CN112694372A

CN112694372A - Epoxy composite propellant, preparation method and application thereof

Info

Publication number: CN112694372A
Application number: CN202011564765.XA
Authority: CN
Inventors: 石仲仑; 白澔烔; 马天任
Original assignee: Individual
Current assignee: Individual
Priority date: 2020-12-25
Filing date: 2020-12-25
Publication date: 2021-04-23

Abstract

The invention discloses an epoxy composite propellant, a preparation method and application thereof, belongs to the field of solid rocket fuels, improves the defects in the prior art, reduces the processing difficulty and processing cost of fuels, and can improve the inner ballistic performance when being applied to a solid rocket engine. The epoxy group composite propellant comprises the following components in percentage by mass: 75% of ammonium perchlorate, 6.5% of 669 epoxy resin active diluent, 0.25% of catalyst, 3.3% -4% of curing agent and the balance epoxy resin. Compared with most of standard composite propellants, the composite propellant provided by the invention does not need to be added with a bonding agent to prevent AP (access point) threshing, simplifies the production flow, has good mechanical property of materials and higher strength, is not easy to become brittle at low temperature, and does not need to be provided with a support structure or can be simplified when being applied to a rocket engine with higher overload.

Description

Epoxy composite propellant, preparation method and application thereof

Technical Field

The invention belongs to the field of solid rocket fuels, and particularly relates to an epoxy composite propellant, a preparation method and application thereof.

Background

The rocket projectile is generally designed by using the already-shaped and produced propellants, including three types of biradical propellants, modified biradical propellants and composite propellants.

The existing small and medium-sized solid rocket engines mostly adopt double-base propellants or modified double-base propellants, such as cobaltic-2, plumbic-2 and the like, the processing mode is an extrusion molding process at high temperature, complex medicine types such as variable cross-section structures and the like are difficult to process, the manufacturing difficulty is high, danger is easy to occur, and multiple malignant explosion accidents occur once. And the long-storage property of the double-base propellant is poor, and a part of additives and liquid ester can seep out during long-storage, so that the fuel cannot be used.

The composite propellant has higher energy than a double-base propellant, high specific impulse, strong combustion controllability and low pressure index. Because of the adoption of casting molding, the large-scale grain can be made, is mainly used for rocket and missile with high thrust. However, the composite propellant is soft in texture and is easy to deform under the condition of high overload, and a supporting structure is required to be added to prevent deformation.

In the laboratory design stage of small and medium-sized solid rocket engines, the processing system type double-base propellant is difficult to prepare so as to meet the experimental requirements, and the characteristics of the existing system type composite propellant are difficult to meet the design requirements of the small and medium-sized rocket engines. The solid fuel which has low fuel cost and easy processing in the early design process of the small and medium-sized solid rocket engine and can reach the standard of combustion characteristics of standard fuel is lacked under the processing condition of a laboratory.

Disclosure of Invention

The invention provides an epoxy composite propellant, a preparation method and application thereof, overcomes the defects in the prior art, reduces the processing difficulty and processing cost of fuel, can improve the internal ballistic performance when being applied to a solid rocket engine, and meets the requirements of low-cost, easy-to-process and high-performance solid propellant required by early design.

In order to achieve the purpose, the invention adopts the following technical scheme:

an epoxy composite propellant consists of the following components in percentage by mass:

70-80% of ammonium perchlorate, 4-6.5% of 669 epoxy resin reactive diluent (ethylene glycol diglycidyl ether), 0.2-1.25% of catalyst, 3.3-4% of curing agent and the balance of epoxy resin;

preferably comprises the following components in percentage by mass:

75% of ammonium perchlorate, 6.5% of 669 epoxy resin active diluent, 0.25% of catalyst, 3.3% -4% of curing agent and the balance epoxy resin.

In the components, the catalyst is lead oxide, and the curing agent is an epoxy resin curing agent; aluminum powder can be added into the composite propellant, so that specific impact is improved, burning speed is improved, and acoustic unstable burning is inhibited; the adding amount of the aluminum powder is 0-16% of the mass of the composite propellant.

The preparation method of the epoxy composite propellant comprises the following steps:

step 1: weighing and mixing liquid components, wherein the liquid components comprise epoxy resin and 669 epoxy resin reactive diluent, and uniformly mixing until bubbles disappear completely for later use;

step 2: weighing and mixing solid components, wherein the solid components comprise ammonium perchlorate and lead oxide catalyst, and uniformly mixing for later use;

and step 3: weighing the curing agent, wherein the content of the curing agent is within a range of 3.3-4%, the curing agent is determined according to actual curing requirements, the content of the curing agent can be properly increased to accelerate the curing speed under the conditions of simple drug form and convenient filling, and otherwise, the content is reduced;

and 4, step 4: adding a curing agent into the liquid component for rapid mixing for 5-10s, adding the solid component into the liquid component, uniformly stirring, wherein the uniformly mixed fuel has good fluidity, is bright orange, has uniform color distribution, and generates a small amount of ammonia gas in the mixing process;

and 5: and pouring the mixture into a fuel mold in vacuum, curing and molding the mixture at normal temperature, and taking the molded mixture out of the mold.

In the above-described step, aluminum is added to the liquid component of step 1 and then mixed with the solid component.

The application of the epoxy composite propellant in the small and medium-sized solid rocket engines can embody the advantages of the epoxy composite propellant.

Has the advantages that: the invention provides an epoxy composite propellant, a preparation method and application thereof, which retain the characteristics of high specific impulse, strong combustion controllability and low pressure index of the composite propellant, further reduce the processing difficulty and the processing cost, can be produced in large quantities under the laboratory condition, can meet the requirements of low fuel cost and easy processing in the early design process of small and medium-sized rocket engines, and can reach the standard of combustion characteristics of standard fuels.

Compared with most of standard composite propellants, the composite propellant disclosed by the invention does not need to be added with a bonding agent to prevent AP (access point) threshing, simplifies the production flow, has good mechanical property of materials and higher strength, is not easy to become brittle at low temperature, and does not need to be provided with a support structure or can be simplified when being applied to a rocket engine with higher overload; compared with a double-base propellant, the composite propellant can be processed and molded at normal temperature, can be cured and molded only in one hour under the environment of 25 ℃ and relative humidity less than 90 percent, and has high curing speed and safe manufacture; the fluidity is good before solidification, the castability is good, and the material can be processed into various complex shapes (such as variable cross-section structures) according to the design requirements of an engine; the long-storage property is good, and the situation of long-storage oil leakage of the double-base propellant cannot occur; the critical pressure of the propellant is only 1MPa, the ignition pressure is low, the combustion is stable, the combustion is tiny, the flame is small, and the propellant is applied to the field of rocket projectile weapons, so that the rocket projectile solid rocket engine works stably and is not easy to expose a launching position during launching.

Drawings

FIG. 1 is a cross-sectional view of a fuel prepared in example 1 of the present invention;

FIG. 2 is a half sectional view of a rocket engine with a diameter of 82mm in example 1 of the present invention, wherein 1 is an engine housing, 2 is a nozzle, 3 is a plug, 4 is a throat insert, 5 is rocket fuel, and 6 is a thermal insulation layer;

FIG. 3 is a graph showing (a) a graph of stagnation pressure and (b) a graph of thrust in example 1 of the present invention;

FIG. 4 is a cross-sectional view of a fuel prepared in example 2 of the present invention;

FIG. 5 is a half sectional view of a rocket engine of 55mm diameter according to example 2 of the present invention, wherein 7 is an engine casing, 8 is a nozzle, 9 is a plug, 10 is a throat insert, 11 is rocket fuel, and 12 is a thermal insulation layer;

FIG. 6 is a graph showing (a) a graph of stagnation pressure and (b) a graph of thrust in example 2 of the present invention;

FIG. 7 is a cross-sectional view of a fuel prepared in example 3 of the present invention;

fig. 8 is a half-sectional view of a 37mm diameter turbo rocket in example 3 of the present invention, in which 13 is an engine case, 14 is a turbo rocket nozzle, 15 is a bulkhead, 16 is a throat insert, 17 is rocket fuel, and 18 is a cladding heat-insulating layer.

Detailed Description

The invention is described in detail below with reference to the following figures and specific examples:

75% of ammonium perchlorate, 6.5% of 669 epoxy resin active diluent, 0.25% of lead oxide catalyst, 3.3% -4% of 593 curing agent and the balance epoxy resin.

Example 1

The propellant is processed and manufactured into a solid rocket engine with the diameter of 82mm, actual tests are carried out, as shown in a section A-A of figure 1, the diameter of a grain is 71mm, a three-section type variable cross-section structure is adopted, the total length is 700mm, the upper end of the fuel is a circular hole, the middle section of the fuel is a circular hole variable cross-section expansion structure, as shown in a section B-B of figure 1, the lower end of the fuel is a star-shaped hole, and epoxy resin mixed fillers are adopted among three sections of grains for inflaming retarding.

As shown in figure 2, the engine adopts an aluminum shell, the spray pipe and the plug are made of bakelite, the throat liner is made of graphite, and a phenolic cloth pipe heat-insulating layer with the thickness of 2mm is arranged outside the fuel grain.

KNO is adopted in the test₃Igniting the ignition powder mixed by Mg and Al, designing the ignition pressure to be 2MPa, measuring the pressure at the plug (namely the upper end of the fuel) to obtain a stagnation pressure and thrust curve diagram shown in figure 3, wherein the time step length is 1.5625e-3s (the sampling rate is 640), and the results are obtained by observing the curve: the ignition pressure peak is low; the highest value (5.44MPa) and the average value (4.11MPa) of the pressure curve are close to theoretical prediction; the turns at each part of the curve shape are the same as the design condition, the curve changes stably without abnormal fluctuation, the test result is well matched with the design, the energy characteristic and the inner ballistic characteristic shown in the table 1 are obtained, and the test result is also well matched with the design as can be seen from the table 1.

TABLE 1 energy characteristics and internal ballistic characteristics

Example 2

The propellant is processed to prepare a solid rocket engine with the diameter of 55mm, and actual tests are carried out. As shown in the section A-A of FIG. 4, the diameter of the grain is 46mm, the total length is 440mm by adopting a four-section structure, and as shown in the section B-B of FIG. 4, the upper end of the fuel is a circular hole with the thickness of 14 mm.

As shown in figure 5, the engine adopts an aluminum shell, the spray pipe and the plug are made of bakelite, the throat liner is made of graphite, and a phenolic cloth pipe heat-insulating layer with the thickness of 2mm is arranged outside the fuel grain.

During testing, ignition is carried out by adopting ignition powder formed by mixing KNO3, Mg and Al, the ignition pressure is designed to be 2MPa, pressure measurement is carried out at a plug (namely the upper end of the fuel), a stagnation pressure and thrust curve diagram shown in figure 6 is obtained, a buffer tube and a pressure accumulation chamber are used during pressure measurement, the obtained curve is convenient for calculating the thrust, and the time step length in the diagram is 2.38e-3s (the sampling rate 420). From the graph observations: 1. the highest value (6.51MPa) and the average value (5.9MPa) of the pressure curve are close to theoretical prediction; 2. the curve is smooth and stable, no abnormal fluctuation exists, the working time is about 2.0s, and the test result is well consistent with the design.

A55 mm diameter solid rocket engine was made into a 55mm diameter rocket for flight testing. The takeoff weight of the rocket is 1.6kg, the takeoff acceleration is about 50g, the acceleration of the rocket after takeoff reaches 90g in 1.2s, and the high overload during takeoff is directly born by the fuel due to good mechanical property of the fuel without adding a supporting structure. The rocket flying condition is observed through a flying experiment to be good, and the expected flying height is achieved.

Example 3

Aluminum powder is added into the composite propellant, so that specific impact is improved, burning speed is improved, and acoustic unstable burning is inhibited; the adding amount of the aluminum powder is 5% of the mass of the composite propellant.

The preparation method of the epoxy composite propellant added with the aluminum powder comprises the following steps:

step 2: weighing and mixing the solid components except the metal combustion agent, wherein the solid components comprise ammonium perchlorate and lead oxide catalyst, and uniformly mixing for later use;

and step 3: weighing a curing agent, wherein the mass ratio content of the curing agent is within a range of 3.3-4%, the curing agent is determined according to actual curing requirements, the content of the curing agent can be properly increased to accelerate the curing speed under the conditions of simple drug type and convenient filling, and otherwise, the content is reduced;

and 4, step 4: adding aluminum powder serving as a metal combustion agent into the liquid component mixed solution, and uniformly mixing;

and 5: adding a curing agent into the liquid component obtained in the step (4) for rapid mixing for 5-10s, adding the other mixed solid components into the liquid component, uniformly stirring, wherein the uniformly mixed fuel has good fluidity, is gray, has uniform color distribution, and generates a small amount of ammonia gas in the mixing process;

step 6: and pouring the mixture into a fuel mold in vacuum, curing and molding the mixture at normal temperature, and taking the molded mixture out of the mold.

The composite propellant added with the aluminum powder with the mass percentage content of 5 percent is processed to prepare a turbine type rocket with the diameter of 37mm, and actual tests are carried out. As shown in section A-A of figure 7, the diameter of the grain is 30.5mm, the length is 100mm, a star-shaped hole structure is adopted, as shown in figure 8, the rocket engine adopts an aluminum shell, the spray pipe and the plug are both made of phenolic composite materials, the throat liner is made of stainless steel, and a polyvinyl chloride coated heat insulation layer with the thickness of 0.75mm is arranged outside the fuel grain.

The rocket is subjected to engine ground test and flight test, and KNO is adopted during the test₃And the ignition powder formed by mixing Mg and Al is ignited, and the ignition pressure is designed to be 2 MPa. The test results show that: the rocket engine works stably and has good flying condition. The takeoff acceleration of the rocket is about 90g, and the rocket can directly bear high overload during takeoff due to good mechanical property. The product is added with 5 percent of aluminum powderGray, bright flame and smoke when burning.

Through the 37mm diameter turbo-type rocket of not adding the product of aluminite powder and making of contrast test, it has obvious promotion effect to the fuel specific impulse to add the aluminite powder, but the smoke volume increases simultaneously, the flame is big, and aggravate to the ablation effect of phenolic aldehyde clad material spray tube.

While the foregoing is directed to the preferred embodiment of the present invention, other and further embodiments of the invention may be devised without departing from the basic scope thereof, and the scope thereof is determined by the claims that follow.

Claims

1. The epoxy composite propellant is characterized by comprising the following components in percentage by mass: 70-80% of ammonium perchlorate, 4-6.5% of 669 epoxy resin active diluent, 0.2-1.25% of catalyst, 3.3-4% of curing agent and the balance of epoxy resin.

2. The epoxy-based composite propellant according to claim 1, wherein the composite propellant consists of the following components in percentage by mass: 75% of ammonium perchlorate, 6.5% of 669 epoxy resin active diluent, 0.25% of catalyst, 3.3% -4% of curing agent and the balance epoxy resin.

3. The epoxy-based propellant complex of claim 1 or 2, wherein the catalyst is lead oxide.

4. The epoxy-based composite propellant according to claim 1 or 2, wherein the composite propellant is added with aluminum powder in an amount of 0-16% by mass of the composite propellant.

5. The preparation method of the epoxy composite propellant is characterized by comprising the following steps:

6. The method for preparing the epoxy-based composite propellant according to claim 5, wherein a metal combustion agent is added to the propellant, comprising the steps of:

and 4, step 4: adding the metal combustion agent into the liquid component mixed solution and uniformly mixing;

7. The method for preparing an epoxy-based composite propellant according to claim 6, wherein the metal combustion agent is aluminum powder.

8. The application of the epoxy composite propellant is characterized in that the epoxy composite propellant is applied to a solid rocket engine.

9. Use of an epoxy-based propellant complex according to claim 8 in small and medium solid rocket engines.