A method for calculating the aerodynamic characteristics of separation of rocket stage-to-stage fragments

Abstract

The present invention discloses a method for calculating the aerodynamic characteristics of separation of rocket inter-stage separation fragments, including: setting the initial state of the calculation of separation of rocket inter-stage separation fragments; updating the following parameters of each time step according to the time step: the radial distance between the center line of the fragment shell and the axis, the displacement of the upper stage, the displacement of the lower stage, the relative displacement between the upper stage and the lower stage, the partial pressure of the fuel gas, the partial pressure of the air, and the temperature of the mixed gas in the interstage, wherein the partial pressure of the fuel gas, the partial pressure of the air, and the temperature of the mixed gas in the interstage at the next time step are determined according to the gas volume and gas exhaust area in the interstage cavity formed by the upper stage, the lower stage, and the fragment separation body; judging whether the gas pressure in the interstage cavity is equal to the ambient pressure again, if so, stopping the calculation and outputting the result. The present invention solves the problem of effectively evaluating the separation aerodynamic force in the complex process of fragment movement exhaust under the joint action of multiple factors such as external flow, fragment separation, and upper stage jet.

Description

Technical Field

[0001] The invention relates to the technical field of multi-stage rocket engine devices, and in particular to a method for calculating aerodynamic characteristics of separation of separated fragments between rocket stages. Background Art

[0002] During the thermal separation process between the stages of the launch vehicle, the jet force is used as the separation impulse between the upper and lower stages, which can be separated quickly to reduce the loss of control time and the risk of collision. However, this solution will bring about a high-pressure and high-temperature environment in the interstage. For this reason, it is necessary to accurately provide the aerodynamic force and aerodynamic heat flow of the upper stage engine jet acting on the interstage during the separation process. In order to avoid the interstage pressure and thermal environment being too harsh, the interstage can be designed with an open flame exhaust method, that is, the interstage is designed as an open separation, and the rod structure interstage or the open flame exhaust window is used to relieve pressure. However, this solution will lead to an increase in the length and structural mass of the rocket body. In addition, the interstage can also be designed in a closed style, that is, the flame exhaust window is omitted, the upper stage engine is ignited first, and the plug is opened when the engine combustion chamber pressure reaches a certain level, and the separation signal is sent at the same time to achieve separation. Under the pressure of the interstage, the upper and lower stages can achieve a faster separation speed. In addition, compared with the open flame exhaust separation method, this solution omits a set of pyrotechnics in the flame exhaust window, and the length and structural mass of the cabin can be further reduced.

[0003] With the development of multi-stage rocket separation technology in recent years, a new interstage thermal separation technology in the form of grid cutting fragment separation has been formed. Its main principle is: after the upper stage engine is ignited, the interstage is cut into four pieces by using a pyrotechnic cutting device, and the four interstage shells are thrown out to the rocket circumference under the action of the cutting impulse and the interstage high-pressure gas generated by the engine, so as to release the constraints of the interstage shell on the nozzle start-up control, thereby achieving interstage separation "without loss of control". The network cutting fragment separation combines the advantages of both open flame separation and closed separation, mainly including: the flame exhaust area is significantly increased during the ejection process, which is conducive to rapid pressure relief and optimization of the load environment in the cabin; the pyrotechnic device for cutting the flame exhaust window is eliminated, reducing the structural complexity and the risks related to pyrotechnics, and improving the structural efficiency; the overlapping length of the separation body along the longitudinal direction is reduced by two transverse cuts, thereby shortening or even eliminating the pull-out distance of the upper stage nozzle, reducing the collision risk, and also reducing the nozzle swing restriction, which is conducive to the upper stage start-up control. The new interstage network cutting and separation technology has been applied in multiple arrow models. Practice has proved that this separation method can better adapt to the separation conditions under high dynamic pressure and high aftereffect thrust.

[0004] There are mature methods for calculating the interstage aerodynamic forces for the traditional open flame separation and closed separation. However, there is no reliable theoretical analysis basis and mature design method for the new grid-cut fragment separation method. Summary of the invention

[0005] The present invention provides a method for calculating the aerodynamic characteristics of separation of rocket inter-stage separation fragments. Through a theoretical approximate model of the grid-cut fragment separation exhaust process, combined with the non-fixed field flow field simulation of the upper and lower stages and the fragment separation motion process, the upper stage jet action and the aerodynamic force between stages in the inter-stage grid-cut fragment separation process are reasonably evaluated in the design stage. This method solves the problem of effectively evaluating the separation aerodynamic force in the complex process of fragment movement exhaust under the joint action of multiple factors such as external flow, fragment separation, and upper stage jet, and supports the design, optimization, and verification of the separation scheme of the model.

[0006] In a first aspect, a method for calculating the aerodynamic characteristics of separation of rocket stage separation fragments is provided, comprising:

[0007] Set the initial state of the rocket stage separation fragment separation calculation, including the initial state of the following parameters: the radial distance from the center line of the fragment shell to the axis , Upper stage displacement , the lower level displacement , relative displacement between the upper stage and the lower stage , partial pressure of gas , partial pressure of air and the mixed gas temperature in the interstage ;

[0008] According to the time step, update the following parameters for each time step: the radial distance from the center line of the fragment shell to the axis , Upper stage displacement , the lower level displacement , relative displacement between the upper stage and the lower stage , partial pressure of gas , partial pressure of air and the mixed gas temperature in the interstage , where the partial pressure of the gas at the next time step is , partial pressure of air and the mixed gas temperature in the interstage The gas volume in the interstage cavity formed by the upper stage, the lower stage and the fragment separation body and gas exhaust area Sure;

[0009] Determine whether the gas pressure in the interstage cavity is met Again with environmental pressure If they are equal, stop the calculation and output the result.

[0010] In combination with the first aspect, in certain implementations of the first aspect, the gas volume in the interstage cavity formed by the upper stage, the lower stage, and the fragment separation body is satisfy:

[0011]

[0012] in, is the initial value of the gas volume in the interstage cavity, is the axial length of the fragment, is the radius of the interstage shell;

[0013] The gas exhaust area in the interstage cavity composed of the upper stage, the lower stage and the fragment separation body satisfy:

[0014]

[0015] in, is the initial flame exhaust window area, .

[0016] In combination with the first aspect, in certain implementations of the first aspect, the partial pressure of the gas at the next time step is , partial pressure of air and the mixed gas temperature in the interstage The gas volume in the interstage cavity formed by the upper stage, the lower stage and the fragment separation body and gas exhaust area Confirm, satisfy:

[0017]

[0018]

[0019] is the density of the gas component, is the density of the air component, is the gas constant of the fuel gas, is the gas constant of air, is the number of upper stage engine nozzles, is the engine nozzle outlet flow rate, is the constant pressure specific heat of the mixture of fuel gas and air, is the constant volume specific heat of the mixed gas, is the total enthalpy of the gas in the engine combustion chamber, is the heat loss coefficient, is the gas constant of the mixed gas, is the gas exhaust flow rate, is the exhaust air flow rate, is the overall exhaust flow rate of the mixed gas, is the mass fraction of the fuel gas component, is the mass fraction of the air component, is the exhaust coefficient of the mixed gas.

[0020] In combination with the first aspect, in certain implementations of the first aspect, the exhaust coefficient of the mixed gas is Calculated as follows:

[0021]

[0022] in, is the exhaust coefficient; is the adiabatic index of the mixed gas itself, The critical pressure for distinguishing between under-expansion and full expansion.

[0023] In combination with the first aspect, in certain implementations of the first aspect, the critical pressure for distinguishing between under-expansion and full-expansion is Calculated as follows:

[0024] .

[0025] In combination with the first aspect, in certain implementations of the first aspect, the constant pressure specific heat of the mixed gas is Specific heat at constant volume and gas constant Calculated as follows:

[0026]

[0027] in, , are the constant volume specific heats of gas and air respectively, , are the constant pressure specific heats of gas and air respectively;

[0028] Mass fraction of gas components and the mass fraction of air components Calculated as follows:

[0029]

[0030] Density of the gas components in the mixed gas and the density of the air component is Calculated as follows:

[0031]

[0032] , They are the gas constants of fuel gas and air respectively, which can be determined from the gas physical property parameter table.

[0033] In combination with the first aspect, in some implementations of the first aspect, the upper stage displacement and the displacement of the following level and the relative displacement between the two satisfy:

[0034]

[0035] , are the masses of the upper and lower stages, respectively; is the axial thrust of the upper stage engine; is the reference pressure corresponding to the engine axial thrust value; is the number of upper stage engine nozzles, Upper stage engine nozzle exit area, is the axial cross-sectional area of ​​the rocket stage, is the aerodynamic drag of the rocket upper stage, It is the aftereffect thrust of the lower stage of the rocket.

[0036] In combination with the first aspect, in certain implementations of the first aspect, the radial distance between the center line of the fragment shell and the axis at the next time step is satisfy:

[0037]

[0038] in, is the inner area of ​​the fragment, The quality of the fragments.

[0039] In conjunction with the first aspect, in certain implementations of the first aspect, an initial state of the rocket stage separation fragment separation calculation is set to satisfy:

[0040] , is the radius of the interstage shell;

[0041] ;

[0042] ;

[0043] , is the atmospheric temperature at the separation point altitude.

[0044] In combination with the first aspect, in certain implementations of the first aspect, the output result also includes the aerodynamic force on the upper stage. and the aerodynamic force on the lower stage ,satisfy:

[0045]

[0046] in, is the axial thrust of the upper stage engine; is the reference pressure corresponding to the engine axial thrust value; is the number of upper stage engine nozzles, Upper stage engine nozzle exit area, is the axial cross-sectional area of ​​the rocket stage, is the aerodynamic drag of the rocket upper stage, It is the aftereffect thrust of the lower stage of the rocket.

[0047] Compared with the prior art, the solution provided by the present invention includes at least the following beneficial technical effects:

[0048] The present invention overcomes the bottleneck of existing methods that rely on unsteady flow field simulation and ground tests to obtain separation process parameters, and proposes an approximate model to model the gas volume and exhaust area in the interstage cavity during the exhaust process, thereby realizing the engineering calculation of the separation process under the influence of multiple factors such as external flow, fragment movement, and multi-component gas flow of gas and air, improving the design efficiency, meeting the deviation permission of engineering applications, and supporting the model's needs for separation scheme design, optimization, and verification and assessment.

[0049] Through the present invention, reliable and rapid engineering calculations can be performed on the exhaust and motion process of the new interstage thermal separation method of grid-cut fragment separation under the combined effects of external flow, fragment separation, and upper stage jet flow, which can support the design and optimization of separation schemes, etc., and save the investment and time cycle of numerical simulation and ground tests. This method has been verified by numerical simulation results and has been successfully applied in multiple launch vehicle models. BRIEF DESCRIPTION OF THE DRAWINGS

[0050] Figure 1 The present invention is a flow chart of a method for calculating the aerodynamic characteristics of separation of rocket stage-to-stage fragments. DETAILED DESCRIPTION

[0051] The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0052] like Figure 1 As shown, the present invention provides a method for calculating the aerodynamic characteristics of separation of rocket stage separation fragments, and the specific steps are as follows.

[0053] Step S1, setting the initial state of the rocket stage separation fragment separation calculation.

[0054] The calculation involves the radial distance between the center line of the fragment shell and the axis. The fragment is a square shell with an arc formed by cutting the interstage section with four longitudinal cutting seams and two circumferential cutting seams through pyrotechnics. Therefore, it is initially located on the cylindrical surface formed by the interstage section. The radial distance between the center line of the fragment shell and the axis is The radius of the interstage shell Consistent, that is .

[0055] The calculation involves the relative displacement between the upper stage and the lower stage. The initial time is generally selected as the separation start time, when the upper stage and the lower stage are in contact with each other, that is, .

[0056] The calculation involves , are the partial pressures of gas and air respectively, and their initial states are:

[0057]

[0058] The calculation involves is the temperature of the mixed gas in the interstage cavity, and its initial state is:

[0059]

[0060] Represents the atmospheric temperature at the separation point altitude, which can be determined by consulting the atmospheric altitude pressure-temperature table.

[0061] Step S2, according to the time step, loop through calculations S2.1 to S2.3, and update the following parameters for each time step: the radial distance from the centerline of the fragment shell to the axis , Upper stage displacement , the lower level displacement , relative displacement between the upper stage and the lower stage , partial pressure of gas , partial pressure of air and the mixed gas temperature in the interstage .

[0062] S2.1, calculate the radial distance of the fragment shell centerline from the axis.

[0063] In some embodiments, considering the movement of the fragment in the radial direction, the radial force depends on the pressure acting on the inner side of the shell. and external pressure , radial distance from the center line of the fragment shell to the axis The acceleration changes with time satisfy Newton's law. Specifically, the radial distance of the fragment shell centerline from the axis in the next time step is updated according to the following formula: :

[0064]

[0065] in, is the inner area of ​​the fragment, The quality of the fragments.

[0066] S2.2, calculate the displacement of the upper stage and the displacement of the following level and the relative displacement between the two .

[0067] Specifically, the displacement of the upper stage is calculated as follows: and the displacement of the following level and the relative displacement between the two :

[0068]

[0069] in, , are the masses of the upper stage and the lower stage respectively. is the axial thrust of the upper stage engine, which can be obtained from the upper stage engine internal ballistic parameter table. is the reference pressure corresponding to the engine axial thrust value, is the ambient pressure at the separation point, which can be obtained from the atmospheric altitude pressure parameter table. is the number of upper stage engine nozzles, Upper stage engine nozzle exit area, is the pressure inside the shell, is the axial cross-sectional area of ​​the rocket stage, is the aerodynamic drag of the rocket upper stage, obtained from the aerodynamic characteristic parameter table. It is the aftereffect thrust of the lower stage of the rocket, which is obtained from the engine internal ballistic parameter table.

[0070] In some embodiments, the separation aerodynamic forces on the upper stage and the lower stage can also be calculated. Considering the relative movement of the upper stage and the lower stage in the axial direction, the circumferential force depends on the force generated by the engine jet and the force of the gas pressure in the interstage section acting on the upper stage rear head.

[0071] Specifically, the aerodynamic force on the upper stage is calculated according to the following formula: and the aerodynamic force on the lower stage ;

[0072]

[0073] in, is the axial thrust of the upper stage engine, which can be obtained from the upper stage engine internal ballistic parameter table. is the reference pressure corresponding to the engine axial thrust value, is the ambient pressure at the separation point, which can be obtained from the atmospheric altitude pressure parameter table. is the number of upper stage engine nozzles, Upper stage engine nozzle exit area, is the axial cross-sectional area of ​​the rocket stage, is the aerodynamic drag of the rocket upper stage, obtained from the aerodynamic characteristic parameter table. It is the aftereffect thrust of the lower stage of the rocket, which is obtained from the engine internal ballistic parameter table.

[0074] S2.3, calculate the partial pressure of the gas by calculating the gas volume and gas exhaust area in the interstage cavity composed of the upper stage, lower stage and fragment separation body , partial pressure of air and the mixed gas temperature in the interstage .

[0075] The present invention adopts an approximate model in S2.3, which assumes that the fragments have only radial translation, and their circumferential translation and rotation around the center of mass of the fragment separation body can be ignored. The inner cavity of the interstage section is composed of two cones and one cylinder surrounded by the upper stage displacement, the lower stage displacement and the radial movement of the fragments.

[0076] S2.3.1, based on the above approximate model, calculate the gas volume in the interstage cavity consisting of the upper stage, the lower stage and the fragment separation body. According to the model assumptions, the following formula is used

[0077]

[0078] in, is the initial value of the gas volume in the interstage cavity, which can be determined by volume detection of the interstage three-dimensional model. is the axial length of the fragment.

[0079] S2.3.2, based on the above approximate model, calculate the gas exhaust area in the interstage cavity composed of the upper stage, the lower stage and the fragment separation body. Specifically, according to the model assumptions, the following formula is used:

[0080]

[0081] in, is the initial flame exhaust window area, which can be determined by detecting the approximate area of ​​the three-dimensional model of the interstage section. .

[0082] S2.3.3, establish a two-component mixed gas model of the engine jet fuel gas and air, and calculate the partial pressure of the fuel gas based on the exhaust process , partial pressure of air and the mixed gas temperature in the interstage .

[0083] Calculate the density, mass fraction, specific heat capacity, and adiabatic exponent of a two-component gas.

[0084] Density of the gas components in the mixed gas and the density of the air component is Calculated as follows:

[0085]

[0086] in , They are the gas constants of fuel gas and air respectively, which can be determined from the gas physical property parameter table.

[0087] Mass fraction of fuel gas components in the mixed gas and the mass fraction of air components Calculated as follows:

[0088]

[0089] The pressure of the gas mixture itself and density According to the following voltage division law:

[0090]

[0091]

[0092] Specific heat of the mixed gas at constant pressure Specific heat at constant volume and gas constant Calculated as follows:

[0093]

[0094] in, , are the constant volume specific heats of gas and air respectively, , They are the constant-pressure specific heats of fuel gas and air respectively, and both can be determined from the gas physical property parameter table.

[0095] Adiabatic index of the gas mixture itself Calculated as follows:

[0096]

[0097] According to the critical pressure that distinguishes underexpansion and full expansion Calculate the exhaust coefficient of the mixed gas .

[0098] Specifically, the critical pressure that distinguishes underexpansion from full expansion Calculated as follows:

[0099]

[0100] Exhaust coefficient of mixed gas Calculated as follows:

[0101]

[0102] in, is the exhaust coefficient, which can be determined by experiment or simulation, and can be taken as 0.9 in this embodiment.

[0103] According to the mixed gas exhaust coefficient Calculate the exhaust flow rate of each component. Specifically, the exhaust flow rate of the gas , Air exhaust flow And the overall exhaust flow of mixed gas Calculated as follows:

[0104]

[0105] Calculate the partial pressure of gas and air at the next time step based on the exhaust flow of gas and air components , and the mixed gas temperature in the interstage :

[0106]

[0107] is the engine nozzle outlet flow rate, is the total enthalpy of the gas in the engine combustion chamber. Both can be determined from the engine performance parameter table. is the heat loss coefficient, which is an empirical parameter and can be taken as 0.8 in this embodiment.

[0108] Step S3, determine whether the gas pressure in the interstage chamber meets Again with environmental pressure If they are equal, stop the calculation and output the result.

[0109] Specifically, the partial pressures of gas and air are obtained , and the mixed gas temperature in the interstage , the aerodynamic force on the upper stage and the aerodynamic force on the lower stage , and the upper stage displacement , the displacement of the following level and the relative displacement between the two , and get the final result.

[0110] Although the present invention is disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications without departing from the spirit and scope of the present invention. Therefore, the scope of protection of the present invention shall be based on the scope defined by the claims of the present invention.

Claims

1. A method for calculating the aerodynamic characteristics of separation of rocket stage separation fragments, characterized in that: include: Set the initial state of the rocket stage separation fragment separation calculation, including the initial state of the following parameters: the radial distance from the center line of the fragment shell to the axis , Upper stage displacement , the lower level displacement , relative displacement between the upper stage and the lower stage , partial pressure of gas , partial pressure of air and the mixed gas temperature in the interstage ; According to the time step, update the following parameters for each time step: the radial distance from the center line of the fragment shell to the axis , Upper stage displacement , the lower level displacement , relative displacement between the upper stage and the lower stage , partial pressure of gas , partial pressure of air and the mixed gas temperature in the interstage , where the partial pressure of the gas at the next time step is , partial pressure of air and the mixed gas temperature in the interstage The gas volume in the interstage cavity formed by the upper stage, the lower stage and the fragment separation body and gas exhaust area Confirm, satisfy: is the density of the gas component, is the density of the air component, is the gas constant of the fuel gas, is the gas constant of air, is the number of upper stage engine nozzles, is the engine nozzle outlet flow rate, is the constant pressure specific heat of the mixture of fuel gas and air, is the constant volume specific heat of the mixed gas, is the total enthalpy of the gas in the engine combustion chamber, is the heat loss coefficient, is the gas constant of the mixed gas, is the gas exhaust flow rate, is the exhaust air flow rate, is the overall exhaust flow rate of the mixed gas, is the mass fraction of the fuel gas component, is the mass fraction of the air component, is the exhaust coefficient of the mixed gas; Determine whether the gas pressure in the interstage cavity is met Again with environmental pressure If they are equal, stop the calculation and output the result.

2. The method according to claim 1, characterized in that: The gas volume in the interstage cavity composed of the upper stage, lower stage and fragment separation body satisfy: in, is the initial value of the gas volume in the interstage cavity, is the axial length of the fragment, is the radius of the interstage shell; The gas exhaust area in the interstage cavity composed of the upper stage, the lower stage and the fragment separation body satisfy: in, is the initial flame exhaust window area, .

3. The method according to claim 1, characterized in that Exhaust coefficient of mixed gas Calculated as follows: in, is the exhaust coefficient; is the adiabatic index of the mixed gas itself, The critical pressure for distinguishing between under-expansion and full expansion.

4. The method according to claim 3, characterized in that Critical pressure for distinguishing underexpansion from full expansion Calculated as follows: 。 5. The method according to claim 1, characterized in that: Specific heat of mixed gas at constant pressure Specific heat at constant volume and gas constant Calculated as follows: in, , are the constant volume specific heats of gas and air respectively, , are the constant pressure specific heats of gas and air respectively; Mass fraction of gas components and the mass fraction of air components Calculated as follows: Density of the gas components in the mixed gas and the density of the air component is Calculated as follows: , They are the gas constants of fuel gas and air respectively, which can be determined from the gas physical property parameter table.

6. The method according to claim 1, characterized in that Upper stage displacement and the displacement of the following level and the relative displacement between the two satisfy: , are the masses of the upper and lower stages, respectively; is the axial thrust of the upper stage engine; is the reference pressure corresponding to the engine axial thrust value; is the number of upper stage engine nozzles, Upper stage engine nozzle exit area, is the axial cross-sectional area of ​​the rocket stage, is the aerodynamic drag of the rocket upper stage, It is the aftereffect thrust of the lower stage of the rocket.

7. The method according to claim 1, characterized in that The radial distance between the centerline of the fragment shell and the axis at the next time step satisfy: in, is the inner area of ​​the fragment, The quality of the fragments.

8. The method according to claim 1, characterized in that Set the initial state of the rocket stage separation fragment separation calculation to satisfy: , is the radius of the interstage shell; ； ； , is the atmospheric temperature at the separation point altitude.

9. The method according to claim 1, characterized in that: The output also includes the aerodynamic forces acting on the upper stage. and the aerodynamic force on the lower stage ,satisfy: in, is the axial thrust of the upper stage engine; is the reference pressure corresponding to the engine axial thrust value; is the number of upper stage engine nozzles, Upper stage engine nozzle exit area, is the axial cross-sectional area of ​​the rocket stage, is the aerodynamic drag of the rocket upper stage, It is the aftereffect thrust of the lower stage of the rocket.

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