrier. In addition, the axial length of the missile is limited during overall design, the axial center of mass of the whole missile is offset to the rear, so that the missile has a risk of instability during flight. In combination with the above design factors, under the premise of meeting the overall impulse requirement of the engine, if a traditional engine combustion chamber charge structure design is adopted, the overall requirement cannot be met. SUMMARY

[0004] The technical problem solved by the application is to overcome the shortcomings of the prior art, provide a single-chamber double-thrust solid rocket engine combustion chamber charge structure and process with a front-mounted center of mass, effectively solve the flight instability problem caused by the axial center of mass offset of the missile, thereby improving the flight reliability of the missile, and at the same time ensuring the initial thrust required for the missile to leave the carrier.

[0005] Further, the combustion chamber charge structure design effectively improves the complexity of the forming process of the traditional series combined single-chamber double-thrust star hole grain, avoids quality problems such as debonding, slag inclusion, bubbles and cracks in the forming process of the transition section of the combined grain, and saves production cost.

[0006] The technical solution of the application is:

[0007] A single-chamber double-thrust solid rocket engine charge structure and process, comprising a grain, an adiabatic layer and a combustion chamber shell, the adiabatic layer is bonded to the inner wall of the combustion chamber shell, the grain is bonded to the inner wall of the adiabatic layer, and an inner hole with a small front and a large rear is arranged at the center axis position of the grain; the grain sequentially comprises a cruising section, a transition section and an initial boost section from the head to the tail, and the inner diameter of the transition section gradually decreases from the initial boost section to the cruising section.

[0008] The ratio of the thickness of the propellant grain to the radius of the propellant grain in the initial boost stage is 1:5, the ratio of the length to the length of the propellant grain is 1:3, the ratio of the thickness of the propellant grain to the radius of the propellant grain in the cruise stage is 1:1.6, and the ratio of the length to the length of the propellant grain is 1:2.2, the initial boost stage of the propellant grain is transitioned to the cruise stage through a circular arc transition section, and the ratio of the length of the transition section to the length of the propellant grain is 1:4.4.

[0009] The heat insulation layer is designed with variable thickness according to the burning law of the propellant grain, and comprises four sections from the head to the tail, i.e., a head artificial debonding heat insulation layer, a barrel section heat insulation layer, a thickened section heat insulation layer, and a tail artificial debonding heat insulation layer.

[0010] The thicknesses of the heat insulation layer from the head artificial debonding heat insulation layer to the tail artificial debonding heat insulation layer are 6mm, 1mm, 4.5mm, and 8mm, respectively.

[0011] The heat insulation layer is provided with the head artificial debonding heat insulation layer and the tail artificial debonding heat insulation layer at both ends, which artificially divides the heat insulation layer into two layers, the outer layer being connected with the combustion chamber shell and the inner layer being connected with the propellant grain, so as to improve the bonding interface stress between the heat insulation layer and the propellant grain caused by the change of the environmental temperature.

[0012] The heat insulation layer adopts homogeneous three-element ethylene-propylene rubber heat insulation material as the base material.

[0013] The propellant grain adopts variable cross-section integrated design, and the propellant grain adopts hydroxyl-terminated polybutadiene (HTPB) propellant.

[0014] The inner surface of the combustion chamber shell is subjected to sand blasting treatment.

[0015] A single-chamber double-thrust solid rocket engine charge structure and process, which adopts the combustion chamber charge structure described in any one of the above embodiments, comprises:

[0016] The inner surface of the combustion chamber shell is subjected to sand blasting treatment and uniform adhesive spraying;

[0017] The outer surface of the heat insulation layer is subjected to interface bonding with the inner surface of the combustion chamber by vacuum process, and is subjected to solidification treatment after bonding;

[0018] The inner surface of the heat insulation layer is subjected to roughening treatment according to process requirements, and the inner surface of the heat insulation layer is uniformly sprayed with a lining layer and an adhesive;

[0019] The propellant grain is integrally solidified on the inner surface of the heat insulation layer by vacuum wall-adhering casting process.

[0020] In summary, the present application has at least the following beneficial technical effects:

[0021] (1) This invention adopts a single-chamber dual-thrust solid rocket engine propellant structure and process, which effectively solves the flight instability problem caused by the axial center of mass shift of the missile, thereby improving the flight reliability of the missile. At the same time, it ensures the initial thrust required for the missile to leave the launcher. In addition, this combustion chamber propellant structure design also effectively improves the complexity of the traditional series combined single-chamber dual-thrust star-hole propellant grain forming process, and avoids quality problems such as debonding, slag inclusion, bubbles and cracks that occur in the transition section forming process of the combined propellant grain, thus saving production costs;

[0022] (2) Make more effective use of existing space, increase the initial thrust of the engine by changing the propellant configuration, ensure the minimum initial thrust requirement for missile deorbiting, and at the same time use the integrated propellant design structure of the solid rocket engine combustion chamber to ensure the forward positioning of the combustion chamber center of mass, thereby correcting the overall center of mass of the missile and improving the flight reliability of the missile. Attached Figure Description

[0023] Figure 1 This is a cross-sectional view of a single-chamber dual-thrust solid rocket motor propellant loading structure according to the present invention;

[0024] Figure 2 This is a schematic diagram of the inner surface of the combustion chamber shell;

[0025] Figure 3 This is a schematic diagram of the inner and outer surfaces of the insulation layer.

[0026] Explanation of reference numerals in the attached diagram: 1. Propellant charge; 2. Insulation layer; 3. Combustion chamber shell;

[0027] 11. Cruise phase; 12. Transition phase; 13. Initial boost phase;

[0028] 21. Head section artificially debonded insulation layer; 22. Tube section insulation layer; 23. Thickened section insulation layer; 24. Tail section artificially debonded insulation layer; Detailed Implementation

[0029] The present application will now be described in further detail with reference to the accompanying drawings and specific embodiments:

[0030] This application discloses a propellant loading structure and process for a single-chamber, dual-thrust solid rocket motor, such as... Figure 1 As shown, the combustion chamber charging structure includes a propellant grain 1, an insulation layer 2, and a combustion chamber shell 3. The insulation layer 2 is bonded to the inner wall of the combustion chamber shell 3, and the propellant grain 1 is bonded to the inner wall of the insulation layer 2. The propellant grain 1 has an inner hole that is smaller at the front and larger at the back at the central axis position.

[0031] like Figure 1As shown, the propellant grain 1 is integrally formed by using a hydroxyl-terminated polybutadiene propellant combined with a vacuum wall-adhering casting process. The propellant grain 1 has a hole diameter of Ф231 mm, a length of Ф382.1 mm and a thickness of 28.1 mm in the initial boost stage 13. The propellant grain 1 has a hole diameter of Ф110 mm, a length of Ф520 mm and a thickness of 92.6 mm in the cruise stage 11. The two-stage propellant grain 1 is connected through a circular arc transition section 12 having a length of 262.7 mm. The propellant grain 1 is designed as a variable cross-section integrated structure. By changing the hole diameter, the entire hole of the propellant grain 1 starts to burn in the initial stage, and a large initial thrust is provided. After the propellant grain 1 at the position of the initial boost stage 13 having a small thickness is burned out, the thrust is reduced, and the two thrusts of the integrated charge are realized. At the same time, by changing the hole diameter of the propellant grain 1, the center of mass of the combustion chamber charge structure is moved forward, and the flight instability problem caused by the axial center of mass deviation of the missile is effectively solved.

[0032] As shown in Figure 1 , the thermal insulation layer 2 uses homogeneous ethylene-propylene-diene rubber as a base material. The thermal insulation layer 2 is designed as a variable thickness in four steps from the head to the tail in a step type. The thickness of the thermal insulation layer 2 from the head to the tail is 6 mm, 1 mm, 4.5 mm and 8 mm in sequence. The position of the thermal insulation layer 2 having a thickness of 1 mm corresponds to the propellant grain of the cruise stage 11. The position of the thermal insulation layer 2 having a thickness of 4.5 mm corresponds to the propellant grain of the transition section 12 and the boost stage 13. By setting the variable thickness of the thermal insulation layer 2, the thermal insulation layer still has a certain residual thickness after the propellant grain 1 is burned out, and thus can play a certain thermal protection role for the combustion chamber shell 3.

[0033] As shown in Figure 1 , the combustion chamber shell 3 is made of 30Cr3SiNiMoVA ultra-high strength steel. The combustion chamber shell 3 has a structure design of a small front opening and a large rear opening. The inner threads of the front and rear openings are designed for connection. The outer diameter of the inner threads is φ300±0.2 mm. The distance between the front and rear abutting surfaces is 1180±2 mm. The wall thickness is 1.5 mm. According to the overall connection requirements, the front section and the rear end of the combustion chamber shell 3 are designed as connection interfaces for external connection.

[0034] As shown in Figure 1 , the thermal insulation layer 2 is provided with a head artificial debonding thermal insulation layer 21 to a tail artificial debonding thermal insulation layer 24 at both ends. The artificial debonding thermal insulation layers 21 to 24 are used to artificially divide the thermal insulation layer 2 into two layers. The outer layer is connected with the combustion chamber shell 3, and the inner layer is connected with the propellant grain 1. The artificial debonding thermal insulation layers 21 to 24 are used to improve the adhesive interface stress between the thermal insulation layer 2 and the propellant grain 1 caused by changes in the environmental temperature.

[0035] Specifically, the shell material of the combustion chamber shell 3 is 30Cr3SiNiMoVA ultra-high strength steel, which is integrally formed by spinning, numerical control processing and welding, the outer diameter is φ300±0.2mm, the distance between the front and rear butt joints is 1180±2mm, and the wall thickness is 1.5mm. The combustion chamber shell 3 adopts a front small opening and rear large opening structure, and the inner threads of the front and rear openings are designed for connection with other components. According to the overall mechanical interface coordination requirements, the front end and rear end of the combustion chamber shell 3 are designed for external connection interfaces to realize connection with the aircraft slide rail.

[0036] Specifically, the heat insulation layer 2 is a prefabricated part, which is prefabricated in advance according to the design requirements of the heat insulation layer and the inner surface of the combustion chamber shell 3 using a mold, and is subjected to vulcanization treatment. The material of the heat insulation layer 2 adopts homogeneous three-element ethylene-propylene rubber heat insulation material as the base material, and the thickness is designed to be variable in four sections from the head to the tail in a stepped manner, and the thicknesses are 6mm, 1mm, 4.5mm and 8mm respectively.

[0037] A single-chamber double-thrust solid rocket engine charging structure and process, comprising the following steps: the inner surface of the combustion chamber shell 3 is subjected to sand blasting treatment and uniformly sprayed with adhesive; the outer surface of the heat insulation layer 2 and the inner surface of the combustion chamber are bonded by vacuum process, and after bonding, solidification treatment is performed; the inner surface of the heat insulation layer 2 is subjected to roughening treatment according to process requirements, and the inner surface of the heat insulation layer 2 is uniformly sprayed with a lining and an adhesive; the grain 1 is integrally solidified on the inner surface of the heat insulation layer 2 by vacuum wall sticking casting process.

[0038] Specifically, after the combustion chamber shell 3 is formed, the inner surface of the combustion chamber shell 3 is subjected to sand blasting treatment, the outer surface of the heat insulation layer 2 is subjected to roughening treatment, and the inner surface of the combustion chamber shell 3 and the outer surface of the heat insulation layer 2 are uniformly sprayed with adhesive. After the adhesive is sprayed, the heat insulation layer 2 and the combustion chamber are bonded by vacuum process using an air bag, and the bonding quality needs to be ensured after bonding.

[0039] Specifically, the inner surface of the bonded heat insulation layer 2 is roughened, and the inner surface of the heat insulation layer 2 is uniformly sprayed with a lining and an adhesive. And the grain 11 is integrally formed by high-temperature solidification by vacuum wall sticking casting process.

[0040] The single-chamber double-thrust solid rocket engine charging structure and process can effectively solve the flight instability problem caused by the axial mass center offset of the missile, thereby improving the flight reliability of the missile. At the same time, the initial thrust required for the missile to leave the rack is ensured, in addition, the combustion chamber charging structure design also effectively improves the complexity of the traditional series combined single-chamber double-thrust star hole grain 1 forming process, avoids the quality problems such as debonding, slag inclusion, bubbles and cracks in the forming process of the transition section 12 of the combined grain 1, and saves the production cost.

[0041] Although the present application has been disclosed in its preferred embodiments with reference to the accompanying drawings, it is not intended to limit the present application thereto, and any person skilled in the art can make possible changes and modifications without departing from the spirit and scope of the present application, and the scope of protection of the present application should be defined by the scope of the claims.

Claims

1. A single chamber dual thrust solid rocket motor grain structure, characterized by: It comprises a propellant grain (1), an insulating layer (2) and a combustion chamber shell (3). The insulating layer (2) is bonded to the inner wall of the combustion chamber shell (3), and the propellant grain (1) is bonded to the inner wall of the insulating layer (2). The propellant grain (1) comprises a cruise section (11), a transition section (12) and an initial boost section (13) from the head to the tail. The ratio of the thickness of the initial boost section (13) of the propellant grain (1) to the radius of the propellant grain (1) is 1:5, the ratio of the length of the initial boost section (13) to the length of the propellant grain (1) is 1:3, the ratio of the thickness of the cruise section (11) of the propellant grain (1) to the radius of the propellant grain (1) is 1:1.6, and the ratio of the length of the cruise section (11) to the length of the propellant grain (1) is 1:2.

2.

2. A single chamber dual thrust solid rocket engine grain structure as claimed in claim 1, wherein: The insulating layer (2) is designed with variable thickness according to the combustion law of the propellant grain, and comprises a head artificial debonding insulating layer (21), a cylinder section insulating layer (22), a thickened section insulating layer (23) and a tail artificial debonding insulating layer (24) from the head to the tail.

3. A single chamber dual thrust solid rocket engine grain structure according to claim 2, wherein: The thickness of the insulating layer (2) from the head to the tail is 6mm, 1mm, 4.5mm and 8mm, respectively.

4. The single chamber dual thrust solid rocket engine grain structure of claim 2, wherein: The head artificial debonding insulating layer (21) and the tail artificial debonding insulating layer (24) are arranged at the two ends of the insulating layer (2), which artificially divides the insulating layer into two layers, the outer layer is connected with the combustion chamber shell (3), and the inner layer is connected with the propellant grain (1), so as to improve the bonding interface stress between the insulating layer (2) and the propellant grain (1) caused by the change of the environmental temperature.

5. A single chamber dual thrust solid rocket engine grain structure as claimed in claim 1, wherein: The insulating layer (2) adopts homogeneous EPDM as the base material.

6. A single chamber dual thrust solid rocket engine grain structure as claimed in claim 1, wherein: The propellant grain (1) adopts variable cross-section integrated design, and the propellant grain (1) adopts HTPB propellant.

7. A single chamber dual thrust solid rocket engine grain structure as claimed in claim 1, wherein: The inner surface of the combustion chamber shell (3) is subjected to sand blasting treatment.

8. A single-chamber dual-thrust solid rocket engine charge process using the combustion chamber charge structure according to any one of claims 1 to 7, characterized by: It comprises The inner surface of the combustion chamber shell (3) is subjected to sand blasting treatment and uniform adhesive spraying; The outer surface of the insulating layer (2) and the inner surface of the combustion chamber are bonded by vacuum process, and the bonding is completed and subjected to curing treatment; The inner surface of the insulating layer (2) is subjected to roughening treatment according to the process requirements, and the inner surface of the insulating layer (2) is uniformly sprayed with a lining and an adhesive; The propellant grain (1) is integrally cured on the inner surface of the insulating layer (2) by vacuum wall-pasting casting process.

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