Multiple power source series ejection discrete pulse timing fitting control method

Abstract

This invention discloses a discrete pulse timing fitting control method for a complex power source series catapult launch, capable of controlling the mass flow rate of multiple small pulse engines. The method achieves consistency between the fitted mass flow rate curve and the theoretical fitting result, maintaining constant overload and stable operation of the missile during launch. The method includes the following steps: Step 1: Solve for the ideal overload using internal ballistics knowledge based on the required missile parameters. Step 2: Obtain the ideal pressure according to the dynamic equations. Step 3: Substitute the ideal pressure into the control equations to obtain the ideal engine mass flow rate. Step 4: Determine whether the interpolation between the ideal engine mass flow rate and the actual mass flow rate meets the requirements. If yes, return to Step 2; otherwise, proceed to Step 5. Step 5: Open a new release point; calculate the actual mass flow rate, and return to Step 2.

Description

Technical Field

[0001] This invention relates to the field of missile launch power source control technology, specifically to a method for fitting the timing of discrete pulses in series launch of complex power sources. Background Technology

[0002] Missile launchers are classified into hot launch and cold launch types according to their launch power. Hot launch means that the missile detaches from the launcher by its own power, that is, the missile ignites and flies directly on the launcher. Cold launch means that the missile is ejected from the tube by an external ejection device, and takes off with the help of auxiliary power. The missile's main engine is then ignited when it reaches a certain altitude.

[0003] Hot launch is currently the most widely used missile launch method. However, this method generates a large amount of high-temperature, high-speed exhaust gas containing solid particles. Improper handling of this high-temperature, high-speed exhaust gas can damage the launch system or even cause the missile to explode. In contrast to hot launch, missiles launched in a cold launch mode already have a certain initial velocity before engine ignition, which increases the missile's range. At the same time, when the missile engine ignites, the missile has already left the ground at a certain altitude, reducing the impact of exhaust gas on the launch system, avoiding complex exhaust gas routing issues, and improving the missile weapon's environmental adaptability.

[0004] Currently, small pulse engine catapult technology, using compressed air with its strong instantaneous expansion and high power density as the working medium, offers advantages such as a high power-to-mass ratio, no pollution, flame retardancy, explosion protection, electromagnetic interference resistance, low working fluid temperature, no need for thermal protection measures, good versatility, and low cost, making it a promising application in missile launches. However, this launch method also has many weaknesses that hinder precise control, such as strong nonlinearity and model uncertainty. Achieving accurate control of the small pulse engine catapult process remains a challenge. Summary of the Invention

[0005] In view of this, the present invention provides a discrete pulse timing fitting control method for multiple power source series catapults, which can control the mass flow rate of multiple small pulse engines, and achieve the result of mass flow rate curve fitting is consistent with the theoretical fitting result, thereby realizing the control of the projectile overload to remain constant and the smooth operation during the launch process.

[0006] To achieve the above objectives, the technical solution of the present invention includes the following steps:

[0007] Step 1: Based on the required missile parameters, use knowledge of internal ballistics to solve for the ideal overload.

[0008] Step 2: Obtain the ideal pressure based on the dynamic equation.

[0009] Step 3: Substitute the ideal pressure into the control equation to obtain the ideal engine mass flow rate.

[0010] Step 4: Determine whether the interpolation between the ideal engine mass flow rate and the actual mass flow rate meets the requirements. If yes, return to step 2; otherwise, proceed to step 5.

[0011] Step 5: Open a new bottle to increase the release point; calculate the actual mass flow rate, and return to Step 2.

[0012] Further, step 1: Based on the required missile parameters, solve for the ideal overload using knowledge of internal ballistics, specifically as follows:

[0013] Missile parameters include initial chamber volume, missile diameter, atmospheric pressure, operating parameters of each pulse engine, missile mass, and friction coefficient f.

[0014] Ideal parameters include ideal overload, ideal outlet velocity, and ideal acceleration-time curve.

[0015] Further, step 2: Obtain the ideal pressure according to the kinetic equation, specifically as follows:

[0016] According to the formula

[0017]

[0018] Where P2 is the ideal pressure of the launch tube, a i M represents the ideal acceleration value at different times. m Let P be the mass of the missile, g be the acceleration due to gravity, f be the coefficient of friction between the missile and the tube wall, and P be the mass of the missile. a Where A is the ambient pressure and A is the cross-sectional area of ​​the launch tube;

[0019] Thus, the ideal pressure P2(i) at different times is obtained, where i is the time.

[0020] Further, step 3: Substitute the ideal pressure into the control equation to obtain the ideal engine mass flow rate;

[0021]

[0022] Where, q i (i) represents the ideal gas mass flow rate at time i, V2(i) represents the internal volume of the launch tube at time i, and R g Let P2(i) represent the constant of the working gas in the engine, P2(i) be the ideal pressure inside the launch tube, P2 = Pa at the initial moment, P2(i+1) be the ideal pressure inside the tube at the next moment, T2 be the temperature inside the launch tube, dt be the duration of each iteration step, A be the cross-sectional area of ​​the launch tube, and v(i) be the missile velocity at the current time step.

[0023] Preferably, the missile velocity v(i) at the current time step is obtained by multiplying the difference Δt between the acceleration at the current iteration step and the acceleration at the previous time step with the time step length.

[0024] Furthermore, to determine whether the interpolation between the ideal engine mass flow rate and the actual mass flow rate meets the requirements, specifically: let the difference between the ideal engine mass flow rate and the actual mass flow rate be Δq, and set a dual criterion to determine whether the difference between the actual mass flow rate and the ideal mass flow rate meets the requirements:

[0025] Criterion 1 is the ratio of Δq to the ideal engine mass flow rate, denoted as r1;

[0026] Criterion 2 is the ratio of Δq to the initial gas flow rate of a single engine, denoted as r2;

[0027] Set a threshold;

[0028] If either parameter r1 or r2 is less than the set threshold, the requirement is met; if all parameters r1 and r2 are not less than the set threshold, the requirement is not met.

[0029] Preferably, the threshold is set to any value between 0 and 1.

[0030] Beneficial effects:

[0031] 1. This invention proposes a discrete pulse timing fitting control method for complex power source tandem catapult launch. Relying on the tandem operation of a single type of pulse engine, it can flexibly realize power source curves for different needs, saving the manufacturing cost of small-scale pulse engines. Compared with traditional methods based on finite element mesh calculations of propellant combustion, the advantages of this invention are: through the above fitting calculation process, an ideal flow curve can be obtained, leading to a time-series superposition optimization algorithm, providing a numerical calculation method for subsequent internal trajectory and flow field simulation coupled calculations.

[0032] 2. The present invention provides a discrete pulse timing fitting control method for complex power source series catapults, which fits the timing from two aspects: total flow and single engine flow. It adopts the "dual criterion" idea to fit the ideal mass flow curve, reasonably controls the fitting process, and improves the fitting accuracy. Attached Figure Description

[0033] Figure 1 This is a flowchart of a complex power source series catapult discrete pulse timing fitting control method according to an embodiment of the present invention;

[0034] Figure 2 This is a schematic diagram of the launch system structure;

[0035] Figure 3 Mass flow rate curve for a single type of pulse engine;

[0036] Figure 4 This is a graph showing the fitting results. Detailed Implementation

[0037] The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0038] This invention provides a method for fitting and controlling the discrete pulse timing of a missile launcher with a complex power source series launch, targeting a missile launch device structure such as... Figure 2 As shown, it should include at least one small pulse engine, where a single engine is insufficient to meet the missile launch requirements. For example... Figure 1 As shown, the method includes the following steps:

[0039] Step 1: Based on the required missile parameters, use knowledge of internal ballistics to solve for the ideal overload.

[0040] Missile parameters include initial chamber volume, missile diameter, atmospheric pressure, operating parameters of each pulse engine, missile mass, and friction coefficient f.

[0041] Ideal parameters include ideal overload, ideal outlet velocity, and ideal acceleration-time curve.

[0042] Step 2: Obtain the ideal pressure based on the kinetic equations; according to the formula...

[0043]

[0044] Where P2 is the ideal pressure of the launch tube, a i M represents the ideal acceleration value at different times. m Let P be the mass of the missile, g be the acceleration due to gravity, f be the coefficient of friction between the missile and the tube wall, and P be the mass of the missile. a Where A is the ambient pressure and A is the cross-sectional area of ​​the launch tube;

[0045] Thus, the ideal pressure P2(i) at different times is obtained, where i is the time.

[0046] Step 3: Substitute the ideal pressure into the control equation to obtain the ideal engine mass flow rate; specifically:

[0047]

[0048] Where, q i (i) represents the ideal gas mass flow rate at time i, V2(i) represents the internal volume of the launch tube at time i, and R gLet P2(i) represent the constant working gas volume of the engine, P2(i) be the ideal pressure inside the launch tube (initially P2 = Pa, P2(i+1) be the ideal pressure inside the tube at the next moment), T2 be the temperature inside the launch tube (set to a constant value of 400 K in this embodiment), dt be the duration of each iteration step (set to 0.001 s in this embodiment), A be the cross-sectional area of ​​the launch tube, and v(i) be the missile velocity at the current time step. The missile velocity v(i) at the current time step is obtained by multiplying the difference Δt between the acceleration of the current iteration step and the acceleration of the previous time step by the time step duration.

[0049] Figure 3 This is a mass flow rate curve for a single type of pulse engine.

[0050] Step 4: Determine whether the interpolation between the ideal engine mass flow rate and the actual mass flow rate meets the requirements. If yes, return to step 2; otherwise, proceed to step 5.

[0051] Specifically: Let the difference between the ideal engine mass flow rate and the actual mass flow rate be Δq, and set two criteria to determine whether the difference between the actual mass flow rate and the ideal mass flow rate meets the requirements:

[0052] Criterion 1 is the ratio of Δq to the ideal engine mass flow rate, denoted as r1;

[0053] Criterion 2 is the ratio of Δq to the initial gas flow rate of a single engine, denoted as r2;

[0054] Set a threshold; set the threshold to any value between 0 and 1.

[0055] If either parameter r1 or r2 is less than the set threshold, the requirement is met; if all parameters r1 and r2 are not less than the set threshold, the requirement is not met.

[0056] Step 5: Open a new bottle to increase release points;

[0057] Step 6: Calculate the actual mass flow rate and return to Step 2. The fitted curve is shown below. Figure 4 As shown.

[0058] In summary, the above are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method of control of a plurality of power sources in series with discrete pulse timing for launch, characterized by, The method comprises the following steps: Step 1: according to the required missile parameters, solve the ideal parameters by using the internal ballistic related knowledge; the missile parameters include initial volume, body diameter, atmospheric pressure, each pulse engine working parameter, missile mass and friction coefficient ; The ideal parameters comprise an ideal overload, an ideal out-of-cylinder speed and an ideal acceleration-time curve; Step 2: obtaining ideal pressure according to a dynamic equation; Step 3: obtaining ideal engine mass flow by substituting the ideal pressure into a control equation; Step 4: judging whether a difference between the ideal engine mass flow and an actual mass flow meets a requirement, if yes, returning to step 2, otherwise, entering step 5; The judgment of whether the difference between the ideal engine mass flow and the actual mass flow meets the requirement is specifically: let the difference between the ideal engine mass flow and the actual mass flow be , and a double-criterion judgment of whether the difference between the actual mass flow and the ideal mass flow meets the requirement is set: The first criterion is The ratio of the actual engine mass flow to the ideal engine mass flow is set as ; The second criterion is that the ratio of the initial gas flow rate of the single engine is set to ; and the ratio of the initial gas flow rate of the single engine is set to ; Setting a threshold value; If With Any one parameter is less than the set threshold, that is, to meet the requirements; if With All parameters are not less than the set threshold, that is, not to meet the requirements; Step 5: opening a new bottle to increase a release point; calculating an actual mass flow, and returning to step 2.

2. The method of claim 1, wherein the plurality of power sources are connected in series and the discrete pulse timing profile is fitted to the plurality of power sources. The step 2: obtaining ideal pressure according to a dynamic equation, specifically comprises: According to a formula wherein is the ideal pressure of the launch tube, is the ideal acceleration value at different times, is the mass of the missile, is the acceleration of gravity, is the friction coefficient between the missile and the wall of the tube, is the ambient pressure, is the cross-sectional area of the launch tube; Thus the ideal pressure at different times is obtained i is the time.

3. The method of claim 2, wherein the plurality of power sources are connected in series and the discrete pulse timing profile is fitted to the plurality of power sources. The step 3: obtaining ideal engine mass flow by substituting the ideal pressure into a control equation; wherein, represents the ideal gas mass flow at time i, represents the launch tube volume at time i, represents the engine working gas constant, is the ideal pressure inside the launch tube at the initial time, , is the ideal pressure inside the launch tube at the next time, is the temperature inside the launch tube; is the time length of each iteration step; is the launch tube cross-sectional area, is the missile velocity at the current time step.

4. The method of claim 3, wherein the plurality of power sources are connected in series and the discrete pulse timing profile is fitted to the plurality of power sources. the current time step of the missile velocity by the difference between the current iteration step acceleration and the previous time step acceleration , multiplied by the time step length.

5. The method of claim 1, wherein the plurality of power sources are connected in series and the discrete pulse timing profile is fitted to the plurality of power sources. The threshold value is set as any value within 0-1.

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