High-pressure low-temperature propellant combustion experiment system and experiment method

Abstract

The invention discloses a high-pressure low-temperature propellant combustion experiment system which comprises a combustion chamber arranged in a low-temperature box and a fiber optic spectrometer connected with the combustion chamber and a computer. The high-speed camera directly faces the combustion chamber and is connected with the computer; the signal collector is respectively connected with the computer and a thermocouple arranged in the propellant, and is used for collecting a voltage signal of the thermocouple in the combustion process of the propellant and transmitting the voltage signal to the computer; the high-pressure transportation system is connected with the combustion chamber and used for providing a high-pressure environment for the combustion chamber; and the ignition device is connected with the propellant through an ignition wire and used for igniting the propellant. The invention further discloses an experiment method of the high-pressure low-temperature propellant combustion experiment system and a method for changing the performance of an aluminum-containing solid composite propellant to enable the aluminum-containing solid composite propellant to be suitable for low temperature. The problem that a propellant combustion experiment is difficult to carry out under the combined action of low temperature and high pressure in the prior art is solved.

Description

Technical Field

[0001] The present invention belongs to the technical field of solid rocket motor propellants, and relates to a high-pressure and low-temperature propellant combustion experiment system and an experimental method. Background Art

[0002] In some cases, an engine must operate in a high-pressure and low-temperature environment. Such an environment not only requires that solid propellants meet safety requirements during storage, transportation, etc., but also requires that solid propellants have good combustion characteristics under high-pressure and low-temperature conditions.

[0003] Traditional composite solid propellant systems are difficult to achieve long-term storage and reliable operation in ultra-low temperature environments. The structural changes of the propellant caused by the decrease in environmental temperature will also affect the ignition characteristics and combustion characteristics of the propellant, affect the combustion rate of the propellant, and reduce the combustion efficiency of the propellant. Under low-temperature conditions, during the combustion process of solid propellants, they will bear huge stresses, and problems such as cracks, breakage, and deformation may occur, resulting in structural damage and causing serious accidents such as unstable combustion and combustion-to-detonation transition. At the same time, under high-pressure environments, the gas-phase reaction during propellant combustion accelerates, the thermal feedback enhances, and the evaporation rate of the surface molten layer increases, jointly promoting the increase in the combustion rate, but may trigger combustion instability (such as acoustic vibration or pressure oscillation) or erosive combustion phenomena.

[0004] As the power source in the propulsion system, in-depth study of the ignition characteristics and combustion characteristics of solid propellants under high-pressure and low-temperature conditions has important guiding significance for evaluating the performance of propellants under high-pressure and low-temperature conditions, and can also provide theoretical guidance and research directions for improving the comprehensive performance of propellants under high-pressure and low-temperature conditions.

[0005] When conducting high-pressure experiments on solid propellants in a low-temperature environment, the experimental device needs to withstand extreme low temperature and high-pressure combustion transient shocks simultaneously, posing extremely high requirements for the compatibility of sealing materials and sensors. These factors make the experimental design need to take into account data reliability and experimental safety. However, there is no complete record of relevant combustion experiment systems under high-pressure and low-temperature environments in the existing technology to solve this problem. Summary of the Invention

[0006] In order to achieve the above object, the present invention provides a high-pressure and low-temperature propellant combustion experiment system and an experimental method, which solve the problem in the existing technology that it is difficult to conduct propellant combustion experiments under the combined action of low temperature and high pressure.

[0007] To solve the above technical problems, the technical solution adopted by the present invention is a high-pressure and low-temperature propellant combustion experiment system for studying methods to change the performance of existing aluminum-containing solid composite propellants to make them suitable for low temperatures. The experimental system includes:

[0008] The combustion chamber is disposed inside the cryogenic tank. The combustion chamber is used to install the propellant and provide a combustion space for the propellant. The cryogenic tank is used to provide a controllable low-temperature environment for the combustion chamber, and the operating temperature range of the cryogenic tank is -80°C to 0°C;

[0009] The fiber optic spectrometer is connected to the combustion chamber and the computer respectively. The fiber optic spectrometer is used to record the combustion intensity of the propellant and transmit relevant data to the computer;

[0010] The high-speed camera is disposed facing the combustion chamber and connected to the computer. It is used to capture the combustion image of the propellant, record the combustion process of the propellant, and transmit the combustion image to the computer;

[0011] The signal collector is connected to the computer and the thermocouple disposed inside the propellant respectively. It is used to collect the voltage signal of the thermocouple during the combustion process of the propellant and transmit the voltage signal to the computer;

[0012] The high-pressure transportation system is connected to the combustion chamber and is used to provide a high-pressure environment of 3 MPa to 15 MPa for the combustion chamber;

[0013] The ignition device is connected to the propellant through an ignition wire and is used to ignite the propellant.

[0014] Furthermore, it further includes a synchronization trigger. The synchronization trigger is connected to the ignition device and the fiber optic spectrometer respectively, and is used to achieve the synchronization of the ignition device turning on and the fiber optic spectrometer recording data.

[0015] Furthermore, the combustion chamber includes:

[0016] The combustion chamber main body has a structure with an internal space and openings at the upper and lower ends. The combustion chamber main body is provided with an observation window; the internal space of the combustion chamber main body is used to accommodate the propellant;

[0017] The upper end cover is fixed at the upper opening of the combustion chamber main body; the propellant is fixed to the lower surface of the upper end cover;

[0018] The lower end cover is fixed at the lower opening of the combustion chamber main body;

[0019] The sealing ring is located at the connection between the observation window and the combustion chamber main body, the connection between the upper end cover and the combustion chamber main body, and the connection between the lower end cover and the combustion chamber main body; the sealing ring is used to ensure the sealing of the combustion chamber in an environment of at least -80°C;

[0020] The combustion chamber main body, the upper end cover, the lower end cover, and the sealing ring form a complete sealed space.

[0021] Furthermore, the high-pressure transportation system includes a high-pressure gas cylinder, an intake pipe, an exhaust pipe, and a pressure sensor. Among them, the high-pressure gas cylinder is connected to the combustion chamber through the intake pipe, and the high-pressure gas cylinder is used to provide high-pressure gas for the combustion chamber; one end of the exhaust pipe is connected to the combustion chamber, and the other end of the exhaust pipe is located outside the cryogenic tank and is used to discharge the high-pressure gas in the combustion chamber; valves are provided on both the intake pipe and the exhaust pipe; the detection head of the pressure sensor is located inside the combustion chamber and is used to detect the internal pressure of the combustion chamber; the pressure sensor is electrically connected to the signal collector and is used to transmit the detection data of the pressure sensor to the signal collector.

[0022] The present invention also provides an experimental method for a high-pressure cryogenic propellant combustion experiment system, including:

[0023] Step S100, preparing the propellant and installing it into the combustion chamber;

[0024] Step S200, increasing the pressure and decreasing the temperature of the combustion chamber;

[0025] Step S300, igniting the propellant for combustion and recording the experimental data;

[0026] Step S400, processing the experimental data;

[0027] Step S500, relieving the pressure and increasing the temperature of the combustion chamber, and collecting the combustion products;

[0028] Step S500, measuring the combustion efficiency of the propellant, as well as the particle size distribution and morphology of the combustion products.

[0029] Furthermore, the specific content of step S100 is as follows:

[0030] Step S110, cutting the propellant, inserting two thermocouples between the opposite surfaces of the cut propellant, keeping a certain distance between the two thermocouples, and finally bonding and fixing the cut propellant firmly. The thermocouples are connected to the signal acquisition system through temperature compensation leads;

[0031] Step S120, passing the ignition wire of the ignition device through the propellant along the length direction of the propellant, and then fixing the upper end of the propellant on the lower surface of the upper end cover of the combustion chamber;

[0032] Step S130, placing an alumina crucible at the bottom of the combustion chamber, and adding absolute ethanol into the alumina crucible as a combustion product collection medium.

[0033] Furthermore, the specific content of step S400 is as follows:

[0034] Determining the ignition delay time of the propellant: the time required for the propellant to ignite minus the time when the fiber optic spectrometer starts recording data;

[0035] Determine the combustion intensity: Obtained from the emission spectrum during the combustion process of the propellant recorded by the fiber optic spectrometer;

[0036] Determine the combustion state: Obtained by shooting with a high-speed camera;

[0037] Determine the combustion temperature: During the test, determine the combustion temperature by the voltage change of the tungsten-rhenium thermocouple collected by the signal collector;

[0038] Use the polynomial fitting method to perform non-linear fitting on the temperature-voltage conversion curve:

[0039]

[0040] where T is the temperature, E is the voltage of the thermocouple, α is the fitting coefficient, and i represents the order;

[0041] Determine the combustion rate: The high-speed camera determines the combustion rate by capturing the combustion process of the propellant and analyzing the regression of the combustion surface during the combustion process: First, convert the video of the propellant combustion process into an image, then extract the combustion surface of the propellant at different times, and then extract the pixel difference that calibrates the position difference of the combustion surface within different time differences; According to the calibration result, convert the pixel difference of the combustion surface position difference into an actual distance difference, and finally divide the distance difference by the time difference to obtain the combustion rate.

[0042] The present invention also discloses a method for changing the performance of the aluminized solid composite propellant to make it suitable for low temperatures, including:

[0043] Select a binder with a lower glass transition temperature to prepare the aluminized solid composite propellant, and then improve the combustion efficiency by reducing the content of the coral-like structure in the propellant and reducing the agglomerate size in the propellant.

[0044] Furthermore, a binder with a higher combustion energy can also be selected to prepare the aluminized solid composite propellant, and then improve the combustion flame heat feedback by increasing the combustion temperature, so as to improve the combustion rate.

[0045] Furthermore, select a binder with a glass transition temperature of -80 °C to prepare the aluminized solid composite propellant.

[0046] The beneficial effects of the present invention are:

[0047] 1. The low-temperature and high-pressure combustion experiment system for propellant provided by the present invention uses a low-temperature box to cool and heat-insulate the combustion chamber, achieving the effect of adding a low-temperature environment to the propellant combustion chamber. By using a polytetrafluoroethylene sealing ring, the problem of low-temperature brittleness failure that may occur in ordinary sealing rings is solved, and by using a stainless-steel combustion chamber, the problem of deformation failure that may occur in a high-pressure tester in a low-temperature environment is solved.

[0048] 2. The propellant low-temperature high-pressure combustion experiment system provided by the present invention can be used to study the influence of low-temperature high-pressure environment on propellant combustion, conduct feasibility tests on the controllability and high energy of solid rocket motor combustion, observe the combustion state of propellant in low temperature, and quantitatively obtain the variation laws of temperature, pressure and burning rate, solving the problem that it is difficult to conduct propellant combustion experiments under the combined action of low temperature and high pressure.

[0049] 3. The test method provided by the present invention is simple and convenient. The combustion rate is calculated by taking pictures of the combustion morphology of the propellant with a high-speed camera, the combustion temperature of the propellant is measured and calculated by a thermocouple and a computer, and the combustion efficiency is evaluated by collecting combustion products and conducting chemical titration. According to the collected experimental data, it is convenient to summarize the combustion mechanism of the propellant at low temperature. Brief Description of the Drawings

[0050] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the following drawings are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0051] Figure 1 It is a schematic diagram of the high-pressure low-temperature propellant combustion experiment system of the embodiment of the present invention.

[0052] Figure 2 It is a schematic diagram of the installation of the propellant in the combustion chamber of the embodiment of the present invention.

[0053] Figure 3 It is the ignition delay at different initial temperatures of the embodiment of the present invention.

[0054] Figure 4 It is the combustion intensity at different initial temperatures of the embodiment of the present invention.

[0055] Figure 5 It is an image of the propellant combustion at different initial temperatures of the embodiment of the present invention.

[0056] Figure 6 It is the propellant temperature curve at different ambient temperatures of the embodiment of the present invention.

[0057] Figure 7 It is the propellant combustion rate at different ambient temperatures of the embodiment of the present invention.

[0058] Figure 8 It is the ccp particle size at different ambient temperatures of the embodiment of the present invention.

[0059] Figure 9is the combustion efficiency of the propellant in the embodiments of the present invention at different ambient temperatures.

[0060] Figure 10 are the ccp scanning electron microscope images of the embodiments of the present invention at different ambient temperatures.

[0061] In the figure, 1. low-temperature box, 2. combustion chamber, 3. fiber optic spectrometer, 4. high-speed camera, 5. computer, 6. signal collector, 7. crucible, 8. propellant, 9. ignition device, 10. high-pressure gas cylinder, 11. intake pipe, 12. exhaust pipe. Detailed implementation manners

[0062] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0063] In a certain environment, the working temperature of the propellant ranges from 25°C to -110°C, and the annual average temperature is -70°C. Therefore, according to the propellant, it not only needs to have a high energy density to achieve a quantitative load, but also the propellant must be stored for a long time in a large-gradient temperature alternating environment (25°C to -110°C) and work reliably in an environment below -70°C. To study a method to change the performance of the existing aluminum-containing solid composite propellant to make it suitable for low temperatures, the embodiments of the present invention provide a high-pressure low-temperature propellant combustion experiment system, as Figures 1 to 2 shown, including a low-temperature box 1, a combustion chamber 2, a fiber optic spectrometer 3, a high-speed camera 4, a signal collector 6 and a computer 5, wherein:

[0064] The combustion chamber 2 is arranged inside the low-temperature box 1. The combustion chamber 2 is used to install the propellant 8 and provide a combustion space for the propellant 8. The low-temperature box 1 is used to provide a controllable low-temperature environment for the combustion chamber 2; the fiber optic spectrometer 3 is respectively connected to the combustion chamber 2 and the computer 5. The fiber optic spectrometer 3 is used to record the combustion intensity of the propellant 8 and transmit relevant data to the computer 5; the high-speed camera 4 is directly opposite the observation window of the combustion chamber 2 and is connected to the computer 5, and is used to capture the combustion image of the propellant 8, record the combustion process of the propellant 8, and determine the combustion rate by analyzing the regression of the combustion surface during the combustion process through the computer 5. The signal collector 6 is respectively connected to a tungsten-rhenium thermocouple, a pressure sensor (not shown in the figure) and the computer 5, and is used to collect the data of the tungsten-rhenium thermocouple and the pressure sensor and transmit them to the computer 5 for processing. The thermocouple is inserted into the propellant 8 and is used to sense the change in the combustion temperature of the propellant 8; the pressure sensor is used to detect the internal pressure of the combustion chamber 2.

[0065] In some examples, the combustion chamber 2 includes a combustion chamber body, an upper end cover, and a lower end cover. The combustion chamber body has a structure with an internal space and openings at both the upper and lower ends. The upper end cover is fixed at the upper opening of the combustion chamber body, and the lower end cover is fixed at the lower opening of the combustion chamber body. The combustion chamber body, the upper end cover, and the lower end cover form a complete sealed space. The combustion chamber body is provided with an observation window made of quartz glass, which has good heat resistance.

[0066] In some examples, the operating temperature range of the cryogenic tank 1 is -80°C to 0°C to provide a low-temperature environment. The combustion chamber 2 is made of 316 stainless steel and is sealed with a polytetrafluoroethylene sealing ring. The polytetrafluoroethylene sealing ring is located at the connection between the observation window and the combustion chamber body, and at the connections between the upper and lower end covers of the combustion chamber 2 and the combustion chamber body respectively. The polytetrafluoroethylene sealing ring can withstand a low temperature of -80°C and ensure the sealing performance to meet the requirements of the high-pressure and low-temperature environment of the combustion chamber 2.

[0067] In some examples, to meet the high-pressure requirements of the combustion chamber 2, the combustion chamber 2 is connected to a high-pressure transportation system (used to provide a high-pressure environment of 3 MPa to 15 MPa for the combustion chamber 2) to fill the combustion chamber 2 with an inert gas, for example, nitrogen. The combustion chamber 2 can create a high-pressure environment by being filled with nitrogen, and the propellant 8 can be ignited in the nitrogen environment, which can isolate the influence of oxygen in the air on the ignition process. In one example, the high-pressure transportation system includes a high-pressure gas cylinder 10, an inlet pipe 11, and an exhaust pipe 12. During use, the exhaust pipe 12 is closed, the inlet pipe 11 is opened, and after observing that the internal pressure value of the combustion chamber 2 collected by the pressure sensor reaches the required value, the inlet pipe 11 is closed, and then the combustion experiment is carried out. When the pressure is too high, the pressure can be released by opening the exhaust pipe 12. The pressure supply upper limit of the high-pressure gas cylinder 10 is relatively low. Considering the possible need for higher pressure supply conditions in the experiment, the pressure supply method of the high-pressure gas cylinder 10 can be adjusted to the pressure supply by a compressor, and the low-pressure gas is converted into high pressure by the compressor for gas supply.

[0068] The connection method between the pressure sensor and the side of the combustion chamber 2 is a threaded connection, and the detection head of the pressure sensor passes through the cryogenic tank 1 and extends into the combustion chamber 2. The pressure sensor is used to detect the internal pressure of the combustion chamber 2.

[0069] In some examples, a synchronous trigger is also provided. The synchronous trigger is respectively connected to the ignition device 9 of the combustion chamber 2 and the fiber optic spectrometer 3 to realize the synchronization of the data recording of the ignition device 9 and the fiber optic spectrometer 3. Due to the existence of the synchronous trigger, the ignition delay time of the propellant 8 is the time interval between the start of data recording by the fiber optic spectrometer 3 and the start of ignition of the propellant 8.

[0070] In some examples, the front and rear sides of the cryogenic chamber 1 are provided with openings, and double-layer hollow glass is installed at the openings. While ensuring the refrigeration effect, it cooperates with the quartz glass observation windows at the front and rear ends of the combustion chamber 2 to observe the combustion process of the propellant 8.

[0071] In some examples, a crucible 7 is placed on the inner bottom surface of the combustion chamber 2 for receiving the combustion products of the propellant 8.

[0072] In some examples, the ignition device 9 is connected to the propellant 8 through an ignition wire for igniting the propellant 8. In this embodiment, a DC power supply is used as the ignition device 9.

[0073] The present invention also provides a high-pressure cryogenic propellant combustion experiment method, including the following steps:

[0074] Step S100: Prepare the propellant 8 and install it into the combustion chamber 2.

[0075] The composition of the propellant 8 used in this embodiment is 68% by weight of ammonium perchlorate (AP, particle size 200 μm), 14% by weight of HTPB (hydroxyl-terminated polybutadiene), and 18% by weight of Al (particle size 13 μm).

[0076] The propellant 8 is made into a cuboid with dimensions of 5×5×20 mm. The propellant sample is cut, and a thermocouple (in this embodiment, a tungsten-rhenium thermocouple with a diameter of 80 μm is used) is inserted between the opposite surfaces of the cut propellant. The number of tungsten-rhenium thermocouples is two and they are kept parallel. The distance between the two tungsten-rhenium thermocouples is 10 mm. Finally, the cut propellant is bonded and fixed firmly, and the tungsten-rhenium thermocouple is connected to the signal collector 6 through a temperature compensation lead wire.

[0077] The ignition wire of the ignition device 9 is passed through the propellant 8 along the length direction of the propellant 8, and then the upper end of the propellant 8 is fixed to the top of the combustion chamber 2 (i.e., the lower surface of the upper end cover).

[0078] Anhydrous ethanol is added as a collection medium into the alumina crucible 7 at the bottom of the combustion chamber 2. The reason is that on the one hand, anhydrous ethanol is simple to obtain and does not react with the combustion products, and on the other hand, anhydrous ethanol has a low freezing point and remains liquid at -80°C, which is convenient for collecting the combustion products.

[0079] Step S200: Increase the pressure and decrease the temperature of the combustion chamber 2;

[0080] The temperature of the combustion chamber 2 is reduced by the cryogenic chamber 1, and the intake pipe 11 is opened to pressurize the combustion chamber 2 through the high-pressure transportation system. When the temperature and pressure reach the required experimental targets, the intake pipe 11 is closed.

[0081] Step S300: Ignite the propellant for combustion and record the experimental data.

[0082] Turn on the ignition device 9 to ignite the propellant 8. During the combustion process, pay attention to the temperature and pressure changes and adjust them in a timely manner through the cryogenic tank 1 and the high-pressure transportation system.

[0083] After ignition, the combustion products are sprayed into the alumina crucible 7 below, and the combustion products are extinguished and precipitated in absolute ethanol.

[0084] Collect experimental data through the signal collector 6, the high-speed camera 4, and the fiber optic spectrometer 3, and transmit the experimental data to the computer 5.

[0085] Step S400: Process the experimental data.

[0086] Determine the ignition delay time of the propellant 8: Subtract the time when the fiber optic spectrometer 3 starts recording data from the time required for the propellant 8 to ignite.

[0087] Determine the combustion intensity: Obtained from the emission spectrum during the combustion of the propellant 8 recorded by the fiber optic spectrometer 3.

[0088] Determine the combustion state: Obtained by shooting with the high-speed camera 4.

[0089] Determine the combustion temperature: During the test, determine the combustion temperature by the voltage change of the tungsten-rhenium thermocouple collected by the signal collector 6.

[0090] The combustion temperature is obtained by the method of embedding tungsten-rhenium thermocouples. At the same time, according to the dynamic response characteristics of the thermocouple, the temperature measurement data is further processed to obtain the final temperature of the gas-phase flame. Use the polynomial fitting method to perform nonlinear fitting on the temperature-voltage conversion curve. In order to ensure a higher fitting accuracy, a 12th-order generating polynomial is used for fitting calculation:

[0091]

[0092] Among them, T is the temperature, E is the voltage of the thermocouple, α is the fitting coefficient, and i represents the order. The fitting coefficients α of the generated polynomial of the obtained fitting curve are shown in Table 1:

[0093] Table 1 Nonlinear fitting coefficients of tungsten-rhenium thermocouple temperature-thermocouple

[0094] Coefficient Term 0 - <![CDATA[1.01111652984×10 2 > E <![CDATA[-1.51347106608×10 1 > <![CDATA[E 2 > +3.22380039590 <![CDATA[E 3 > <![CDATA[-4.86725776569×10 -1 > <![CDATA[E 4 > <![CDATA[5.06502023695×10 -2 > <![CDATA[E 5 > <![CDATA[-3.63280170367×10 -3 > <![CDATA[E 6 > <![CDATA[1.80222548835×10 -4 > <![CDATA[E 7 > <![CDATA[-6.15045292476×10 -6 > <![CDATA[E 8 > <![CDATA[1.41349974762×10 -7 > <![CDATA[E 9 > <![CDATA[-2.08470109484×10 -9 > <![CDATA[E 10 > <![CDATA[1.77759784942×10 -11 > <![CDATA[E 11 > <![CDATA[-6.64350492029×10 -14 > <![CDATA[E 12 >

[0095] Measure the combustion wave temperature distribution of the propellant 8 through the thermocouple. The thermocouple is pre-embedded in the propellant 8 strip. As the strip burns layer by layer, the thermocouple gradually approaches the combustion surface, and then enters the gas phase region through the combustion surface, so as to measure the entire combustion wave temperature distribution of the propellant 8 from the condensed phase to the gas phase.

[0096] Determine the burning rate: The high-speed camera 4 determines the burning rate by capturing the combustion process of the propellant 8 and analyzing the regression of the burning surface during combustion. First, the video of the combustion process of the propellant 8 is converted into images, then the burning surfaces of the propellant 8 at different times are extracted, and then the pixel differences in the position differences of the burning surfaces within different time differences are extracted. According to the calibration results, the pixel differences in the position differences of the burning surfaces are converted into actual distance differences, and finally the burning rate is obtained by dividing the distance difference by the time difference.

[0097] Step S500: Release pressure and raise the temperature, and collect the combustion products.

[0098] Open the exhaust pipe 12 to release the pressure. At the same time, control the low-temperature box 1 to heat up. After the pressure returns to normal and the temperature rises to room temperature, open the combustion chamber 2, take out the mixture of the combustion products and absolute ethanol, and filter and dry it to obtain the final ccp (condensed-phase combustion products).

[0099] Step S600: Measure the combustion efficiency of the propellant and the particle size distribution and morphology of the combustion products.

[0100] The combustion efficiency of the ccp powder is measured by the potassium dichromate titration method. The active Al in the ccp powder is oxidized by Fe 3+ to generate Al 3+ , while Fe 3+ is reduced to Fe 2+ . During the titration process, potassium dichromate reacts the Fe 2+ in the solution back to Fe 3+ . When all the Fe 2 + is consumed, potassium dichromate reacts with the indicator, turning the solution purple. During the titration process, the consumption of the potassium dichromate solution allows the calculation of the content of active Al in the ccp to determine the combustion efficiency.

[0101] The combustion efficiency calculation formula is as follows:

[0102] w = 1 - cV i M / (3G)

[0103] In the formula, w is the combustion efficiency of the propellant, c is the molar concentration of the potassium dichromate solution, V i is the volume of the consumed potassium dichromate solution, M is the molar mass of aluminum, and G is the mass of the combustion products.

[0104] In this embodiment, the laser particle size analysis method is used to measure the particle size distribution of the ccp. In addition, a scanning electron microscope (SEM) is used to characterize the morphology of the ccp.

[0105] Example

[0106] In this embodiment, five groups of data were established, and the high-pressure and low-temperature ignition experiment of propellant 8 in nitrogen was carried out using the high-pressure and low-temperature propellant combustion experiment system of the present invention. The range of the initial temperature was -80 to 0 °C, the interval was 20 °C, and the high-pressure environment was 7 MPa.

[0107] To compare the ignition characteristics of propellant 8 at different initial temperatures, the ignition delay time of propellant 8 was calculated using the 10% threshold of the maximum intensity (wavelength 486 nm), as Figure 3 shown. As the initial temperature of propellant 8 decreased from 0 °C to -80 °C, the ignition delay time increased monotonically, from 66.5 ms to 180.5 ms, an increase of 171.4%. This indicates that under the experimental conditions, a lower initial temperature of propellant 8 results in poorer ignition performance. This may be due to the fact that as the initial temperature of propellant 8 decreases, the decomposition peak temperature of AP in propellant 8 increases, resulting in a decrease in the AP decomposition ability and a weakening of ignition. In addition, a lower initial temperature requires a higher ignition energy, resulting in an increase in the ignition delay time.

[0108] The emission spectrum generated during the combustion of propellant 8 was recorded using the fiber optic spectrometer 3, as Figure 4 shown. The emission spectrum obtained from the combustion of propellant 8 is broad and continuous, showing characteristic peaks at specific wavelengths. The peaks at 471 nm, 486 nm, and 512 nm correspond to AlO, which is related to the combustion of Al in propellant 8. The peaks at 589 nm, 670 nm, and 768 nm are Na, Li, and K, respectively. Although their concentrations may be very low (impurities inherent in the AP manufacturing process), their colors strongly affect the appearance of the flame, thus producing obvious characteristic spectral lines. The spectral intensity of propellant 8 at an initial temperature of 0 °C is significantly higher than that at other temperatures. As the initial temperature of propellant 8 decreases from 0 °C to -80 °C, the spectral intensity of propellant 8 decreases, further indicating that reducing the initial temperature of the propellant weakens the combustion intensity of the propellant.

[0109] The change process of the combustion of propellant 8 was captured using the high-speed camera 4, as Figure 5As shown. During the test, the exposure parameters of the camera remained unchanged. During the combustion of propellant 8, a complex aggregation and agglomeration process occurred on the combustion surface of propellant 8, which was characterized by the aggregation of visible coral-like structures. These structures consisted of natural aluminum particles, oxidizer, binder, and their decomposition products. These structures adhered to the combustion surface and were continuously deformed by the action of combustion gases and finally detached from the surface. At an initial temperature of 0 °C, there was less accumulation of coral-like structures on the combustion surface. As the initial temperature of propellant 8 decreased from 0 °C to -80 °C, the accumulation of coral-like structures on the combustion surface increased significantly. This finding indicates that lowering the initial temperature of propellant 8 reduces the combustion surface temperature, resulting in the generation of more irregular coral-like structures. As the initial temperature of propellant 8 decreases, both the combustion temperature and the agglomerate size increase.

[0110] Figure 6 Fig. shows the combustion waveforms of propellant 8 at different initial temperatures under atmospheric pressure. The temperature curves of propellant 8 at different initial temperatures showed a slow increase first and then a rapid increase, and finally tended to be stable. The final combustion temperature of propellant 8 exceeded 2000 °C. As the initial temperature of propellant 8 decreased, the flame temperature gradually decreased. The initial temperature of propellant 8 had a significant effect on the temperature rise of the combustion waveform. To compare the temperature gradients near the combustion surface, the data in this temperature range were fitted by linear regression, and the temperature gradients between 500 °C and 1500 °C were determined. The slope obtained from the linear fit represents the temperature gradient, and the temperature gradients of propellant 8 in the range of initial temperatures from 0 °C to -80 °C were 3506.5 °C / mm, 2794.5 °C / mm, 2441.4 °C / mm, 1909.7 °C / mm, and 1817.6 °C / mm in turn. This indicates that the lower the initial temperature of propellant 8, the slower the temperature rise rate of propellant 8. The initial temperature of propellant 8 also significantly affected the final combustion temperature of propellant 8.

[0111] The combustion rate of propellant 8 was evaluated at different initial temperatures, as Figure 7As shown. The propellant 8 exhibits the highest burning rate at an initial temperature of 20 °C. As the initial temperature of the propellant 8 decreases, the burning rate gradually decreases from 7.80 mm / s to 5.69 mm / s, a decrease of 27.1%. In terms of energy balance, the burning rate of the propellant 8 is mainly affected by the heat generated by the solid-phase reaction and the thermal interaction in the gas phase, including conduction and radiation feedback. For the propellant 8, radiation feedback makes a significant contribution to the total heat exchange. Reducing the initial temperature of the propellant 8 results in a smaller temperature gradient, leading to a reduction in conduction feedback, thus limiting the burning rate. As the initial temperature of the propellant 8 decreases, the ignition delay time increases, the accumulation of the coral-like structure increases, and the burning surface temperature of the propellant 8 decreases, reducing the radiative heat feedback, thereby reducing the burning rate. Therefore, the burning rate can be increased by adopting high-energy components to increase the burning temperature and thermal feedback accordingly.

[0112] In addition, at atmospheric pressure, ccp can be collected from the combustion products at different initial temperatures to further study the effects of different initial temperatures on the combustion efficiency and agglomeration characteristics of the propellant 8. The particle size distribution of ccp shows a typical three-peak distribution. For the propellant 8 with an initial temperature of 0 °C (refer to Figure 8 ), the three-peak distribution of ccp is more obvious, with the distribution ranges of 1 μm - 10 μm, 10 μm - 100 μm, and 100 μm - 1000 μm, and the average particle size (D43) is 36 μm. In the propellant 8 at a temperature of -20 °C, the first peak is hardly visible, while the third peak is more prominent, indicating that the average particle size increases to 59 μm. As the initial temperature of the propellant 8 is further reduced to -40 °C, -60 °C, and -80 °C, the first peak is almost invisible, the second peak shifts relative to the third peak, and the average particle size gradually increases to 81 μm, 93 μm, and 135 μm respectively. Generally speaking, reducing the initial temperature of the propellant 8 results in the displacement of the third peak and the average particle size increasing from 36 μm to 135 μm, indicating that reducing the initial temperature of the propellant 8 increases the size of ccp.

[0113] To more accurately evaluate the combustion efficiency of the propellant 8, chemical titration was performed on ccp, as Figure 9 shown. The chemical titration method determines the content of active aluminum in ccp, thereby calculating the combustion efficiency of the propellant 8. The combustion efficiency of the propellant 8 is the highest at 0 °C, reaching 89.5%, while the combustion efficiencies at -20 °C, -40 °C, -60 °C, and -80 °C are 86.9%, 84.2%, 80.3%, and 70.1% respectively. As the initial temperature of the propellant 8 decreases, the combustion efficiency of the propellant 8 also decreases. The morphology of CCP was characterized by SEM, as Figure 10As shown. When the initial temperature is 0 °C, the ccp of the propellant 8 mainly shows regular spherical particles with a smooth surface. As the initial temperature of the propellant 8 decreases, the particle size in the field of view significantly increases and the particle shape becomes irregular. When the propellant is at a low temperature, especially when the ambient temperature is lower than the glass transition temperature, the binder of the propellant will deteriorate, and the coral-like structure in the combustion process of the propellant increases, resulting in an increase in the agglomeration of combustion products and a decrease in combustion efficiency. Therefore, a binder with a lower glass transition temperature can be used to reduce the influence of low temperature on the binder, thereby improving the combustion efficiency.

[0114] Based on the above experimental results, the present invention also provides a method for changing the performance of an aluminized solid composite propellant to make it suitable for low temperatures, including:

[0115] Select a binder with a lower glass transition temperature to prepare the aluminized solid composite propellant, and then by reducing the content of the coral-like structure in the propellant and reducing the agglomerate size in the propellant, so as to improve the combustion efficiency.

[0116] It is also possible to select a binder with a higher combustion energy to prepare the aluminized solid composite propellant, and then by increasing the combustion temperature to increase the heat feedback of the combustion flame, so as to increase the combustion rate.

[0117] In some examples, a binder with a glass transition temperature of -80 °C is selected to prepare the aluminized solid composite propellant.

[0118] Each embodiment in this specification is described in a related manner. The same or similar parts between the embodiments can be referred to each other, and the key points of each embodiment are the differences from other embodiments.

[0119] The above is only the preferred embodiment of the present invention and is not used to limit the protection scope of the present invention. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention are included in the protection scope of the present invention.

Claims

1. A high-pressure and low-temperature propellant combustion experimental system, characterized in that: Used to study a method for changing the performance of existing aluminum-containing solid composite propellants to make them suitable for low temperatures, the experimental system includes: A combustion chamber (2) is arranged inside the cryogenic box (1), the combustion chamber (2) is used to install the propellant (8) and provide a combustion space for the propellant (8), and the cryogenic box (1) is used to provide a controllable low-temperature environment for the combustion chamber (2); an optical fiber spectrometer (3), connected to the combustion chamber (2) and the computer (5), respectively, and the optical fiber spectrometer (3) is used to record the combustion intensity of the propellant (8) and transmit relevant data to the computer (5); A high-speed camera (4) is arranged facing the combustion chamber (2) and connected to a computer (5) for capturing a combustion image of the propellant (8), recording the combustion process of the propellant (8), and transmitting the combustion image to the computer (5); A signal collector (6) is connected to the computer (5) and a thermocouple disposed inside the propellant (8) respectively, and is used to collect a voltage signal of the thermocouple during the combustion of the propellant, and transmit the voltage signal to the computer (5); A high-pressure transport system connected to the combustion chamber (2) and used to provide a high-pressure environment for the combustion chamber (2); An ignition device (9) is connected to the propellant (8) via an ignition line and is used to ignite the propellant (8).

2. A high-pressure and low-temperature propellant combustion experimental system according to claim 1, characterized in that: It also comprises a synchronization trigger, which is respectively connected to the ignition device (9) and the optical fiber spectrometer (3) and is used to synchronize the start of the ignition device (9) and the recording of data by the optical fiber spectrometer (3).

3. A high-pressure and low-temperature propellant combustion experimental system according to claim 1, characterized in that: The combustion chamber (2) comprises: The combustion chamber body is a structure with an internal space and upper and lower openings, and the combustion chamber body is provided with an observation window; the internal space of the combustion chamber body is used to accommodate the propellant (8); An upper end cover is fixed at the upper opening of the combustion chamber body; the propellant (8) is fixed to the lower surface of the upper end cover; A lower end cover, fixed at the lower opening of the combustion chamber body; Sealing rings are located at the connection between the observation window and the combustion chamber body, at the connection between the upper end cover and the combustion chamber body, and at the connection between the lower end cover and the combustion chamber body; the sealing rings are used to ensure the sealing of the combustion chamber (2) in an environment of at least -80°C; The combustion chamber body, the upper end cover, the lower end cover and the sealing ring form a complete closed space.

4. A high-pressure and low-temperature propellant combustion experimental system according to claim 1, characterized in that: The high-pressure transportation system comprises a high-pressure gas cylinder (10), an air intake pipe (11), an exhaust pipe (12) and a pressure sensor, wherein the high-pressure gas cylinder (10) is connected to the combustion chamber (2) via the air intake pipe (11), and the high-pressure gas cylinder (10) is used to provide high-pressure gas to the combustion chamber (2); one end of the exhaust pipe (12) is connected to the combustion chamber (2), and the other end of the exhaust pipe (12) is located outside the low-temperature box and is used to discharge the high-pressure gas from the combustion chamber (2); the air intake pipe (11) and the exhaust pipe (12) are both provided with valves; the detection head of the pressure sensor is located inside the combustion chamber (2) and is used to detect the internal pressure of the combustion chamber (2); the pressure sensor is electrically connected to a signal collector and is used to transmit the detection data of the pressure sensor to the signal collector.

5. The experimental method of a high-pressure and low-temperature propellant combustion experimental system according to any one of claims 1 to 4, characterized in that: include: Step S100, preparing a propellant (8) and installing it into a combustion chamber (2); Step S200, increasing the pressure and lowering the temperature of the combustion chamber (2); Step S300, igniting the propellant to burn and recording experimental data; Step S400, processing the experimental data; Step S500, depressurizing and heating the combustion chamber, and collecting combustion products; Step S600: Determine the propellant combustion efficiency and the particle size distribution and morphology of the combustion products.

6. The experimental method of a high-pressure and low-temperature propellant combustion experimental system according to claim 5, characterized in that: The step S100 is specifically as follows: Step S110, cutting the propellant, inserting two thermocouples between opposite surfaces of the cut propellant, maintaining a certain distance between the two thermocouples, and finally bonding and fixing the cut propellant firmly, and connecting the thermocouples to the signal acquisition system through temperature compensation leads; Step S120, passing the ignition wire of the ignition device (9) through the propellant (8) along the length direction of the propellant (8), and then fixing the upper end of the propellant (8) to the lower surface of the upper end cover of the combustion chamber (2); Step S130: placing an alumina crucible (7) at the bottom of the combustion chamber (2), and adding anhydrous ethanol into the alumina crucible (7) as a combustion product collection medium.

7. The experimental method of a high-pressure and low-temperature propellant combustion experimental system according to claim 6, characterized in that: The step S400 is specifically as follows: Determine the ignition delay time of the propellant (8): the time required for the ignition of the propellant (8) minus the time when the fiber optic spectrometer (3) starts recording data; Determine the combustion intensity: obtain the emission spectrum of the propellant (8) during the combustion process recorded by the fiber optic spectrometer (3); Determine the combustion state: obtain it by taking pictures with a high-speed camera (4); Determining the combustion temperature: During the test, the combustion temperature is determined by the voltage change of the tungsten-rhenium thermocouple collected by the signal collector (6); Use polynomial fitting method to perform nonlinear fitting on the temperature-voltage conversion curve: Where T is the temperature, E is the voltage of the thermocouple, α is the fitting coefficient, and i represents the order; Determine the burning rate: The high-speed camera (4) captures the burning process of the propellant (8) and determines the burning rate by analyzing the regression of the burning surface during the burning process: first, the video of the burning process of the propellant (8) is converted into an image, and then the burning surface of the propellant (8) at different times is extracted, and then the pixel difference of the position difference of the burning surface within different time differences is extracted and calibrated; according to the calibration result, the pixel difference of the position difference of the burning surface is converted into an actual distance difference, and finally the distance difference is divided by the time difference to obtain the burning rate.

8. A method for changing the properties of aluminum-containing solid composite propellant to make it suitable for low temperature, characterized in that: A binder with a lower glass transition temperature is selected to prepare aluminum-containing solid composite propellant, thereby improving combustion efficiency by reducing the content of coral-like structures in the propellant and reducing the size of agglomerates in the propellant.

9. The method of changing the performance of aluminum-containing solid composite propellant to make it suitable for low temperature according to claim 8, characterized in that: It is also possible to select a binder with higher combustion energy to prepare aluminum-containing solid composite propellants, thereby increasing the combustion temperature to increase the thermal feedback of the combustion flame, thereby increasing the combustion rate.

10. The method of changing the performance of aluminum-containing solid composite propellant to make it suitable for low temperature according to claim 8, characterized in that: A binder with a glass transition temperature of -80°C is selected to prepare aluminum-containing solid composite propellant.

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