A design method for lateral interference of thermal separation nozzle between rocket stages

Abstract

This invention discloses a design method for lateral interference of the nozzle during interstage thermal separation in rockets, relating to the technical field of multi-stage rocket engine devices. Addressing the lateral interference phenomenon of the nozzle during the pressurization process of interstage thermal separation in multi-stage rockets, this invention overcomes the bottleneck of existing design methods that rely on unsteady flow field simulations and ground tests to obtain the lateral interference force and moment characteristics of the nozzle. It proposes an approximate model to model the pressure distribution deviation inside and outside the nozzle during the interstage thermal separation process. This enables the calculation of the lateral interference force and moment of the nozzle during the separation process, considering the engine pressurization process, nozzle oscillation, and other factors. This improves design efficiency and supports the requirements of model design, optimization, and verification of separation schemes.

Description

Technical Field

[0001] The invention relates to the technical field of multi-stage rocket engine devices, in particular to a method for designing lateral interference of thermal separation nozzles between rocket stages. Background Art

[0002] During the interstage thermal separation process of a launch vehicle, the circumferential non-uniform state of the flow inside and outside the nozzle can lead to additional lateral aerodynamic interference torque acting on the nozzle. If the lateral interference force exceeds the tolerance limit of the nozzle servo, the servo will not be able to track the control instructions normally in a short time. However, this interference is generated at a stage where multiple factors such as interstage separation timing control, engine pressure buildup and plug opening, and upper stage nozzle control work together. The direction and magnitude of the torque change within a few tens of milliseconds of separation, and the process is extremely complicated. In the design of interstage thermal separation of traditional launch vehicles, a servo system with a large load capacity is used, which has sufficient anti-interference ability. In order to improve the carrying capacity, the new generation of launch vehicles often need to reduce the mass and power of the servo. Therefore, the design of the lower limit of the servo power selection is not only related to the separation and attitude control design schemes, but also requires quantitative anti-interference against the aerodynamic interference of the separation process.

[0003] There is currently no reliable theoretical analysis basis or mature design method for the above-mentioned aerodynamic interference torque. Summary of the invention

[0004] The present invention provides a method for designing the lateral interference of the hot separation nozzle between rocket stages. The present invention aims at the lateral interference phenomenon of the nozzle in the process of hot separation and pressure building between multi-stage rocket stages, overcomes the bottleneck of the existing design method that must rely on unsteady flow field simulation and ground test to obtain the lateral interference force and torque characteristics of the nozzle, and proposes an approximate model to model the pressure distribution deviation inside and outside the nozzle during the hot separation process between stages, and realizes the calculation of the lateral interference force and torque of the nozzle in the separation process under the influence of multiple factors such as the engine pressure building process, nozzle swing, etc., which improves the design efficiency and supports the needs of the model for separation scheme design, optimization and verification assessment.

[0005] In a first aspect, a method for designing lateral interference of a rocket interstage thermal separation nozzle is provided, comprising:

[0006] Calculate the cross-sectional radius corresponding to the shock wave surface position under typical conditions during the thermal separation process between rocket stages and the Mach number upstream of the shock wave ;

[0007] According to the flow parameters under typical conditions, the corresponding pressure downstream of the shock wave inside the nozzle is calculated. ;

[0008] Combined with the sinking angle of the upper stage rocket nozzle caused by the deformation of the engine pressure building process And preset swing angle and servo command swing angle , find out the actual swing range of the nozzle under typical conditions ;

[0009] The stationary shock wave shape in the nozzle is fitted into an elliptical plane perpendicular to the engine axis; according to the pressure downstream of the shock wave inside the nozzle and the actual swing range of the nozzle , calculate the lateral disturbance force generated by the pressure inside the nozzle and torque ;

[0010] According to the calculation method of aerodynamic characteristics of interstage separation process, the pressure change process of the gas in the interstage cavity during the separation process is calculated, and the gas pressure in the cavity outside the nozzle under typical conditions is obtained. ;

[0011] According to the actual swing range of the nozzle , the offset of the gap area between the nozzle outlet and the interstage shell relative to the uniform gap width is obtained; on this basis, the nozzle outer pressure under the influence of the offset is determined Distribution; Calculate the lateral interference force generated by the nozzle according to the pressure distribution outside the nozzle and torque ;

[0012] The lateral disturbance force generated by the pressure inside and outside the nozzle and , select the maximum value as the engineering design result of the nozzle lateral interference; according to the lateral interference torque generated by the pressure inside and outside the nozzle and , and select the maximum value as the engineering design result of the nozzle lateral interference.

[0013] In conjunction with the first aspect, in certain implementations of the first aspect, the typical state includes a flight test state and a ground test state;

[0014] For the flight test state, there is only one stationary shock wave from the nozzle throat of the multi-stage rocket upper stage to the front head of the multi-stage rocket lower stage, which is located at the nozzle exit. is the nozzle exit radius, and the corresponding shock wave upstream Mach number is the nozzle exit Mach number;

[0015] For the ground test state, there is only one stationary shock wave from the throat of the upper stage nozzle of the multi-stage rocket to the front head of the lower stage of the multi-stage rocket, which is located at 1 / 3~2 / 3 of the length from the throat to the nozzle outlet on the inside of the nozzle. The corresponding shock wave section radius The nozzle geometry is determined from the engine performance parameters, and the corresponding shock wave upstream Mach number Satisfies the following equation

[0016]

[0017] is the specific heat ratio coefficient of the fuel gas, which can be determined from the fuel gas thermophysical property parameter table in the engine performance parameters; is the radius of the rocket upper stage engine nozzle throat.

[0018] In combination with the first aspect, in certain implementations of the first aspect, the pressure downstream of the shock wave inside the nozzle is satisfy:

[0019]

[0020] is the Mach number upstream of the shock wave; is the gas specific heat ratio coefficient; is the engine combustion chamber pressure.

[0021] In combination with the first aspect, in certain implementations of the first aspect, the sinking angle and engine combustion chamber pressure The relationship satisfies:

[0022]

[0023] in, is the proportional coefficient of the sinking angle to the combustion chamber pressure;

[0024] Actual swing range of the nozzle satisfy:

[0025] .

[0026] In combination with the first aspect, in certain implementations of the first aspect, the stationary shock wave is fitted into an ellipse formed by the intersection of a plane perpendicular to the engine axis and the engine nozzle; in the elliptical plane stationary shock wave form, the high-pressure area formed downstream of the shock wave acts on the inner surface of the nozzle and forms an asymmetric distribution.

[0027] In combination with the first aspect, in certain implementations of the first aspect, the lateral interference force generated by the pressure inside the nozzle and torque satisfy:

[0028]

[0029] in, is the radius of the nozzle at the approximate stationary shock wave plane under typical conditions, is the axial distance from the center of the approximate stationary shock wave plane to the nozzle swing center.

[0030] In combination with the first aspect, in certain implementations of the first aspect, according to the actual swing angle range of the nozzle , calculate the offset of the gap width between the nozzle outlet and the interstage shell wall relative to the uniform gap width, satisfying:

[0031]

[0032] is the gap width between the nozzle outlet and the interstage shell wall. When the actual nozzle swing angle is 0°, is the uniform gap width between the nozzle outlet and the interstage shell wall, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, with a value range of 0~360°; is the displacement length at the nozzle outlet due to the swing, , It is the distance from the nozzle outlet center to the nozzle swing center.

[0033] In combination with the first aspect, in certain implementations of the first aspect, the nozzle outer side pressure satisfy:

[0034]

[0035] is the uniform gap width between the nozzle outlet and the interstage shell wall, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, with a value range of 0~360°; is the displacement length at the nozzle outlet due to the swing, , It is the distance from the nozzle outlet center to the nozzle swing center.

[0036] In combination with the first aspect, in some implementations of the first aspect, the lateral interference force Lateral disturbance torque satisfy:

[0037]

[0038]

[0039] is the axial distance from the nozzle swing center The axial section radius at is the uniform gap width between the nozzle outlet and the interstage shell wall, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, with a value range of 0~360°; is the displacement length at the nozzle outlet due to the swing, , is the distance from the nozzle outlet center to the nozzle swing center, It is the distance from the rear head flange to the center of the swing.

[0040] In combination with the first aspect, in some implementations of the first aspect, the lateral interference force Lateral disturbance torque satisfy:

[0041]

[0042] is the lateral projection area of ​​the nozzle from the rear head flange to the outlet, is the centroid of the lateral section of the nozzle from the rear head flange to the outlet relative to the swing center, is the uniform gap width between the nozzle outlet and the interstage shell wall, is the displacement length at the nozzle outlet due to the swing, , It is the distance from the nozzle outlet center to the nozzle swing center.

[0043] Compared with the prior art, the solution provided by the present invention includes at least the following beneficial technical effects:

[0044] The present invention adopts an approximate model to model the deviation of the pressure distribution inside and outside the nozzle during the interstage thermal separation process, thereby realizing the calculation of the lateral interference force and torque of the nozzle during the separation process under the influence of multiple factors such as the engine pressure building process, nozzle swing, etc., and meeting the deviation allowance of engineering applications, thereby improving the design efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Figure 1 Schematic diagram of engine swing angle.

[0046] Figure 2 Schematic diagram of asymmetric pressure distribution in the nozzle.

[0047] Figure 3 Schematic diagram of gap deviation.

[0048] Figure 4 The present invention is a flow chart of a design method for lateral interference of thermal separation nozzles between rocket stages. DETAILED DESCRIPTION

[0049] The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0050] like Figure 4As shown, the present invention provides a method for designing lateral interference of thermal separation nozzles between rocket stages.

[0051] Step S1, calculating the cross-sectional radius corresponding to the shock wave surface position under typical conditions during the thermal separation process between rocket stages and the Mach number upstream of the shock wave Typical states include flight test state and ground test state.

[0052] Typical state 1: For the flight test state, there is only one stationary shock wave from the nozzle throat of the upper stage of the multi-stage rocket to the front head of the lower stage of the multi-stage rocket, and its position is exactly at the nozzle outlet. is selected as the nozzle exit radius, and the corresponding shock wave upstream Mach number It is the nozzle outlet Mach number, which can be determined from the gas flow outlet state parameters in the engine performance parameters.

[0053] Typical state 2: For the ground test state, there is only one stationary shock wave from the throat of the upper stage of the multi-stage rocket to the front head of the lower stage of the multi-stage rocket. Its position is inside the nozzle and can be determined based on design experience or numerical simulation. This method recommends that the shock wave position be selected at 1 / 3~2 / 3 of the length from the throat to the nozzle outlet. The corresponding shock wave section radius The nozzle geometry can be determined from the engine performance parameters. The corresponding shock wave upstream Mach number can be determined based on the isentropic flow parameters of the complete gas in the variable cross-section pipe. Specifically, the shock wave upstream Mach number Satisfies the following equation

[0054]

[0055] is the fuel gas specific heat ratio coefficient, which can be determined from the fuel gas thermophysical property parameter table in the engine performance parameters. is the radius of the rocket upper stage engine nozzle throat.

[0056] Step S2, calculate the corresponding nozzle inner shock wave downstream pressure according to the flow parameters under the typical state .

[0057] This method proposes a method to calculate the pressure of the high-pressure area inside the nozzle by using the total pressure loss relationship before and after the shock wave. Specifically, according to the total pressure loss relationship before and after the shock wave, the pressure of the high-pressure area downstream of the shock wave inside the nozzle is It can be expressed as

[0058]

[0059] in, is the shock wave upstream Mach number determined in step S1; is the specific heat ratio coefficient of the fuel gas, which can be determined from the fuel gas thermophysical property parameter table in the engine performance parameters; It is the engine combustion chamber pressure, which can be determined from the engine rising section performance parameters.

[0060] Step S3, according to the engine combustion chamber pressure under the typical state calculated in step S2, determine the sinking angle of the upper stage rocket nozzle caused by the deformation of the engine pressure building process , combined with the preset swing angle and servo command swing angle , find out the actual swing range of the nozzle under typical conditions .

[0061] Specifically, the engine preset swing angle is a known parameter, the sinking angle and engine combustion chamber pressure The relationship between the engine performance parameters can be obtained by looking up the engine performance parameters. In this embodiment, the following formula is used according to the engine combustion chamber pressure Calculate the sink angle :

[0062]

[0063] in, It is the proportional coefficient of the sinking angle-combustion chamber pressure, which can be obtained by looking up the engine performance parameters.

[0064] During the rocket separation process, the control system issues a servo command for the nozzle to swing, and the corresponding servo command swing angle The value will be within a certain range, and the specific range can be extracted from the simulation results of the attitude control system.

[0065] Consider the preset swing angle , sink angle and servo command swing angle , the actual swing angle of the nozzle is obtained by superposition , which is calculated according to the following formula

[0066]

[0067] The unit of all variables is angle (°). It is usually a negative value. When the engine reaches the pressure plateau, the sinking angle Set and preset the swing angle Offset each other.

[0068] In this embodiment, for the typical state described in step S1, the corresponding engine combustion chamber pressure is selected , and determine the servo command swing angle In the value range of the corresponding state, the actual swing angle of the nozzle in this state is obtained At this actual swing angle, the nozzle axis and the engine axis form a value of The angle of Figure 1 shown.

[0069] Step S4, fit the stationary shock wave shape in the nozzle into an elliptical plane perpendicular to the engine axis (note that it is not the nozzle axis). According to the shock wave downstream parameters calculated in step S2 and the actual nozzle swing angle obtained in step S3, calculate the lateral interference force and torque generated by the pressure inside the nozzle.

[0070] Specifically, the stationary shock wave is fitted as an ellipse formed by the intersection of a plane perpendicular to the engine axis and the engine nozzle. When , that is, the angle between the engine axis and the nozzle axis is When the above approximate stationary shock wave is located on the plane and the nozzle axial plane, a corresponding size of Angle.

[0071] Furthermore, under the above approximate stationary shock wave model, the high pressure area formed downstream of the shock wave acts on the inner surface of the nozzle and forms an asymmetric distribution, such as Figure 2 shown.

[0072] Under this asymmetric distribution, according to the shock wave downstream pressure calculated in step S2 , calculate the lateral interference force generated by the pressure inside the nozzle according to the following formula: and torque

[0073]

[0074] in, is the radius of the nozzle at the approximate stationary shock wave plane under the typical state described in step S1, is the axial distance from the center of the approximate stationary shock wave plane to the nozzle swing center.

[0075] Step S5, according to the aerodynamic characteristics calculation method of the interstage separation process (this method is consistent with the aerospace standard), calculate the pressure change process of the gas in the interstage cavity during the separation process, and obtain the gas pressure in the cavity outside the nozzle under the typical state described in step S1 .

[0076] Step S6, based on the actual swing range of the nozzle calculated in step S3, obtain the offset of the gap area between the nozzle outlet and the interstage shell wall relative to the uniform gap area. On this basis, a gap height-pressure approximation model based on linear proportion is used to model the offset of the pressure distribution outside the nozzle relative to the uniform distribution. According to the pressure distribution outside the nozzle, the lateral interference force and torque generated by it are calculated.

[0077] Sub-step S61, according to the actual swing angle range of the nozzle, calculate the offset of the gap width between the nozzle outlet and the interstage shell wall relative to the uniform gap width, such as Figure 3 shown.

[0078] Specifically, when the actual nozzle swing angle is 0°, the gap width between the nozzle outlet and the interstage shell wall is uniform along the circumferential direction. This uniform gap width is recorded as , which can be measured and obtained from the interstage 3D model using 3D modeling software.

[0079] When the actual nozzle swing angle is When the gap width between the nozzle outlet and the interstage shell wall is It is non-uniformly distributed along the circumference and can be calculated as follows

[0080]

[0081] in, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, and the value range is 0~360°. is the displacement length at the nozzle outlet caused by the swing, calculated according to the following formula:

[0082]

[0083] in, It is the distance between the nozzle outlet center and the nozzle swing center, which can be measured from the interstage 3D model using 3D modeling software.

[0084] When the nozzle actual swing angle When the range of values ​​described in step S3 is used, the gap width between the nozzle outlet and the interstage shell wall is In this embodiment, the actual nozzle swing angle can be adjusted at intervals of 0.1°. Take a set of discrete values ​​and calculate the displacement length of the corresponding nozzle outlet due to swing , providing subsequent steps in a list format.

[0085] Sub-step S62, according to the width of the gap between the nozzle outlet and the interstage shell wall The mathematical formula for the nozzle outer pressure at the corresponding position is The calculation can be approximately performed using the gap height-pressure linear proportional model, which can be expressed as the following formula:

[0086]

[0087] Under this nozzle outer side pressure distribution, the generated nozzle lateral interference force and the lateral disturbance torque relative to the nozzle swing center It can be expressed as the following formula after derivation:

[0088]

[0089] in, is the axial distance from the nozzle swing center The axial section radius at is the lateral projection area of ​​the nozzle from the rear head flange to the outlet, is the distance from the rear head flange to the center of the swing, It is the centroid of the lateral section of the nozzle from the rear head flange to the outlet relative to the swing center. The above geometric parameters can be measured and obtained from the three-dimensional model of the jet engine nozzle through three-dimensional modeling software.

[0090] Step S7, based on the lateral interference force and torque generated by the pressures on the inside and outside of the nozzle obtained in step S4 and step S6, the maximum value between the two is selected as the engineering design result of the lateral interference of the nozzle.

[0091] Specifically, the lateral interference force generated by the pressure inside and outside the nozzle and the lateral disturbance torque relative to the nozzle swing center It can be expressed as

[0092]

[0093] 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 rocket interstage thermal separation nozzle lateral interference design method, characterized in that: include: Calculate the cross-sectional radius corresponding to the shock wave surface position under typical conditions during the thermal separation process between rocket stages and the Mach number upstream of the shock wave ; According to the flow parameters under typical conditions, the corresponding pressure downstream of the shock wave inside the nozzle is calculated. ; Combined with the sinking angle of the upper stage rocket nozzle caused by the deformation of the engine pressure building process And preset swing angle and servo command swing angle , find out the actual swing range of the nozzle under typical conditions ; The stationary shock wave shape in the nozzle is fitted into an elliptical plane perpendicular to the engine axis; according to the pressure downstream of the shock wave inside the nozzle and the actual swing range of the nozzle , calculate the lateral disturbance force generated by the pressure inside the nozzle and torque , lateral disturbance force and torque satisfy: in, is the radius of the nozzle at the approximate stationary shock wave plane under typical conditions, is the axial distance from the center of the approximate stationary shock wave plane to the nozzle swing center; According to the calculation method of aerodynamic characteristics of interstage separation process, the pressure change process of the gas in the interstage cavity during the separation process is calculated, and the gas pressure in the cavity outside the nozzle under typical conditions is obtained. ; According to the actual swing range of the nozzle , the offset of the gap area between the nozzle outlet and the interstage shell relative to the uniform gap width is obtained; on this basis, the nozzle outer pressure under the influence of the offset is determined Distribution; Calculate the lateral interference force generated by the nozzle according to the pressure distribution outside the nozzle and torque ; The lateral disturbance force generated by the pressure inside and outside the nozzle and , select the maximum value as the engineering design result of the nozzle lateral interference; according to the lateral interference torque generated by the pressure inside and outside the nozzle and , and select the maximum value as the engineering design result of the nozzle lateral interference.

2. The method according to claim 1, characterized in that: Typical states include flight test state and ground test state; For the flight test state, there is only one stationary shock wave from the nozzle throat of the multi-stage rocket upper stage to the front head of the multi-stage rocket lower stage, which is located at the nozzle exit. is the nozzle exit radius, and the corresponding shock wave upstream Mach number is the nozzle exit Mach number; For the ground test state, there is only one stationary shock wave from the throat of the upper stage nozzle of the multi-stage rocket to the front head of the lower stage of the multi-stage rocket, which is located at 1 / 3~2 / 3 of the length from the throat to the nozzle outlet on the inside of the nozzle. The corresponding shock wave section radius The nozzle geometry is determined from the engine performance parameters, and the corresponding shock wave upstream Mach number Satisfies the following equation is the specific heat ratio coefficient of the fuel gas, which can be determined from the fuel gas thermophysical property parameter table in the engine performance parameters; is the radius of the rocket upper stage engine nozzle throat.

3. The method according to claim 1, characterized in that Pressure downstream of the shock wave inside the nozzle satisfy: is the Mach number upstream of the shock wave; is the gas specific heat ratio coefficient; is the engine combustion chamber pressure.

4. The method according to claim 1, characterized in that: Sinking Angle and engine combustion chamber pressure The relationship satisfies: in, is the proportional coefficient of the sinking angle-combustion chamber pressure; Actual swing range of the nozzle satisfy: 。 5. The method according to claim 1, characterized in that The stationary shock wave is fitted as an ellipse formed by the intersection of a plane perpendicular to the engine axis and the engine nozzle; in the elliptical plane stationary shock wave form, the high-pressure area formed downstream of the shock wave acts on the inner surface of the nozzle and forms an asymmetric distribution.

6. The method according to claim 1, characterized in that According to the actual swing angle range of the nozzle , calculate the offset of the gap width between the nozzle outlet and the interstage shell wall relative to the uniform gap width, satisfying: is the gap width between the nozzle outlet and the interstage shell wall. When the actual nozzle swing angle is 0°, is the uniform gap width between the nozzle outlet and the interstage shell wall, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, with a value range of 0~360°; is the displacement length at the nozzle outlet due to the swing, , It is the distance from the nozzle outlet center to the nozzle swing center.

7. The method according to claim 1, characterized in that Nozzle outer pressure satisfy: is the uniform gap width between the nozzle outlet and the interstage shell wall, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, with a value range of 0~360°; is the displacement length at the nozzle outlet due to the swing, , It is the distance from the nozzle outlet center to the nozzle swing center.

8. The method according to claim 1, characterized in that: Lateral interference force Lateral disturbance torque satisfy: is the axial distance from the nozzle swing center The axial section radius at is the uniform gap width between the nozzle outlet and the interstage shell wall, Represents the circumferential azimuth of different positions at the nozzle exit relative to the reference quadrant, with a value range of 0~360°; is the displacement length at the nozzle outlet due to the swing, , is the distance from the nozzle outlet center to the nozzle swing center, It is the distance from the rear head flange to the center of the swing.

9. The method according to claim 1, characterized in that: Lateral interference force Lateral disturbance torque satisfy: is the lateral projection area of ​​the nozzle from the rear head flange to the outlet, is the centroid of the lateral section of the nozzle from the rear head flange to the outlet relative to the swing center, is the uniform gap width between the nozzle outlet and the interstage shell wall, is the displacement length at the nozzle outlet due to the swing, , It is the distance from the nozzle outlet center to the nozzle swing center.

Images