Composite material plasma erosion and ablation method based on solid rocket engine working condition calibration

Abstract

The invention discloses a composite material plasma erosion ablation method based on solid rocket engine working condition calibration, and belongs to the field of resin material plasma ablation. The method comprises the following steps: selecting a sample material and structure according to a calibration part of the solid rocket engine; obtaining calibration parameters of a calibration part of the solid rocket engine as calibration indexes, wherein the calibration parameters comprise an ablation rate, a back temperature and microstructure characteristics of fibers observed by an electronic scanning microscope; setting an initial plasma ablation process, performing plasma flame ablation on the sample, and measuring calibration parameters of the sample as test indexes after cooling; and dividing the calibration index by the test index to compare the calibration ratios, and if the calibration ratios range from 0.45 to 0.55, completing the calibration. The result of the prepared sample in the plasma flame ablation state is kept consistent with the ablation of the actual solid rocket engine, and the ablation fitting of the actual working condition of the solid rocket engine can be realized.

Description

Technical Field

[0001] The invention belongs to the field of plasma ablation of resin materials, and in particular relates to a composite material plasma scouring and ablation method based on solid rocket engine working condition calibration. Background Art

[0002] With the widespread application of composite materials in the field of solid rocket engines, it is difficult to obtain the real-time ablation of materials in the thermal environment of high temperature, high pressure and high scouring, and thus it is impossible to study its ablation mechanism. Conventional plasma ablation tests are designed based on GJB 323.B, which makes it difficult to fit the complex ablation environment, limiting the design of hot end components of composite materials.

[0003] The conventional plasma ablation method only limits the heat flux density of the jet heat flow. The sample is 30mm in diameter and 10mm in thickness, and is ablated perpendicular to the flame direction. This method has limited simulated processes and is difficult to fit the actual oxidation reaction, mass transfer and heat transfer, interface recoil dynamics, etc. during the operation of the solid rocket engine. It is impossible to provide guidance for the actual design, which is a technical bottleneck problem that needs to be solved urgently. Summary of the invention

[0004] Technical issues to be solved:

[0005] In order to avoid the shortcomings of the prior art, the present invention provides a composite material plasma scouring and ablation method based on solid rocket engine working condition calibration, plasma ablation is performed on the prepared sample, and the ablation process is adjusted, and the ablation condition of the actual solid rocket engine working component is used as the input condition for fitting. The test is calibrated with three indicators: ablation rate, back temperature, and microscopic morphology characteristics of the fiber under electron scanning microscope observation, so that the plasma ablation result is the same as the actual solid rocket engine ablation result, thereby ensuring the accuracy of ablation fitting. The results of the plasma ablation of the sample prepared by the present invention under the plasma flame ablation state are consistent with the ablation of the actual solid rocket engine under the coupling of the three regression mechanisms of oxidation reaction, heat transfer heat, and particle erosion, and can achieve ablation fitting of the actual solid rocket engine working condition.

[0006] The technical solution of the present invention is: a composite material plasma erosion ablation method based on solid rocket engine working condition calibration, the specific steps are as follows:

[0007] Select the sample material and structure according to the solid rocket motor calibration location;

[0008] Acquiring calibration parameters of a solid rocket engine calibration part as calibration indicators, wherein the calibration parameters include ablation rate, back temperature, and microscopic morphology characteristics of the fiber under electron scanning microscope observation;

[0009] Set the initial plasma ablation process, ablate the sample with plasma flame, and measure the calibration parameters of the sample as test indicators after cooling;

[0010] The calibration index is divided by the test index to compare the calibration ratio. If the calibration ratios are all between 0.45 and 0.55, the calibration is completed. If a calibration ratio deviates from 0.5 by the most and is greater than 0.5, the plasma ablation process is adjusted to increase or decrease the severity of ablation. Each time the process is adjusted, the ablation test is repeated until the calibration ratios are all between 0.45 and 0.55.

[0011] A further technical solution of the present invention is: the sample includes a cylinder with a diameter of 10 mm and a disc with a diameter of 30 mm, the cylinder and the disc are coaxially fixedly connected, during the experiment, the cylinder faces the direction of the plasma flame outlet, and a thermocouple is set on the back of the disc to measure the back temperature of the sample.

[0012] A further technical solution of the present invention is that the material and thickness of the sample are consistent with the calibration part of the solid rocket engine.

[0013] A further technical solution of the present invention is: the calculation method of the calibration ablation rate in the calibration parameters is to use a calibrator to divide the retreat length L1 of the material at the calibration position of the solid rocket engine before and after operation by the engine working time t1 to obtain the calibration ablation rate, recorded as A1; the calculation method of the sample ablation rate is to measure the cylindrical length L2 of the sample and divide it by the engine working time t1 to obtain the sample ablation rate, which is A2.

[0014] A further technical solution of the present invention is: the calibration back temperature in the calibration parameters is the temperature rise slope k1 and the maximum temperature T1 of the measured temperature curve of the back wall of the solid rocket engine; the sample back temperature is the slope k2 and the maximum temperature T2 of the thermocouple temperature rise curve set on the back wall of the sample.

[0015] A further technical solution of the present invention is: the calibration microscopic morphology feature in the calibration parameters is that the material at the calibration position is sampled and observed under an electron scanning microscope, and the fiber tip bevel angle is recorded as J1; the sample microscopic morphology feature is that the cylinder of the sample is placed under an electron scanning microscope for observation, and the fiber tip angle is recorded as J2.

[0016] A further technical solution of the present invention is: the initial plasma ablation process is: the distance between the cylinder of the sample and the plasma flame outlet is recorded as S, and S is set to 10mm; the heat flux density of the plasma flame is R, and R is set to 4Mw / m2; the injection flow rate of the aluminum oxide particles is C, and C is set to 1g / s; the injection speed of the aluminum oxide particles is v, and v is set to 100m / s; the diameter of the aluminum oxide particles is D, and D is set to 100μm; the ablation time is t1.

[0017] A further technical solution of the present invention is: the calibration ratio expressions of the calibration index divided by the test index are: A=A2 / (A1+A2), k=k2 / (k2+k1), T=T2 / (T2+T1), J=J2 / (J1+J2).

[0018] A further technical solution of the present invention is: the method for adjusting the plasma ablation process is:

[0019]

[0020] A composite plasma erosion and ablation system based on solid rocket engine working condition calibration includes a plasma ablation device, a sample mounted on a sample clamping table of the plasma ablation device, and a host computer, and the ablation process of the plasma ablation device is controlled by the host computer;

[0021] The host computer includes at least one processor and a memory communicatively connected to the at least one processor; wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor so that the at least one processor can perform a comparison of a calibration ratio by dividing the calibration index with the test index.

[0022] Beneficial Effects

[0023] The beneficial effect of the present invention is that the present invention utilizes the ablation result of a certain part of the component during the actual operation of the solid rocket engine to purposefully adjust the plasma ablation process, so that the plasma ablation result can better represent the actual ablation state of the actual solid rocket engine. By fitting the ablation rate, back temperature, and microscopic morphology, the ablation and degradation process of the material under the influence of heat and mass transfer, oxidation degradation, and physical erosion during the actual operation of the solid rocket engine can be simulated, and then the ablation mechanism research with lower cost, higher efficiency, and wider adjustment range can be carried out through plasma.

[0024] The present invention uses a calibrated plasma erosion ablation method to subject the sample to a thermal environment similar to that of a solid rocket engine, with the heat conduction level, chemical oxidation level, and particle erosion level being substantially the same. Compared with the conventional plasma method, the application of this method can conduct research on the ablation mechanism of materials in a solid rocket engine environment without reaching a high pressure state. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 This is a composite plasma erosion and ablation test process based on solid rocket engine working condition calibration according to an embodiment of the present invention. DETAILED DESCRIPTION

[0026] The embodiments described below with reference to the accompanying drawings are exemplary and are intended to be used to explain the present invention, but should not be construed as limiting the present invention.

[0027] Based on the conventional plasma ablation method, only the heat flux density of the jet heat flow is limited. The sample is 30mm in diameter and 10mm in thickness, and is ablated perpendicular to the flame direction. The method can simulate limited processes, and it is difficult to fit the actual oxidation reaction, mass transfer and heat transfer, interface degeneration mechanics, etc. during the operation of the solid rocket engine. It is impossible to guide the actual design, which is a technical bottleneck problem that needs to be solved urgently. The present invention provides a composite material plasma scouring and ablation method based on solid rocket engine working condition calibration, including determination of a solid rocket engine calibration position, sample preparation, and an adjustment method for feedback of the ablation process according to actual ablation results; determination of the solid rocket engine calibration position determines various calibration parameters of the solid rocket engine; sample preparation must meet the actual material thickness and material composition of the calibration position; the adjustment method of the ablation process after feedback includes the heat flux density R, the speed v of the aluminum oxide particles, the concentration C of the aluminum oxide particles, the distance S between the test piece 1 and the plasma flame outlet, and the diameter D of the aluminum oxide particles.

[0028] The solid rocket engine calibration area must be at least a circle with a diameter of 6 mm.

[0029] The calibration parameters of the solid rocket engine include ablation rate, thickness of internal heat protection material, thickness of back wall material and metal shell, temperature measurement data of shell part, and microscopic morphology of material surface of calibration part.

[0030] The sample consists of two parts, namely, a sample 1 directly contacting the plasma flame and a sample 2 consisting of a back wall material and a metal shell.

[0031] The adjustment method of the ablation process after feedback is as follows: according to the experimental results, the ablation rate, the slope of the heating curve, the maximum temperature, and the fiber tip bevel angle, the adjustment directions of different process parameters will be different according to the deviation from the calibration parameters of the solid rocket engine calibration part.

[0032] The specific steps are as follows:

[0033] Step 1: (Selection method of various calibration parameters of solid rocket engines) Determine the position where the solid rocket engine needs to be fitted as the calibration position, and use a calibrator to divide the material's retreat length L1 at this position before and after work by the engine working time t1 to obtain the calibration ablation rate, recorded as A1. The design thickness of the back metal shell at this position is recorded as F1. Record the temperature rise slope k1 and the maximum temperature T1 of the measured temperature curve of the back wall of the solid rocket engine.

[0034] Step 2: (Method for determining the microscopic morphological characteristics of materials) Sample the material at the calibration position, observe it under an electron scanning microscope, and record the fiber tip bevel angle as J1.

[0035] Step 2: (Test piece preparation method) Process the composite material with the same material as that at this position to form a cylinder with a diameter of 10 mm and a length of L1, which is the test piece 1, and prepare a metal round cake with a diameter of 30 mm and a thickness of F1, which is the test piece 2. Bond the test pieces 1 and 2 at the axis of their cylinders to form the test piece, and install the test piece on the sample clamping table of the plasma ablation test machine, so that the test piece 1 faces the direction of the plasma flame outlet. A thermocouple is attached to the back of the test piece 2, and its temperature is recorded under the test state.

[0036] Step 3: (Initial plasma ablation process record before the test) Set the following test conditions of the plasma ablation equipment: the distance between the test piece 1 and the plasma flame outlet is recorded as S, let S = 10mm; the heat flux density of the plasma flame is R, let R = 4Mw / m2; the injection flow rate of the aluminum oxide particles is C, let C = 1g / s; the injection speed of the aluminum oxide particles is v, let v = 100m / s; the diameter of the aluminum oxide particles is D, let D = 100μm; the ablation time is t1.

[0037] Step 4: (Test index recording) Ablate the specimen 1 with a plasma flame and remove it after it has cooled sufficiently. Measure the length L2 of the specimen 1 and the sample ablation rate A2 = L2 / t1. Record the slope k2 and the highest temperature T2 of the thermocouple heating curve. Observe the specimen 1 under an electron scanning microscope and record the fiber tip angle as J2.

[0038] Step 5: (Ablation process adjustment method based on test result feedback) Divide the calibration index by the test index to compare the calibration ratio: A=A2 / (A1+A2), k=k2 / (k2+k1), T=T2 / (T2+T1), J=J2 / (J1+J2). Sort A, k, T, and J by size. When a value deviates from 0.5 by the most and is greater than 0.5, select a suitable plasma ablation process for adjustment to increase or decrease the severity of ablation. Each time the process is adjusted, repeat step 4 to conduct an ablation test, repeat the sorting of the calibration ratios in step 5, and adjust the ablation process until the four calibration ratios are between 0.45 and 0.55. The adjustment method is shown in Table 1.

[0039] Table 1 Ablation process adjustment method based on test result feedback

[0040]

[0041] The above technical solution is further described below with reference to examples and drawings:

[0042] Reference Figure 1 As shown, when conducting an ablation test, when the calibration position is a certain position of the solid rocket engine, the ablation process adjustment method in Table 1 can be used to make the material ablation closer to the actual situation.

[0043] Example 1: Plasma calibration ablation of the front end surface of a C / C composite material of a certain engine submersible nozzle.

[0044] Step 1: Measure the calibration parameters of the front end of the submerged nozzle, record A1, k1, T1, J1. Also record the thickness F1 of the nozzle metal shell at this location, and the ablation time is 5s.

[0045] Step 2: Prepare a C / C composite material sample 1 and a steel sample 2, and bond them along the axis so that the sample 1 faces the plasma flame.

[0046] Step 3: Set the initial ablation process of the test as follows: the distance between the test piece 1 and the plasma flame outlet is recorded as S, let S = 10mm; the heat flux density of the plasma flame is R, let R = 4Mw / m2; the injection flow rate of the aluminum oxide particles is C, let C = 1g / s; the injection speed of the aluminum oxide particles is v, let v = 100m / s; the diameter of the aluminum oxide particles is D, let D = 100μm; the ablation time is 5s. Perform the first ablation.

[0047] Step 4: Record the results of the first ablation and adjust the process according to Table 1. It is found that the ablation rate deviation is the largest and is a positive deviation. Therefore, in the second test, other test processes remain unchanged, and the heat flux density is adjusted to 3.6Mw / m2 for the second ablation test.

[0048] Step 5: Record the results of the second ablation and adjust the process according to Table 1. It is found that the maximum temperature deviation of the back wall is the largest and negative. Therefore, the other test processes are kept unchanged in the third test, and the incident particle speed is adjusted to 110m / s for the third ablation test.

[0049] Step 6: Record the results of the third ablation and adjust the process according to Table 1. It is found that the maximum temperature deviation of the back wall is still the largest and is a negative deviation. Therefore, the other test processes are kept unchanged in the fourth test, and the incident particle speed is adjusted to 120m / s for the fourth ablation test.

[0050] Step 7: Record the fourth ablation result and find that all deviation values ​​are between 0.45 and 0.55, which indicates that the test is successful. The ablation process of the front end surface of the C / C composite material of a certain engine submersible nozzle is thus obtained.

[0051] Example 2: Plasma calibration ablation at the inlet of a resin composite expansion section of a certain engine submersible nozzle.

[0052] Step 1: Measure the calibration parameters of the resin composite material at the entrance of the expansion section, record A1, k1, T1, J1. Also record the thickness F1 of the nozzle metal shell at this location, and the ablation time is 5s.

[0053] Step 2: Prepare a resin composite material sample 1 and an aluminum sample 2, and bond them along the axis so that the sample 1 faces the plasma flame.

[0054] Step 3: Set the initial ablation process of the test as follows: the distance between the test piece 1 and the plasma flame outlet is recorded as S, let S = 10mm; the heat flux density of the plasma flame is R, let R = 4Mw / m2; the injection flow rate of the aluminum oxide particles is C, let C = 1g / s; the injection speed of the aluminum oxide particles is v, let v = 100m / s; the diameter of the aluminum oxide particles is D, let D = 100μm; the ablation time is 5s. Perform the first ablation.

[0055] Step 4: Record the results of the first ablation and adjust the process according to Table 1. It is found that the slope k of the heating curve has the largest deviation and is a negative deviation. Therefore, in the second test, other test processes remain unchanged, and the heat flux is adjusted to 4.4Mw / m2 for the second ablation test.

[0056] Step 5: Record the results of the second ablation and adjust the process according to Table 1. It is found that the maximum temperature T of the back wall has the largest deviation and is a positive deviation. Therefore, the other test processes are kept unchanged in the third test, and the incident particle speed is adjusted to 90m / s for the third ablation test.

[0057] Step 6: The third ablation result was recorded, and the process was adjusted according to Table 1. It was found that the fiber tip angle J had the largest deviation and was a negative deviation. Therefore, the other test processes were kept unchanged in the fourth test, and the incident particle concentration was adjusted to 1.1 g / s for the fourth ablation test.

[0058] Step 7: Record the fourth ablation result and find that all deviation values ​​are between 0.45 and 0.55, and the test is deemed successful. Thus, the ablation process of plasma calibration ablation at the inlet of the expansion section of the resin composite material of a certain engine submersible nozzle is obtained.

[0059] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and are not to be construed as limitations on the present invention. A person skilled in the art may change, modify, replace and modify the above embodiments within the scope of the present invention without departing from the principles and purpose of the present invention.

Claims

1. A composite material plasma erosion ablation method based on solid rocket engine working condition calibration, characterized in that The specific steps are as follows: Select the sample material and structure according to the solid rocket motor calibration location; Acquiring calibration parameters of a solid rocket engine calibration part as calibration indicators, wherein the calibration parameters include ablation rate, back temperature, and microscopic morphology characteristics of the fiber under electron scanning microscope observation; Set the initial plasma ablation process, ablate the sample with plasma flame, and measure the calibration parameters of the sample as test indicators after cooling; The calibration index is divided by the test index to compare the calibration ratio. If the calibration ratios are all between 0.45 and 0.55, the calibration is completed. If a calibration ratio deviates from 0.5 by the most and is greater than 0.5, the plasma ablation process is adjusted to increase or decrease the severity of ablation. Each time the process is adjusted, the ablation test is repeated until the calibration ratios are all between 0.45 and 0.

55.

2. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 1, characterized in that: The sample includes a cylinder with a diameter of 10 mm and a disk with a diameter of 30 mm. The cylinder and the disk are coaxially fixedly connected. During the experiment, the cylinder faces the direction of the plasma flame outlet, and a thermocouple is set on the back of the disk to measure the back temperature of the sample.

3. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 1, characterized in that: The material and thickness of the sample are consistent with the calibration part of the solid rocket engine.

4. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 2, characterized in that: The calibration ablation rate in the calibration parameters is calculated by using a caliper to divide the material withdrawal length L1 at the calibration position of the solid rocket engine before and after operation by the engine operating time t1 to obtain the calibration ablation rate, recorded as A1; the sample ablation rate is calculated by measuring the cylindrical length L2 of the sample and dividing it by the engine operating time t1 to obtain the sample ablation rate, which is A2.

5. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 4, characterized in that: The calibration back temperature in the calibration parameters is the temperature rise slope k1 and the maximum temperature T1 of the measured temperature curve of the back wall of the solid rocket engine; the sample back temperature is the slope k2 and the maximum temperature T2 of the thermocouple temperature rise curve set on the back wall of the sample.

6. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 5, characterized in that: The calibration microscopic morphology feature in the calibration parameters is that the material at the calibration position is sampled and observed under an electron scanning microscope, and the fiber tip bevel angle is recorded as J1; the sample microscopic morphology feature is that the cylinder of the sample is placed under an electron scanning microscope for observation, and the fiber tip angle is recorded as J2.

7. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 6, characterized in that: The initial plasma ablation process is as follows: the distance between the cylinder of the sample and the plasma flame outlet is recorded as S, and S=10mm; the heat flux density of the plasma flame is R, and R=4Mw / m2; the injection flow rate of the aluminum oxide particles is C, and C=1g / s; the injection speed of the aluminum oxide particles is v, and v=100m / s; the diameter of the aluminum oxide particles is D, and D=100μm; and the ablation time is t1.

8. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 7, characterized in that: The calibration ratio expressions of the calibration index divided by the test index are: A=A2 / (A1+A2), k=k2 / (k2+k1), T=T2 / (T2+T1), J=J2 / (J1+J2).

9. The composite material plasma erosion ablation method based on solid rocket engine working condition calibration according to claim 8, characterized in that: The method for adjusting the plasma ablation process is:

10. A composite material plasma erosion and ablation system based on solid rocket engine working condition calibration, used to implement the composite material plasma erosion and ablation method based on solid rocket engine working condition calibration according to any one of claims 1 to 9; characterized in that: It includes a plasma ablation device, a sample installed on a sample clamping table of the plasma ablation device, and a host computer, and the ablation process of the plasma ablation device is controlled by the host computer; The host computer includes at least one processor and a memory communicatively connected to the at least one processor; The memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor so that the at least one processor can compare the calibration ratio by dividing the calibration index with the test index.

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