Miniature solid rocket nozzle structure and miniature solid rocket engine
By designing the throat liner structure and swirl channel, the problem of alumina deposition in the throat liner of the micro solid rocket motor was solved, achieving stable engine operating pressure and structural integrity, and meeting the design weight and cost requirements.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- SHAANXI PULIMEI MATERIAL TECH CO LTD
- Filing Date
- 2026-02-02
- Publication Date
- 2026-05-12
AI Technical Summary
When using aluminum-containing propellants, the deposition of alumina at the throat liner in micro solid rocket engines leads to a reduction in throat diameter, and in severe cases, the engine casing disintegrates. Furthermore, propellants without aluminum powder have low energy density, or high-energy composite propellants are expensive and have low density, making it difficult to meet design requirements.
A throat liner structure was designed, including a converging section, a straight section, and an expanding section. A gradient functional structure was formed by using high thermal conductivity materials and plasma spraying process. A swirling groove was set inside the throat liner. The swirling flow field and high-temperature gas were used to heat the alumina to keep it in a liquid state and avoid solid deposition.
This effectively prevents aluminum oxide buildup in the throat, ensuring stable engine operating pressure, avoiding damage to the casing, and meeting engine design weight and cost constraints.
Smart Images
Figure CN121630604B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of solid rocket engine technology, and specifically relates to a micro solid rocket nozzle structure and a micro solid rocket engine. Background Technology
[0002] In practical applications, the small overall mass of micro solid rocket motors results in relatively low engine thrust, leading to nozzle throat diameters of less than φ3.5mm. The propellant typically contains aluminum powder. When alumina and other metal oxides produced in the combustion chamber pass through the throat, the initial throat liner temperature is below 60℃, causing the liquid metal oxides to deposit. This results in the throat liner size shrinking to less than φ2.0mm in diameter, causing the engine operating pressure to severely exceed the design maximum pressure. In severe cases, this can lead to engine casing disintegration and mission failure.
[0003] Using aluminum-free propellants can usually solve the problem of aluminum powder deposition at the throat liner, but it has the following problems: 1. Conventional double-base propellants have low energy density, which increases the size and weight of the engine structure and does not meet the engine design weight limit requirements. 2. High-energy composite propellants are expensive and have low density, and they also exceed the cost limit design requirements of low-cost engines. Summary of the Invention
[0004] The purpose of this invention is to overcome the defects in the prior art and provide a micro solid rocket nozzle structure.
[0005] In a first aspect, the present invention provides a micro solid rocket nozzle structure, including a nozzle housing, wherein an insulation layer is provided on the inner wall of the nozzle housing, and a throat liner is provided on the inner wall of the insulation layer.
[0006] The throat liner includes a converging section, a straight section, and an expanding section, wherein the converging half-angle of the converging section is 15° to 30°;
[0007] The throat liner has a thickness of 0.5 to 1.0 mm and is conformal to the insulation layer.
[0008] A further option is that the insulation layer is made of high-silica or carbon phenolic resin through molding and is bonded to the inside of the nozzle housing.
[0009] A further option is that the converging section, straight section, and expanding section of the throat liner are integrally formed.
[0010] A further option is to integrally form the throat liner by plasma spraying onto the insulation layer.
[0011] A further embodiment is that the throat liner includes a support liner, a gradient liner, and a working liner that are integrally sprayed from the inside out.
[0012] The supporting liner is ceramic powder plasma-sprayed onto the insulation layer;
[0013] The gradient liner is a mixture of ceramic powder and metal powder plasma-sprayed onto the support liner, and the metal powder content is higher than the ceramic powder content on the side of the gradient layer away from the support liner.
[0014] The working liner is metal powder plasma-sprayed onto the gradient liner.
[0015] A further embodiment is that the ceramic powder is at least one of zirconium carbide, silicon carbide, zirconium oxide, and zirconium boride;
[0016] The metal powder is at least one of molybdenum powder, molybdenum alloy powder, tungsten powder, and tungsten alloy powder.
[0017] A further embodiment is that a swirling groove is formed on the inner surface of the throat liner so that the gas flows through the converging section and generates swirling flow. The swirling groove extends from the converging section to the straight section.
[0018] A further option is to polish the inner surface of the throat liner and the inner surface of the swirl groove to a mirror finish.
[0019] A further embodiment is that the pitch of the swirl channel is 0.5 to 1.5 times the inner diameter of the throat liner converging section inlet, and the helix angle is 10° to 25°; the depth of the swirl channel is 0.1 to 0.3 mm, and the width is 0.2 to 0.5 mm.
[0020] In a second aspect, the present invention provides a micro solid rocket motor comprising the micro solid rocket nozzle structure described above.
[0021] Compared with the prior art, the beneficial effects of the present invention are as follows:
[0022] The throat liner of this invention comprises a convergent section, a straight section, and an expanding section of one integrated structure. This eliminates the gaps and thermal resistance discontinuities that may exist in traditional segmented or spliced throat liners, ensuring a smooth heat conduction path. This allows the entire throat area to reach the required high temperature uniformly and quickly, reducing the formation of solid deposits at the source. Furthermore, the throat liner of this invention employs a thin-walled structure made of a high thermal conductivity material, enabling the inner surface temperature of the throat liner to be rapidly heated from an initial low temperature to above the melting point of alumina within milliseconds of engine ignition. Liquid alumina remains in a liquid state as it flows through the throat liner wall, and its adhesion to the solid wall is much less than that of solid deposits, facilitating its removal using airflow.
[0023] The convergence section half-angle design of this invention is within the range of 15° to 30°. While ensuring a compact structure, it allows for a smooth flow acceleration process when the combustion gas enters the throat from the combustion chamber. Furthermore, as the combustion gas flows along the convergence section, its velocity vector can be decomposed into a normal component perpendicular to the wall and a tangential component parallel to the wall. The tangential force exerted by the airflow on the liquid alumina on the wall is significantly enhanced. Under the strong airflow, the liquid alumina film is more easily detached from the wall and flows along the convergence section to the straight section and the expansion section, ultimately being carried out of the nozzle by the high-speed mainstream gas, effectively preventing alumina accumulation in the throat.
[0024] This invention employs a plasma spraying process to directly and integrally form a throat liner with a gradient functional structure on the insulation layer. The material composition of the support liner, gradient liner, and working liner of the throat liner achieves a continuous gradient and is integrally sprayed. The working layer is coated with metal powder to achieve rapid heat conduction, maintain a high wall temperature, and withstand the direct impact of combustion gases. The composition of the gradient liner continuously changes from ceramic to metal, achieving a gradient match in the coefficient of thermal expansion between the metal working liner and the ceramic support liner. The support liner is coated with ceramic powder, effectively preventing excessive heat transfer to the rear insulation layer and shell, playing a crucial role in thermal protection and protecting the main engine structure from damage.
[0025] This invention features a swirling groove extending from a converging section to a straight section on the inner surface of the throat liner. When high-temperature combustion gas flows through the inlet of the converging section of the throat liner with the spiral groove, the spiral groove guides the combustion gas, creating a swirling field within the throat channel. The centrifugal force generated by the swirling field causes alumina particles to move towards the throat liner wall. This reduces the particle throughput directly flowing through the throat, and the swirling flow transforms the path of the alumina particles into a spiral motion, extending the residence time of the alumina particles in the high-temperature combustion gas region. This ensures that the alumina particles are completely heated and melted, preventing them from solidifying on the inner wall of the throat liner. Attached Figure Description
[0026] The following figures are for illustrative purposes only and are not intended to limit the scope of the invention, wherein:
[0027] Figure 1 : Schematic diagram of the cross-section of the nozzle structure of the present invention;
[0028] Figure 2 : Schematic diagram of the location of the vortex channel;
[0029] In the diagram: 1. Throat liner; 2. Insulation layer; 3. Nozzle housing; 4. Swirl channel. Detailed Implementation
[0030] To make the objectives, technical solutions, design methods, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are for illustrative purposes only and are not intended to limit the invention.
[0031] like Figure 1 As shown, this invention provides a micro solid rocket nozzle structure, including a nozzle shell 3, an insulation layer 2 disposed on the inner wall of the nozzle shell 3, and a throat liner 1 disposed on the inner wall of the insulation layer 2. The insulation layer 2 is made of high-silica or carbon phenolic resin through molding and is bonded to the inside of the nozzle shell 3. The throat liner 1 is integrally formed by plasma spraying on the insulation layer 2. The throat liner 1 is conformal to the insulation layer 2 and includes a converging section, a straight section, and a expanding section, which are integrally formed. The convergence half-angle of the converging section is 15°–30°, and the thickness of the throat liner 1 is 0.5–1.0 mm. This reduces the heat capacity of the throat liner, allowing its inner surface temperature to be rapidly heated from an initial low temperature to above the melting point of alumina within milliseconds of engine ignition. This prevents the solidification of liquid alumina as it flows through the throat liner and facilitates the removal of liquid alumina by airflow.
[0032] In this embodiment, the throat liner 1 is composed of a thermally conductive gradient structure made of two different materials, with its thermal conductivity gradually decreasing from the inner surface in contact with the combustion gas towards the insulation layer 2. This ensures both rapid heat conduction and heat insulation. Specifically, the thermally conductive gradient structure includes: a support liner, a gradient liner, and a working liner, integrally sprayed from the inside out. In this embodiment, the support liner is ceramic powder plasma-sprayed onto the insulation layer; the gradient liner is a mixture of ceramic powder and metal powder plasma-sprayed onto the support liner, with the metal powder content being higher than the ceramic powder content on the side of the gradient layer furthest from the support liner; the working liner is metal powder plasma-sprayed onto the gradient liner. The ceramic powder is at least one of zirconium carbide, silicon carbide, zirconium oxide, and zirconium boride; the metal powder is at least one of molybdenum powder, molybdenum alloy powder, tungsten powder, and tungsten alloy powder.
[0033] like Figure 2 As shown, to further reduce alumina particle deposition at the throat liner 1, a swirling groove 4 is machined on the inner surface of the throat liner 1 using laser etching. This allows the gas flow to generate swirling flow as it passes through the converging section. The swirling groove 4 extends from the converging section to the straight section, ensuring that the swirling flow fully develops within the converging section. Furthermore, the inner surface of the throat liner 1 and the inner surface of the swirling groove 4 are precision ground and polished to achieve a surface roughness Ra ≤ 0.8 μm, reducing airflow resistance and the probability of particle adsorption. In this embodiment, the pitch of the swirling groove 4 is 0.5 to 1.5 times the inner diameter of the converging section inlet of the throat liner. Specifically, the pitch is 2 mm (i.e., 1 times the inner diameter of the converging section inlet of the throat liner 1, with an inner diameter of 2 mm), the helix angle is 18°, the groove depth is 0.2 mm, and the groove width is 0.3 mm.
[0034] Comparative example:
[0035] Using a high-silica phenolic molded part as the insulation layer 2, since the throat liner 1 is conformal to the insulation layer 2, the high-silica phenolic molded part presents a contour of a converging section, a straight section and an expanding section, and the convergence half angle of the converging section contour is 20°. The high-silica phenolic molded part is fixed to the inner wall of the nozzle housing 3 by an adhesive.
[0036] The throat liner 1 is integrally formed on the inner surface of the insulation layer 2 using a plasma spraying process, including the following steps:
[0037] A ceramic composite powder composed of zirconium carbide and silicon carbide is sprayed to form a support liner with a thickness of about 0.4 mm. The mass ratio of the ceramic composite powder is 8:2.
[0038] By using a dual powder feeding system, the powder feeding ratio of molybdenum powder to the above-mentioned ceramic composite powder is continuously changed, transitioning from pure ceramic to pure metal, forming a gradient liner with a continuous compositional gradient change of about 0.3 mm in thickness.
[0039] Molybdenum powder is sprayed to form a working liner with a thickness of about 0.3 mm, and the final total thickness of the throat liner is 1.0 mm.
[0040] A spiral swirl groove 4 is machined on the inner surface from the inlet of the converging section to the straight section of the throat liner 1. The groove is 0.2 mm deep, 0.3 mm wide, and has a helix angle of 15°. The pitch is 1.0 times the inner diameter of the inlet of the converging section. The inner wall of the throat liner 1 and the swirl groove 4 are finely polished to achieve a mirror effect.
[0041] The aforementioned nozzle was assembled into a micro solid rocket motor, and a static ignition test was conducted using a composite propellant containing aluminum powder. The motor's operating time was 500 ms. Before and after operation, the measured nozzle throat diameter changed from φ3.5 mm to φ3.3 mm, with a contraction rate of approximately 5.7%. The combustion chamber operating pressure remained stable within the range of 9 MPa to 10 MPa, and the motor structure remained intact and operated smoothly.
[0042] A comparative test was conducted using the same engine model, with only the nozzle throat liner replaced with a traditional design. Ignition tests were performed under the same aluminum-containing propellant formulation and test conditions. After engine operation, severe alumina slag deposition and erosion occurred on the graphite throat liner surface. The measured nozzle throat diameter decreased from the initial φ3.5mm to φ2.5mm, a shrinkage rate greater than 40%. This caused the combustion chamber pressure to rise from the design value of 9MPa to 29MPa, severely exceeding the casing's pressure bearing limit. After the test, the engine casing bulged and deformed.
[0043] In the comparative example, due to the poor thermal conductivity and large wall thickness of graphite material, the throat liner temperature rises slowly, making it impossible to keep alumina in a liquid state. This leads to the direct adhesion and accumulation of solid particles, causing rapid blockage of the throat. This invention achieves instantaneous heating to keep alumina in a liquid state through high thermal conductivity materials and a thin-walled structure. Furthermore, the inclusion of a swirl channel reduces particle throughput through centrifugal separation, preventing the accumulation of low-thermal-conductivity alumina in the converging and straight sections of the throat liner, thus preventing the large-scale deposition of alumina at the throat liner.
[0044] This application also discloses a micro solid rocket motor, including the nozzle structure disclosed in the above embodiments. Because it has the above-described nozzle structure, the micro solid rocket motor disclosed in this application possesses all the technical effects of the above-described nozzle structure, which will not be repeated here.
[0045] The various embodiments of the present invention have been described above. These descriptions are exemplary and not exhaustive, nor are they limited to the disclosed embodiments. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the described embodiments. The terminology used herein is chosen to best explain the principles, practical application, or technical improvements to the embodiments in the market, or to enable others skilled in the art to understand the embodiments disclosed herein.
Claims
1. A micro solid rocket nozzle structure, characterized in that, Includes a nozzle housing (3), the inner wall of which is provided with an insulation layer (2), and the inner wall of the insulation layer (2) is provided with a throat liner (1). The throat liner (1) includes a converging section, a straight section and an expanding section. The converging section, straight section and expanding section of the throat liner (1) are integrally formed. The converging half angle of the converging section is 15° to 30°. The throat liner (1) has a thickness of 0.5 to 1.0 mm, and the throat liner (1) is conformal with the insulation layer (2); The throat liner (1) is integrally formed by plasma spraying on the insulation layer (2); The throat liner (1) includes a support liner, a gradient liner, and a working liner that are integrally sprayed from the inside out; The supporting liner is ceramic powder plasma-sprayed onto the insulation layer; The gradient liner is a mixture of ceramic powder and metal powder plasma-sprayed onto the support liner, and the metal powder content is higher than the ceramic powder content on the side of the gradient layer away from the support liner. The working liner is metal powder plasma-sprayed onto the gradient liner.
2. The micro solid rocket nozzle structure according to claim 1, characterized in that, The insulation layer (2) is made of high silica or carbon phenolic resin and is bonded to the inside of the nozzle housing (3).
3. The micro solid rocket nozzle structure according to claim 2, characterized in that, The ceramic powder is at least one of zirconium carbide, silicon carbide, zirconium oxide, and zirconium boride. The metal powder is at least one of molybdenum powder, molybdenum alloy powder, tungsten powder, and tungsten alloy powder.
4. The micro solid rocket nozzle structure according to claim 3, characterized in that, The throat liner (1) has a swirling groove (4) on its inner surface so that the gas flows through the converging section and generates a swirling flow. The swirling groove (4) extends from the converging section to the straight section.
5. The micro solid rocket nozzle structure according to claim 4, characterized in that, The inner surface of the throat liner (1) and the inner surface of the swirl groove (4) are polished to a mirror finish.
6. The micro solid rocket nozzle structure according to claim 5, characterized in that, The pitch of the swirl channel (4) is 0.5 to 1.5 times the inner diameter of the throat liner converging section inlet, and the helix angle is 10° to 25°; the depth of the swirl channel (4) is 0.1 to 0.3 mm, and the width is 0.2 to 0.5 mm.
7. A miniature solid rocket motor, characterized in that, The micro solid rocket nozzle structure includes any one of claims 1-6.