CN114479431A - Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof - Google Patents

Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof Download PDF

Info

Publication number
CN114479431A
CN114479431A CN202111564616.8A CN202111564616A CN114479431A CN 114479431 A CN114479431 A CN 114479431A CN 202111564616 A CN202111564616 A CN 202111564616A CN 114479431 A CN114479431 A CN 114479431A
Authority
CN
China
Prior art keywords
retardant layer
flame
parts
ablatable
mass
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
CN202111564616.8A
Other languages
Chinese (zh)
Inventor
吴晨颖
吴战武
沈兆宏
詹国柱
宋洪标
翟杰明
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Shanghai Aerospace Chemical Engineering Institute
Original Assignee
Shanghai Aerospace Chemical Engineering Institute
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Shanghai Aerospace Chemical Engineering Institute filed Critical Shanghai Aerospace Chemical Engineering Institute
Priority to CN202111564616.8A priority Critical patent/CN114479431A/en
Publication of CN114479431A publication Critical patent/CN114479431A/en
Pending legal-status Critical Current

Links

Images

Classifications

    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08L—COMPOSITIONS OF MACROMOLECULAR COMPOUNDS
    • C08L75/00—Compositions of polyureas or polyurethanes; Compositions of derivatives of such polymers
    • C08L75/04—Polyurethanes
    • C08L75/14—Polyurethanes having carbon-to-carbon unsaturated bonds
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08G—MACROMOLECULAR COMPOUNDS OBTAINED OTHERWISE THAN BY REACTIONS ONLY INVOLVING UNSATURATED CARBON-TO-CARBON BONDS
    • C08G18/00—Polymeric products of isocyanates or isothiocyanates
    • C08G18/06—Polymeric products of isocyanates or isothiocyanates with compounds having active hydrogen
    • C08G18/28—Polymeric products of isocyanates or isothiocyanates with compounds having active hydrogen characterised by the compounds used containing active hydrogen
    • C08G18/67—Unsaturated compounds having active hydrogen
    • C08G18/69—Polymers of conjugated dienes
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08G—MACROMOLECULAR COMPOUNDS OBTAINED OTHERWISE THAN BY REACTIONS ONLY INVOLVING UNSATURATED CARBON-TO-CARBON BONDS
    • C08G18/00—Polymeric products of isocyanates or isothiocyanates
    • C08G18/06—Polymeric products of isocyanates or isothiocyanates with compounds having active hydrogen
    • C08G18/70—Polymeric products of isocyanates or isothiocyanates with compounds having active hydrogen characterised by the isocyanates or isothiocyanates used
    • C08G18/72—Polyisocyanates or polyisothiocyanates
    • C08G18/74—Polyisocyanates or polyisothiocyanates cyclic
    • C08G18/75—Polyisocyanates or polyisothiocyanates cyclic cycloaliphatic
    • C08G18/751—Polyisocyanates or polyisothiocyanates cyclic cycloaliphatic containing only one cycloaliphatic ring
    • C08G18/752—Polyisocyanates or polyisothiocyanates cyclic cycloaliphatic containing only one cycloaliphatic ring containing at least one isocyanate or isothiocyanate group linked to the cycloaliphatic ring by means of an aliphatic group
    • C08G18/753—Polyisocyanates or polyisothiocyanates cyclic cycloaliphatic containing only one cycloaliphatic ring containing at least one isocyanate or isothiocyanate group linked to the cycloaliphatic ring by means of an aliphatic group containing one isocyanate or isothiocyanate group linked to the cycloaliphatic ring by means of an aliphatic group having a primary carbon atom next to the isocyanate or isothiocyanate group
    • C08G18/755—Polyisocyanates or polyisothiocyanates cyclic cycloaliphatic containing only one cycloaliphatic ring containing at least one isocyanate or isothiocyanate group linked to the cycloaliphatic ring by means of an aliphatic group containing one isocyanate or isothiocyanate group linked to the cycloaliphatic ring by means of an aliphatic group having a primary carbon atom next to the isocyanate or isothiocyanate group and at least one isocyanate or isothiocyanate group linked to a secondary carbon atom of the cycloaliphatic ring, e.g. isophorone diisocyanate
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08K—Use of inorganic or non-macromolecular organic substances as compounding ingredients
    • C08K3/00—Use of inorganic substances as compounding ingredients
    • C08K3/30—Sulfur-, selenium- or tellurium-containing compounds
    • C08K2003/3045—Sulfates
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08L—COMPOSITIONS OF MACROMOLECULAR COMPOUNDS
    • C08L2201/00—Properties
    • C08L2201/02—Flame or fire retardant/resistant

Abstract

The invention relates to an ablatable flame-retardant layer of a solid rocket engine, a preparation method and application thereof, wherein the ablatable flame-retardant layer comprises the following components in parts by weight: 35-45 parts of hydroxyl-terminated polybutadiene, 6-8 parts of bis (2-ethylhexyl) sebacate, 0.025-0.05 g of dimethyl silicone oil, 2-3 parts of chain extender, 0.15-0.2 part of curing assistant, 40-55 parts of high-heat-sink filler, 5-8 parts of curing agent and 0.6-0.8 part of dibutyl phthalate solution; wherein the dibutyl phthalate solution contains 2.0-2.5 percent of ferric acetylacetonate; the flame-retardant layer is arranged between the primary propellant and the secondary propellant, has flame-retardant and ablation effects, realizes the heat absorption and ablation loss of the flame-retardant layer by igniting the primary propellant in the solid rocket engine, protects and does not influence the ignition of the secondary propellant, and realizes the grading independent work among the multi-stage propellants.

Description

Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof
Technical Field
The invention relates to a flame-retardant layer material and a preparation method thereof, in particular to an ablatable flame-retardant layer of a solid rocket engine, a preparation method and application thereof, and belongs to the technical field of thermal protection of solid rocket engines.
Background
In the background of typical air-defense missile application, in order to meet the requirements of fine control and good combat efficiency of missile weapons, a solid rocket engine is required to have multiple start-stop capabilities, a mode of independently packaging thrust units by a plurality of pulses is generally adopted at present, each thrust unit is provided with an independent combustion chamber filled with propellant and an independent signal control system, and rocket engines with the structures have the limitations of large structural size, high manufacturing cost, complex control system and the like, so the application of the solid rocket engine is limited.
The propellant in solid rocket engines is a solid propellant and the composition contains an oxidizer that provides oxygen for self-sustaining combustion and, once combustion conditions are established, combustion continues until the propellant is burned off. Therefore, in order to realize the starting and stopping of the propellant of the solid rocket engine, a redundant structure of a multi-pulse independently packaged thrust unit is not adopted, a partition material is required to be used, the combustion of the propellant is purposefully retarded, and the next ignition starting cannot be influenced.
Disclosure of Invention
The invention aims to provide an ablative flame-retardant layer of a solid rocket engine, which is arranged between a primary propellant and a secondary propellant in the same combustion chamber and has the functions of flame retardance and ablation, the ignition of the primary propellant in the solid rocket engine realizes the heat absorption ablation loss of the flame-retardant layer, the protection does not influence the ignition of the secondary propellant, and the grading independent work among the multi-stage propellants is realized.
The invention also aims to provide a preparation method and application of the ablatable flame-retardant layer of the solid rocket engine.
The above purpose of the invention is mainly realized by the following technical scheme:
the solid rocket engine ablatable flame-retardant layer comprises the following raw materials in parts by mass:
35-45 parts of hydroxyl-terminated polybutadiene, 6-8 parts of bis (2-ethylhexyl) sebacate, 0.025-0.05 g of simethicone, 2-3 parts of chain extender, 0.15-0.2 part of curing assistant, 40-55 parts of high-heat-sink filler, 5-8 parts of curing agent and 0.6-0.8 part of dibutyl phthalate solution;
wherein the dibutyl phthalate solution contains 2.0-2.5 percent of ferric acetylacetonate.
In the solid rocket motor ablatable flame-retardant layer, the chain extender is any one of N, N' -bis (2-hydroxypropyl) aniline, ethylene glycol, diethylene glycol or triethylene glycol.
In the ablative flame-retardant layer of the solid rocket engine, the curing auxiliary agent is any one of polyamide resin or liquid paraffin.
In the ablative flame-retardant layer of the solid rocket engine, the high heat sink filler is any one of ammonium sulfate, paraformaldehyde or melamine.
In the ablative flame-retardant layer of the solid rocket motor, the curing agent is any one of isophorone diisocyanate or toluene diisocyanate.
In the ablative flame-retardant layer of the solid rocket engine, the thickness of the flame-retardant layer is 1-2 mm.
A preparation method of an ablatable flame-retardant layer of a solid rocket engine comprises the following steps:
(1) weighing ferric acetylacetonate and dibutyl phthalate to prepare a solution with the ferric acetylacetonate content of 2.0-2.5% by mass for later use;
(2) placing the chain extender and the curing auxiliary agent in an oven at 58-62 ℃ for storage for 12-24 h for later use;
(3) placing the high-heat-deposition filler in a drying box for storage for 6-8 days for later use;
(4) weighing 35-45 parts by mass of hydroxyl-terminated polybutadiene, and adding into a container;
(5) weighing 6-8 parts by mass of bis (2-ethylhexyl) sebacate, adding into a container, adding 1-2 drops of simethicone by using a rubber head dropper, namely 0.025 g-0.05 part by mass, and stirring for 3-5 min;
(6) taking out the chain extender in the oven, weighing 2-3 parts by mass, adding into a container, and stirring for 3-5 min;
(7) taking out the curing assistant in the oven, weighing 0.15-0.2 parts by mass, adding into a container, and rapidly stirring for 3-5 min until no insoluble substance remains;
(8) taking out the high-heat-deposition filler in the drying oven, grinding for 1-2 min, weighing 40-55 parts by mass, adding into a container, and slowly stirring for 5-10 min until the powder is completely dissolved;
(9) weighing 5-8 parts by mass of a curing agent, adding into a container, and stirring for 3-5 min;
(10) weighing 0.6-0.8 part by mass of dibutyl phthalate solution with the mass percentage of ferric acetylacetonate of 2.0-2.5%, adding the dibutyl phthalate solution into a container, and stirring for 7-8 min;
(11) placing the slurry stirred in the containers in the steps (4) to (10) into a vacuum pump, and vacuum degassing for 20-40 min;
(12) and after degassing, pouring the slurry into a film mold, scraping the slurry flatly, sealing, and then placing in a 58-62 ℃ oven for curing for 160-176 h to finish the preparation of the flame-retardant layer film.
In the preparation method of the solid rocket engine ablative flame retardant layer, the chain extender is any one of N, N' -bis (2-hydroxypropyl) aniline, ethylene glycol, diethylene glycol or triethylene glycol; the curing auxiliary agent is any one of polyamide resin or liquid paraffin; the high heat sink filler is any one of ammonium sulfate, paraformaldehyde or melamine; the curing agent is any one of isophorone diisocyanate or toluene diisocyanate.
In the preparation method of the ablatable flame-retardant layer of the solid rocket motor, the film mould in the step (12) is made of aluminum, and the coating material on the surface of the film mould is polytetrafluoroethylene.
In the preparation method of the ablative flame-retardant layer of the solid rocket engine, the thickness of the prepared flame-retardant layer rubber sheet is 1-2 mm.
The application of the ablative flame-retardant layer of the solid rocket engine is that the flame-retardant layer is arranged between the primary propellant and the secondary propellant, plays a flame-retardant role, and absorbs heat to ablate after the primary propellant ignites.
Compared with the prior art, the invention has the following beneficial effects:
(1) the solid rocket engine ablatable flame-retardant layer material provided by the invention is prepared by taking hydroxyl-terminated polybutadiene and isophorone diisocyanate curing system as matrix materials, adding high-heat-sinking filler, arranging the flame-retardant layer between the primary propellant and the secondary propellant, having flame-retardant and ablation effects, effectively realizing heat absorption ablation loss after the ignition work of the primary propellant of the flame-retardant layer is finished, and protecting the secondary propellant from being influenced.
(2) The obtained flame-retardant layer can meet the requirement of realizing heat absorption ablation loss of the flame-retardant layer after ignition of a first-stage propellant in the staged ignition of the solid rocket engine, protects the condition of not influencing ignition of a second-stage propellant, completes staged independent work among multiple stages of propellants, replaces the condition of adopting two-stage thrust chambers in the prior art, simplifies the structure of the solid rocket engine, reduces the cost and improves the efficiency.
(3) The tests of the invention show that the peak value and the working time of the primary thrust curve and the secondary thrust curve are consistent with the calculation prediction result.
Drawings
FIG. 1 is a schematic view of the structure of a primary propellant, a flame retardant layer and a secondary propellant according to the present invention;
FIG. 2 is a thermal analysis curve of a sample in example 1 of the present invention;
FIG. 3 is a thermal analysis curve of a sample in example 2 of the present invention;
FIG. 4 is a drawing showing a sample of a flame retardant layer in example 1 of the present invention;
FIG. 5 is a drawing showing a sample of a flame-retardant layer in example 2 of the present invention;
fig. 6 is a graph comparing the test results and the predicted results of the ablative flame retardant layer of example 1 of the present invention in a two-stage pulse engine, where a is the pressure-time curve and b is the thrust-time curve.
The specific implementation mode is as follows:
the invention is described in further detail below with reference to the following figures and specific examples:
the invention prepares the flame-retardant layer partition among the multi-level propellants, and has the following characteristics according to the application requirement of the flame-retardant layer in the secondary ignition pulse engine: the heat insulation layer material realizes ablation after primary ignition without residues such as carbon deposition and the like, otherwise secondary ignition is influenced; the heat insulating layer material has heat absorbing properties that would otherwise ignite a secondary propellant; the heat-insulating layer material should have high long-term storage stability and good processing formability.
As shown in fig. 1, the structure of the primary propellant, the flame retardant layer and the secondary propellant is schematically shown, the flame retardant layer needs to absorb heat to complete ablation and weakening after the primary propellant finishes working, so that the normal working of the secondary propellant is not influenced, and the grading independent working among multiple stages of propellants is realized.
The raw materials of the solid rocket engine ablative flame-retardant layer comprise the following components in parts by mass:
35-45 parts of hydroxyl-terminated polybutadiene, 6-8 parts of bis (2-ethylhexyl) sebacate, 1-2 drops (0.025 g-0.05 part) of a dimethicone rubber head dropper, 2-3 parts of a chain extender, 0.15-0.2 part of a curing assistant, 40-55 parts of a high-heat-sink filler, 5-8 parts of a curing agent and 0.6-0.8 part of a dibutyl phthalate solution;
wherein the dibutyl phthalate solution contains 2.0-2.5 percent of ferric acetylacetonate.
In an alternative embodiment, the chain extender is any one of N, N' -bis (2-hydroxypropyl) aniline, ethylene glycol, diethylene glycol, or triethylene glycol.
In an alternative embodiment, the curing assistant is any one of polyamide resin or liquid paraffin.
In an alternative embodiment, the high heat-sinking filler is any one of ammonium sulfate, paraformaldehyde or melamine.
In an alternative embodiment, the curing agent is any one of isophorone diisocyanate or toluene diisocyanate.
In an optional embodiment, the thickness of the flame retardant layer is 1 to 2 mm.
The preparation method of the solid rocket engine ablatable flame-retardant layer comprises the following steps:
(1) weighing ferric acetylacetonate and dibutyl phthalate to prepare a solution with the ferric acetylacetonate content of 2.0-2.5% by mass for later use;
(2) placing the chain extender and the curing assistant in a drying oven at 60 +/-2 ℃ for storage for 12-24 hours for later use;
(3) placing the high-heat-deposition filler in a drying box for storage for 6-8 days for later use;
(4) weighing 35-45 parts by mass of hydroxyl-terminated polybutadiene, and adding into a container;
(5) weighing 6-8 parts by mass of bis (2-ethylhexyl) sebacate, adding into a container, adding 1-2 drops (0.025 g-0.05 part by mass) of simethicone by using a rubber head dropper, and stirring for 3-5 min;
(6) taking out the chain extender in the oven, weighing 2-3 parts by mass, adding into a container, and stirring for 3-5 min until the color is uniform;
(7) taking out the curing assistant in the oven, weighing 0.15-0.2 parts by mass, adding into a container, and rapidly stirring for 3-5 min until no insoluble substances remain and the color is uniform;
(8) taking out the high-heat-deposition filler in the drying oven, grinding for 1-2 min, weighing 40-55 parts by mass, adding into a container, and slowly stirring for 5-10 min until the powder is completely dissolved and the color is uniform;
(9) weighing 5-8 parts by mass of a curing agent, adding into a container, and stirring for 3-5 min;
(10) weighing 0.6-0.8 part by mass of dibutyl phthalate solution with the mass percentage content of ferric acetylacetonate of 2.0-2.5%, adding the dibutyl phthalate solution into a container, and stirring for 7-8 min until the color is uniform;
(11) placing the slurry stirred in the containers in the steps (4) to (10) into a vacuum pump, and vacuum degassing for 20-40 min;
(12) and after degassing, pouring the slurry into a film mold, scraping the slurry flatly by using a scraper, sealing, and then placing the sealed slurry in a baking oven at 58-62 ℃ for curing for 160-176 h to finish the preparation of the flame-retardant layer film.
In an optional embodiment, the film mold is made of aluminum, the coating material on the surface of the film mold is polytetrafluoroethylene, and the thickness of the flame-retardant layer film is 1 mm.
The chain extender is any one of N, N' -bis (2-hydroxypropyl) aniline, ethylene glycol, diethylene glycol or triethylene glycol; the curing auxiliary agent is any one of polyamide resin or liquid paraffin; the high heat sink filler is any one of ammonium sulfate, paraformaldehyde or melamine; the curing agent is any one of isophorone diisocyanate or toluene diisocyanate.
The flame-retardant layer film prepared by the formula of the ablatable flame-retardant layer material provided by the invention can meet the requirement of realizing the heat absorption ablation loss of the flame-retardant layer after the ignition of the primary propellant in the staged ignition of the solid rocket engine, protect and not influence the ignition of the secondary propellant, and complete the staged independent work among the multi-stage propellant.
Example 1
An ablatable flame-retardant layer material of a solid rocket engine is composed of the following raw materials: 35.5g of hydroxyl-terminated polybutadiene, 6g of bis (2-ethylhexyl) sebacate, 1 drop (0.025g) of a dimethicone rubber head dropper, 2g of N, N' -bis (2-hydroxypropyl) aniline, 0.15g of polyamide resin, 50g of ammonium sulfate, 5.8g of isophorone diisocyanate, and 0.7g of dibutyl phthalate solution with 2.2% of iron acetylacetonate.
A preparation method of an ablatable flame-retardant layer material of a solid rocket engine comprises the following steps:
(1) weighing ferric acetylacetonate and dibutyl phthalate to prepare a 2.2% solution for later use;
(2) placing N, N' -bis (2-hydroxypropyl) aniline and polyamide resin in a 60 ℃ oven for storage for 24 hours for later use;
(3) placing the high heat-sinking filler in a drying box for storage for 6 days for later use;
(4) weighing 35.5g of hydroxyl-terminated polybutadiene, and adding into a preparation container;
(5) weighing 6g of sebacic acid di (2-ethylhexyl) ester, adding into a container, adding 1 drop of simethicone by using a rubber head dropper, and stirring for 3 min;
(6) taking out the N, N' -bis (2-hydroxypropyl) aniline in the oven, weighing 2g of aniline, adding the weighed aniline into a container, and stirring for 5min until the color is uniform;
(7) taking out the polyamide resin in the oven, weighing 0.15g, adding into a container, and rapidly stirring for 5min until no insoluble substances remain and the color is uniform;
(8) taking out the ammonium sulfate in the drying oven, grinding for 2min, weighing 50g, adding into a container, and slowly stirring for 5-10 min until the powder is completely dissolved and the color is uniform;
(9) weighing 5.8g of isophorone diisocyanate, adding into a container, and stirring for 4 min;
(10) weighing 0.7g of dibutyl phthalate solution with the ferric acetylacetonate content of 2.2%, adding the solution into a container, and stirring for 7min until the color is uniform;
(11) placing the stirred slurry into a vacuum pump, and vacuum degassing for 20 min;
(12) after degassing is finished, pouring the slurry into a polytetrafluoroethylene square film die, scraping the slurry flatly by using a scraper, sealing and flatly placing the slurry in a 60 ℃ oven for curing for 168 hours, and manufacturing a flame-retardant layer film and a test piece.
The main test results of the ablatable flame retardant layer material prepared in example 1 are as follows: fig. 2 is a thermogravimetric analysis curve of a sample of example 1 of the present invention, table 1 shows the main performance test results, and fig. 4 is a macroscopic digital photograph of the sample of example 1 of the present invention. Fig. 6 is a graph comparing the test results and the predicted results of the ablative flame retardant layer of example 1 of the present invention in a two-stage pulse engine, where a is the pressure versus time curve and b is the thrust versus time curve.
The results in fig. 6 show that the thrust and pressure rapidly drop to 0 after the end of the primary propellant operation and that a subsequent re-ignition is successfully achieved after a predetermined 20 s. The test pressure and thrust curves are consistent with the prediction curve, and the working time of the first propellant and the second propellant is consistent.
Table 1 example 1 sample main test results
Figure BDA0003421422790000091
Example 2
An ablatable flame-retardant layer material of a solid rocket engine is composed of the following raw materials: 42.4g of hydroxyl-terminated polybutadiene, 7.2g of a bis (2-ethylhexyl) sebacate solution, 1 drop of a dimethicone rubber head dropper, 2.3g of N, N' -bis (2-hydroxypropyl) aniline, 0.17g of polyamide resin, 40g of paraformaldehyde, 7g of isophorone diisocyanate and 0.8g of a dibutyl phthalate solution with the content of iron acetylacetonate of 2.2%.
A preparation method of an ablatable flame-retardant layer material of a solid rocket engine comprises the following steps:
(1) weighing ferric acetylacetonate and dibutyl phthalate to prepare a 2.2% solution for later use;
(2) placing N, N' -bis (2-hydroxypropyl) aniline and polyamide resin in a 60 ℃ oven for storage for 24 hours for later use;
(3) placing the high heat-sinking filler in a drying box for storage for 8 days for later use;
(4) weighing 42.4g of hydroxyl-terminated polybutadiene in a container;
(5) weighing 7.2g of sebacic acid di (2-ethylhexyl) ester, adding into a container, adding 1 drop of simethicone by using a rubber head dropper, and stirring for 3 min;
(6) taking out the N, N' -bis (2-hydroxypropyl) aniline in the oven, weighing 2.3g, adding into a container, and stirring for 4min until the color is uniform;
(7) taking out the polyamide resin in the oven, weighing 0.17g, adding into a container, and rapidly stirring for 5min until no insoluble substances remain and the color is uniform;
(8) taking out paraformaldehyde in a drying oven, grinding for 2min, weighing 40g, adding into a container, and slowly stirring for 10min until the powder is completely dissolved and the color is uniform;
(9) weighing 7g of isophorone diisocyanate, adding into a container, and stirring for 5 min;
(10) weighing 0.8g of dibutyl phthalate solution with the ferric acetylacetonate content of 2.2%, adding into a container, and stirring for 8min until the color is uniform;
(11) placing the stirred slurry into a vacuum pump, and vacuum degassing for 30 min;
(12) after degassing is finished, pouring the slurry into a polytetrafluoroethylene square film die, scraping the slurry flatly by using a scraper, sealing and flatly placing the slurry in a 60 ℃ oven for curing for 168 hours, and manufacturing a flame-retardant layer film and a test piece.
The main test results of the ablatable flame retardant layer material prepared in example 2 are as follows: FIG. 3 is a thermogravimetric analysis curve of the sample of example 2, Table 2 shows the main performance test results, and FIG. 5 is a macroscopic digital photograph of the sample of example 2.
Table 2 example 2 main test results of samples
Test item Test value
Tensile Strength (kPa) 405
Elongation at Break (%) 68
Density (g/cm)3) 0.89
Thermal decomposition temperature T5、Tmax(℃) T5＝108℃，Tmax1＝145℃，Tmax2＝465℃
The above description is only for the best mode of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.

Claims (11)

1. The solid rocket engine ablatable flame-retardant layer is characterized by comprising the following components in parts by mass:
35-45 parts of hydroxyl-terminated polybutadiene, 6-8 parts of bis (2-ethylhexyl) sebacate, 0.025-0.05 g of simethicone, 2-3 parts of chain extender, 0.15-0.2 part of curing assistant, 40-55 parts of high-heat-sink filler, 5-8 parts of curing agent and 0.6-0.8 part of dibutyl phthalate solution;
wherein the dibutyl phthalate solution contains 2.0-2.5 percent of ferric acetylacetonate.
2. The solid rocket motor ablatable flame retardant layer of claim 1, wherein the chain extender is any one of N, N' -bis (2-hydroxypropyl) aniline, ethylene glycol, diethylene glycol, or triethylene glycol.
3. The solid-rocket motor ablatable flame retardant layer of claim 1, wherein the curing aid is any one of polyamide resin or liquid paraffin.
4. The solid-rocket motor ablatable flame retardant layer of claim 1, wherein the high heat sink filler is any one of ammonium sulfate, paraformaldehyde, or melamine.
5. The solid-rocket motor ablatable flame retardant layer of claim 1, wherein the curing agent is any one of isophorone diisocyanate or toluene diisocyanate.
6. The solid-rocket motor ablatable flame retardant layer of claim 1, wherein the flame retardant layer has a thickness of 1-2 mm.
7. A preparation method of an ablatable flame-retardant layer of a solid rocket engine is characterized by comprising the following steps:
(1) weighing ferric acetylacetonate and dibutyl phthalate to prepare a solution with the ferric acetylacetonate content of 2.0-2.5% by mass for later use;
(2) placing the chain extender and the curing auxiliary agent in an oven at 58-62 ℃ for storage for 12-24 h for later use;
(3) placing the high-heat-deposition filler in a drying box for storage for 6-8 days for later use;
(4) weighing 35-45 parts by mass of hydroxyl-terminated polybutadiene, and adding into a container;
(5) weighing 6-8 parts by mass of bis (2-ethylhexyl) sebacate, adding into a container, adding 1-2 drops of simethicone by using a rubber head dropper, namely 0.025 g-0.05 part by mass, and stirring for 3-5 min;
(6) taking out the chain extender in the oven, weighing 2-3 parts by mass, adding into a container, and stirring for 3-5 min;
(7) taking out the curing assistant in the oven, weighing 0.15-0.2 parts by mass, adding into a container, and rapidly stirring for 3-5 min until no insoluble substance remains;
(8) taking out the high-heat-deposition filler in the drying oven, grinding for 1-2 min, weighing 40-55 parts by mass, adding into a container, and slowly stirring for 5-10 min until the powder is completely dissolved;
(9) weighing 5-8 parts by mass of a curing agent, adding into a container, and stirring for 3-5 min;
(10) weighing 0.6-0.8 part by mass of dibutyl phthalate solution with the mass percentage of ferric acetylacetonate of 2.0-2.5%, adding the dibutyl phthalate solution into a container, and stirring for 7-8 min;
(11) placing the slurry stirred in the containers in the steps (4) to (10) into a vacuum pump, and vacuum degassing for 20-40 min;
(12) and after degassing, pouring the slurry into a film mold, scraping the slurry flatly, sealing, and then placing in a 58-62 ℃ oven for curing for 160-176 h to finish the preparation of the flame-retardant layer film.
8. The method for preparing an ablatable, flame-retardant layer of a solid rocket engine as recited in claim 7, wherein said chain extender is any one of N, N' -bis (2-hydroxypropyl) aniline, ethylene glycol, diethylene glycol or triethylene glycol; the curing auxiliary agent is any one of polyamide resin or liquid paraffin; the high heat sink filler is any one of ammonium sulfate, paraformaldehyde or melamine; the curing agent is any one of isophorone diisocyanate or toluene diisocyanate.
9. The method for preparing the ablatable flame retardant layer of the solid rocket motor according to claim 7, wherein the film mold in step (12) is made of aluminum and the coating material on the surface of the film mold is polytetrafluoroethylene.
10. The method for preparing the ablatable flame-retardant layer of the solid rocket engine according to claim 7, wherein the thickness of the prepared flame-retardant layer rubber sheet is 1-2 mm.
11. Use of the ablatable, flame retardant layer of a solid-rocket engine as recited in any one of claims 1-6, wherein said ablatable, flame retardant layer is disposed between a primary propellant and a secondary propellant, and is configured to perform flame retardant action and to absorb heat to ablate upon ignition of the primary propellant.
CN202111564616.8A 2021-12-20 2021-12-20 Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof Pending CN114479431A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202111564616.8A CN114479431A (en) 2021-12-20 2021-12-20 Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202111564616.8A CN114479431A (en) 2021-12-20 2021-12-20 Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof

Publications (1)

Publication Number Publication Date
CN114479431A true CN114479431A (en) 2022-05-13

Family

ID=81494000

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202111564616.8A Pending CN114479431A (en) 2021-12-20 2021-12-20 Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof

Country Status (1)

Country Link
CN (1) CN114479431A (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115784815A (en) * 2022-11-21 2023-03-14 北京理工大学 Hydroxyl-terminated polybutadiene flame-retardant material and preparation method and application thereof

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20040209987A1 (en) * 2002-06-26 2004-10-21 Gajiwala Himansu M. Low-cost, low-density, ablative rubber insulation for rocket motors
CN110982475A (en) * 2019-11-19 2020-04-10 上海航天化工应用研究所 Non-precured butylated hydroxyl lining and preparation method thereof
CN112853770A (en) * 2021-01-08 2021-05-28 上海毅榕电力技术服务有限公司 Fire-resistant flame-retardant heat-insulating material, fire-resistant flame-retardant heat-insulating explosion-proof cloth and preparation method

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20040209987A1 (en) * 2002-06-26 2004-10-21 Gajiwala Himansu M. Low-cost, low-density, ablative rubber insulation for rocket motors
CN110982475A (en) * 2019-11-19 2020-04-10 上海航天化工应用研究所 Non-precured butylated hydroxyl lining and preparation method thereof
CN112853770A (en) * 2021-01-08 2021-05-28 上海毅榕电力技术服务有限公司 Fire-resistant flame-retardant heat-insulating material, fire-resistant flame-retardant heat-insulating explosion-proof cloth and preparation method

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
毕大鹏: "基于互穿网络的消融防热涂料", 《工程科技Ⅰ辑》, no. 2, pages 5 - 8 *
马天信;: "有机消融涂层防热隔热机理及组成分析", 航天制造技术, no. 02, pages 10 - 13 *

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115784815A (en) * 2022-11-21 2023-03-14 北京理工大学 Hydroxyl-terminated polybutadiene flame-retardant material and preparation method and application thereof
CN115784815B (en) * 2022-11-21 2024-01-30 北京理工大学 Butyloxapol flame retardant material and preparation method and application thereof

Similar Documents

Publication Publication Date Title
US6059906A (en) Methods for preparing age-stabilized propellant compositions
KR101334732B1 (en) Insensitive propellant composition
CN109305869B (en) Carborane propellant and preparation method thereof
CN111592436A (en) Low-vulnerability propellant and preparation method thereof
CN110156548B (en) Casting type high-energy insensitive explosive and preparation method thereof
CN114479431A (en) Meltable flame-retardant layer of solid rocket engine, preparation method and application thereof
DE3207811A1 (en) Castable magnesium fuel having a base of silicon
CN110423184B (en) Propellant for solid rocket engine and preparation method thereof
KR101557925B1 (en) Solid Propellant Compositions of Nitrate Ester Polyether Family in Moderate Temperature
US6726788B2 (en) Preparation of strengthened ammonium nitrate propellants
KR101664236B1 (en) Propellant for ducted rocket
CN110981649B (en) Coating layer suitable for testing burning rate of low-burning-rate propellant and preparation method thereof
CA2273335A1 (en) Ammonium nitrate propellants with molecular sieve
CN114196454B (en) Solid fuel containing high nitrogen compound
CN107512997B (en) Low-burning-rate high-energy solid propellant
US4116131A (en) Solid propellant motor
US3215028A (en) Means and method for restricting a solid propellant
CN112608205A (en) Hydroxyl-terminated polyethylene oxide and propylene oxide random co-polyether type propellant and preparation method thereof
US3170283A (en) Compacted hydrazine bisborane fuel and method of operating gas generators
US2938778A (en) Ammonium nitrate gas-generating composition
US5192379A (en) Densifying and stabilizing ingredient
US3791140A (en) Operation of hybrid rocket systems
US3077734A (en) Solid-propellant rocket motor
CN115849999B (en) Hydrocarbon fuel-rich propellant and preparation method thereof
US2973256A (en) Ammonium nitrate solid composite propellant composition

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination