CN113047981A

CN113047981A - Method for judging effectiveness of initial experimental data in solid propellant burning rate test by impulse method

Info

Publication number: CN113047981A
Application number: CN202110283176.2A
Authority: CN
Inventors: 王英红; 张鑫鹏; 唐红梅; 吴世曦; 杨虹
Original assignee: Northwestern Polytechnical University; Hubei Institute of Aerospace Chemical Technology
Current assignee: Northwestern Polytechnical University; Hubei Institute of Aerospace Chemical Technology
Priority date: 2021-03-16
Filing date: 2021-03-16
Publication date: 2021-06-29
Anticipated expiration: 2041-03-16
Also published as: CN113047981B

Abstract

The invention discloses a method for judging the effectiveness of initial experimental data in a solid propellant burning rate test by an impulse method, which comprises the steps of firstly recording the outline dimension of a tubular grain of a solid propellant, then coating the outer surface and the end surface of the tubular grain, and only keeping the inner surface as an initial burning surface; then filling the tubular explosive column into a combustion chamber of the solid rocket engine; the solid rocket engine is ignited, the combustion pressure and the generated thrust in the working process of the solid rocket engine are recorded, pressure and thrust time curves are drawn on the same graph, the pressure and thrust time curves are analyzed, and the effectiveness of the thrust curve is judged, namely ignition synchronization and whether erosive combustion exists are judged. The invention provides technical support for the specific realization of the impulse method and mass flow rate method burning rate testing technology, and guarantees the consistency and effectiveness of the testing result.

Description

Method for judging effectiveness of initial experimental data in solid propellant burning rate test by impulse method

Technical Field

The invention belongs to the technical field of rocket engines, and particularly relates to an experimental data validity judgment method.

Background

The burning speed of the solid propellant is a basic parameter for designing the rocket engine and is an important parameter for predicting the ballistic performance of the rocket engine. The burning rate is defined as the distance of the propellant charge combustion surface moving back along its normal direction per unit time, and is referred to as the burning rate of the solid propellant for short. The method not only directly determines the release rate of the energy of the solid propellant, but also calculates other combustion performances of the solid propellant, such as a combustion rate coefficient, a combustion rate pressure index, a combustion rate temperature sensitive coefficient, an erosion ratio and other core parameters.

The research on the high-pressure dynamic burning rate and pressure index of the solid propellant in the paper of the bulletin of dynamite academic newspaper in 3.2019 and the research on the burning rate test of the mass flow rate method based on the working principle of the solid rocket engine in the paper of the bulletin of dynamite academic newspaper in 6.2020 proposes that the burning rates under different pressure spans are calculated by using a certain configuration of surface-enhanced combustion explosive in an engine test through a thrust-time curve and a pressure-time curve measured by one experiment. The experimental principles of impulse-and mass flow-rate-based burn rate testing are described in the two papers, both of which require that the initial combustion surface of the propellant during combustion be the exposed interior surface of the charge and that the combustion surface recede in parallel layers. There are two cases, that is, when the initial burning surface is not ignited synchronously, the initial burning surface is not the inner surface of the tubular grain; and if corrosion combustion exists, the combustion surface of the propellant does not move back along the parallel layer. However, how to determine ignition synchronization and whether there is erosive combustion is not mentioned in the two papers.

Disclosure of Invention

In order to overcome the defects of the prior art, the invention provides a method for judging the effectiveness of original experimental data in a solid propellant burning rate test by an impulse method, which comprises the steps of firstly recording the outline dimension of a tubular grain of the solid propellant, then coating the outer surface and the end surface of the tubular grain, and only keeping the inner surface as an initial burning surface; then filling the tubular explosive column into a combustion chamber of the solid rocket engine; the solid rocket engine is ignited, the combustion pressure and the generated thrust in the working process of the solid rocket engine are recorded, pressure and thrust time curves are drawn on the same graph, the pressure and thrust time curves are analyzed, and the effectiveness of the thrust curve is judged, namely ignition synchronization and whether erosive combustion exists are judged. The invention provides technical support for the specific realization of the impulse method and mass flow rate method burning rate testing technology, and guarantees the consistency and effectiveness of the testing result.

The technical scheme adopted by the invention for solving the technical problem comprises the following steps:

step 1: the solid propellant is in the shape of a tubular grain; recording the external dimensions of the tubular explosive column, including the external diameter D, the internal diameter D and the length L, and weighing the mass m of the tubular explosive column;

step 2: coating the outer surface and the end surface of the tubular grain, and only keeping the inner surface as an initial combustion surface;

and step 3: filling the tubular charge in a combustion chamber of a solid rocket engine;

and 4, step 4: igniting the solid rocket engine, recording the combustion pressure p (t) and the generated thrust F (t) of the solid rocket engine in the working process, drawing curves of p (t) and F (t) on the same graph, wherein t is time;

and 5: analyzing the p (t) curve and the F (t) curve, and judging the effectiveness of the F (t) curve;

step 5-1: if the p (t) curve and the F (t) curve both continuously increase along with time, the pushing force F (t) is reduced after the pushing force F (t) is increased to the maximum value Fmax, and the time for the F (t) curve to be reduced from the maximum value Fmax to 70% Fmax is less than or equal to 80ms, judging that the F (t) curve is effective, namely the solid propellant finishes burning at the maximum burning surface;

step 5-2: if the p (t) curve and the F (t) curve are continuously increased along with time, after the pushing force F (t) is increased to the maximum value Fmax, the time from the maximum value Fmax to the point where the F (t) curve is reduced to 70% Fmax is more than 80ms, the F (t) curve is judged to be invalid, namely the maximum burning surface of the solid propellant is not the final burning surface, because the naked surface of the propellant is not ignited synchronously or has erosive burning.

Preferably, said thrust force f (t) has a maximum value Fmax comprised between 1.5KN and 2.5 KN.

Preferably, the solid propellant is in surface-increasing combustion, the solid rocket engine has an increasing trend after ignition, a maximum thrust point generated by combustion corresponds to the time when the solid propellant is burnt out, and an F (t) curve after the maximum point is a tail section.

The invention has the following beneficial effects:

the method takes the results caused by 'asynchronous ignition on the exposed surface of the propellant' and 'erosive combustion' as the starting point, solves the problem that 'asynchronous ignition on the exposed surface of the propellant' and 'erosive combustion' cannot be judged, provides technical support for the specific realization of the impulse method and mass flow rate method combustion rate testing technology, and ensures the consistency and the effectiveness of the testing results.

Drawings

FIG. 1 is a schematic view of a solid rocket engine used in the method of the present invention.

FIG. 2 is a schematic view of a combustion pressure and thrust force acquisition system for use in the method of the present invention.

FIG. 3 is a schematic view of the structure of the cartridge used in the method of the present invention.

FIG. 4 is a graph illustrating effective pressure p (t) and thrust F (t) curves according to an embodiment of the present invention;

fig. 5 is a simplified model of the solid propellant ignition non-uniform combustion of the present invention.

FIG. 6 is an exemplary erosive combustion of an inner bore combustion charge of an embodiment of the present invention, wherein the left figure is the inner bore combustion charge; the right picture is the inner hole of the rear bell-mouth charge column caused by erosion combustion.

FIG. 7 is a diagram of exemplary invalid raw data.

FIG. 8 is a graph of raw data for the effectiveness of nozzle ablation in accordance with an embodiment of the present invention.

FIG. 9 is a test raw curve of Bisby 3 according to example of the present invention.

FIG. 10 is a graph comparing the results of three test burn rates for the same propellant of examples of the present invention.

In the figure, 1 — thrust sensor; 2-a pressure sensor; 3-an igniter; 4-a combustion chamber housing; 5-a solid propellant charge; 6, spraying a pipe; 7-nozzle throat insert.

Detailed Description

The invention is further illustrated with reference to the following figures and examples.

In the burning rate test, the ignition synchronism of the initial burning surface of the propellant grain is consistent, namely the uncoated surface of the propellant is instantly and completely ignited by the ignition powder, and the propellant is in steady-state combustion, and abnormal combustion including erosion combustion is not considered, which is a precondition of the burning rate test. The invention analyzes the experimental phenomena respectively generated by 'asynchronous ignition' and 'erosive combustion', takes the result caused by 'asynchronous ignition on the exposed surface of the propellant' and 'erosive combustion' as the starting point, and solves the problem that 'asynchronous ignition on the exposed surface of the propellant' and 'erosive combustion' can not be judged.

A method for judging the effectiveness of original experimental data of a solid propellant burning rate test by an impulse method comprises the following steps:

step 2: as shown in fig. 3, the outer surface and the end surface of the tubular charge are coated, and only the inner surface is reserved as an initial combustion surface;

and step 3: freely filling the tubular explosive column in a combustion chamber of a solid rocket engine; as shown in fig. 1, a solid rocket engine is comprised of an igniter, a combustion chamber, a solid propellant charge, and a nozzle;

and 4, step 4: as shown in fig. 2, the solid rocket engine is ignited, the combustion pressure p (t) and the generated thrust force F (t) of the solid rocket engine in the working process are recorded, curves p (t) and F (t) are drawn on the same graph, and t is time; as shown in fig. 4, since the solid propellant is in the area-increasing combustion, the solid rocket engine has an increasing trend after ignition, the maximum thrust point generated by combustion corresponds to the time when the solid propellant is burnt out, and the curve f (t) after the maximum point is a tail section;

step 5-1: as shown in fig. 4, both the p (t) and f (t) curves continue to increase with time, or as shown in fig. 8, there is no rapid decrease or even an equal segment of pressure after the combustion pressure p (t) increases to a maximum, but the thrust force f (t) continues to increase throughout the combustion segment, including the combustion pressure p (t) is constant, and the thrust force f (t) also increases; if the pushing force F (t) is rapidly reduced after the pushing force F (t) is increased to the maximum value Fmax, and the time for the F (t) curve to be reduced from the maximum value Fmax to 70% Fmax is less than or equal to 80ms, judging that the F (t) curve is effective, namely the solid propellant finishes burning at the maximum burning surface; the combustion pressure in the combustion process is slowly increased or not increased due to the ablation of the nozzle, but is always increased in area, so the thrust F (t) is increased in whole process, and the data is effective;

step 5-2: as shown in fig. 7, if both the p (t) curve and the f (t) curve continue to increase with time, there is no rapid decrease after increasing to a maximum value, or even a constant pressure and thrust segment, because the ignition is not synchronized or there is erosive combustion, the maximum combustion surface is not the final combustion surface. After the pushing force F (t) is increased to the maximum value Fmax, the time from the maximum value Fmax to the point where the F (t) curve is reduced to 70% Fmax is longer than 80ms, and the F (t) curve is judged to be invalid, namely the maximum burning surface of the solid propellant is not the final burning surface, because the naked surface of the propellant is not ignited synchronously or eroded and burnt.

The exposed surface of the propellant is ignited asynchronously, namely the combustion surface close to one side of the ignition charge is ignited firstly, and the combustion surface far away from one side of the ignition charge is ignited later, as shown in fig. 5, Lx is the length of the combustion surface which is burnt firstly, and theta is the included angle between the combustion surface which is burnt firstly and the combustion surface which is burnt later. The shape surface of the powder column is a front end bell mouth, namely, the exposed initial burning surface is not synchronously ignited. FIG. 6 is a schematic view of a typical erosive combustion of an inner bore combustion charge having a flared rearward profiled end of the charge. Thus, both "asynchronous ignition on the exposed surface of the propellant" and "erosive combustion" result in the final combustion surface of the charge not being the maximum combustion surface.

The reason for the slow decline of the f (t) curve is that the ignition synchronism is poor, so that the combustion is not according to the expected combustion surface, namely the combustion termination time is the maximum combustion surface, and the whole combustion is complete, because the ignition synchronism is poor, so that the residual medicine still burns after the maximum thrust Fmax in the graph is reached, a certain compensation is given to the thrust, so that the thrust curve does not directly decline, but slowly declines, and the original data is invalid.

The specific embodiment is as follows:

the method is adopted to analyze and judge the test data of different solid propellants and select the solid propellant with excellent performance.

1. The shape of the solid propellant to be measured is measured, and the shape of the propellant is tubular grain, so the inner diameter and the outer diameter of the grain and the diameter of the spray pipe need to be recorded, and the grain is different propellant size parameters as shown in table 1.

TABLE 1 three shot Diety propellant size parameters

2. And coating the outer surface and the end surface of the solid propellant grain to be detected according to requirements, and only reserving the inner surface as an initial combustion surface.

3. The solid propellant coated according to the requirement is freely filled in a combustion chamber of the rocket engine. And connecting an ignition circuit, starting an ignition button after the circuit connection is confirmed to be correct, and simultaneously observing the real-time change condition of the pressure intensity in the combustion chamber displayed by a pressure sensor connected with the combustion chamber and the thrust change condition displayed by a thrust sensor connected with the engine. The original thrust-time curve and pressure-time curve of the whole combustion process are obtained, and the sizes of the grains corresponding to fig. 7, 8 and 9 are biradical 1, biradical 2 and biradical 3 in table 1 respectively.

4. The burning rate results of 3 tests obtained by the impulse burning rate calculation method are shown in table 2, and the comparison of the response results is shown in fig. 10. It can be seen that bistatic 2 and bistatic 3 burn rate consistency is good, while bistatic 1 has no comparability as a result of the invalidity of the original data.

TABLE 2 results of three experiments with bis-based propellants

Claims

1. A method for judging the effectiveness of original experimental data in a solid propellant burning rate test by an impulse method is characterized by comprising the following steps:

2. The method for determining the validity of the original experimental data of the impulse-method solid propellant burning rate test according to claim 1, wherein the maximum value Fmax of the thrust force f (t) is between 1.5KN and 2.5 KN.

3. The method for determining the effectiveness of the initial experimental data of the impulse-method solid propellant burning rate test according to claim 1, wherein the solid propellant is surface-enhanced combustion, the solid rocket engine has an increasing trend after ignition, a maximum thrust point generated by combustion corresponds to a time when the solid propellant is burnt out, and an F (t) curve after the maximum point is a trailing section.