阅读说明：本技术 一种固体运载火箭入轨修正方法 (Solid carrier rocket in-orbit correction method ) 是由 艾之恒 舒畅 马超 于 2021-08-25 设计创作，主要内容包括：本申请公开了一种固体运载火箭入轨修正方法,包括：在三级发动机关机时,获取箭体的速度位置状态量；根据速度位置状态量及轨道参数信息确定滑行姿态指令和四级点火时间；在滑行调姿段根据滑行姿态指令对火箭进行姿态调整,直至调至四级定轴姿态；在四级点火时间到达时,发出四级点火指令,使得火箭进入四级主动段；根据发动机的内弹道曲线及发动机状态参数,计算四级主动段全程每间隔时间Δt-(4)的预测入轨轨道根数；通过预测入轨轨道根数与目标轨道根数计算获得四级姿态角指令的补偿值；根据补偿值对四级姿态角指令进行更新,直到四级发动机耗尽关机。本申请使得固体火箭发动机实现精准关机,由于不需要设置推力终止装置,降低了入轨成本。(The application discloses a solid carrier rocket in-orbit correction method, which comprises the following steps: when the three-stage engine is shut down, acquiring the speed position state quantity of the rocket body; determining a sliding attitude instruction and four-stage ignition time according to the speed position state quantity and the track parameter information; carrying out attitude adjustment on the rocket in the sliding attitude adjustment section according to the sliding attitude command until the rocket is adjusted to a four-stage dead axle attitude; when the fourth-stage ignition time is reached, a fourth-stage ignition instruction is sent out, so that the rocket enters a fourth-stage active section; according to the internal trajectory curve and the engine state parameters of the engine, calculating the whole-course interval time delta t of the four-stage active section 4 The predicted number of in-track tracks; calculating to obtain a compensation value of the four-level attitude angle instruction by predicting the number of the orbit entering tracks and the number of the target tracks; and updating the four-stage attitude angle command according to the compensation value until the four-stage engine is exhausted and shut down. The application enables the solid rocket engine to be accurately shut down, and the in-orbit cost is reduced because a thrust termination device is not required to be arranged.)

1. A method for correcting the in-orbit of a solid launch vehicle, the method comprising:

when the three-stage engine is shut down, acquiring the speed position state quantity of the rocket body;

determining a sliding attitude instruction and four-stage ignition time according to the speed position state quantity and the track parameter information; the track parameter information is track parameter information during four-level track entry;

carrying out attitude adjustment on the rocket in a sliding attitude adjusting section according to the sliding attitude instruction until the rocket is adjusted to a four-level dead axle attitude; the sliding attitude adjusting section is a section from the shutdown of the three-stage engine to the ignition of the four-stage transmitter;

when the fourth-stage ignition time is reached, a fourth-stage ignition instruction is sent out, so that the rocket enters a fourth-stage active section;

according to the internal trajectory curve and the engine state parameters of the engine, calculating the whole course of the four-stage active section at intervals delta t₄The predicted number of in-track tracks;

calculating to obtain a compensation value of the four-level attitude angle instruction by predicting the number of the orbit entering tracks and the number of the target tracks;

and updating the four-stage attitude angle command according to the compensation value until the four-stage engine is exhausted and shut down.

2. The solid launch vehicle in-orbit correction method of claim 1, wherein the fourth stage ignition timing is determined according to the following steps:

the speed position state quantity [ r ] under the earth center inertia system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]Track number a converted into guidance instruction for calculating time_sub、e_sub；

Calculating the apogee height R of the current sliding orbit of the rocket according to the following formula (I)_ap；

R_ap＝a_sub(1+e_sub) (ii) a (A)

Judgment of R_apWhen the track height is less than the target track, the equivalent pulse time t is determined according to the following formula group (II)_impAnd four-stage ignition time t_ig：

t_imp＝t_{i_ap}

R_ap＝a_sub(1+e_sub)

t_ig＝t_{i_ap}-t_core4

t_core4＝t₄-R_M4/W_M4(ii) a (II)

Wherein E_impIs the approximate point angle of the equivalent pulse point; r_orbIs the geocentric distance of the target track; t is t_core4Is a core time of four levels, t₄For four engine operating times, R_M4In four-level apparent displacement increments, W_M4The four-stage apparent velocity increment is shown, and E is a deviation from a point angle;

judgment of R_apWhen the track height is larger than or equal to the target track, determining the equivalent pulse time t according to the following formula (III)_impAnd four-stage ignition time t_ig：

t_ig＝t_imp-t_core4

t_core4＝t₄-R_M4/W_M4(ii) a (III)

If t is determined according to formula set (three)_impAnd t determined according to formula set (two)_{i_ap}Satisfy _ t_{i_ap}-t_imp≤R_M4/W_M4If so, let t_imp＝t_{i_ap}-R_M4/W_M4

Wherein E_impIs the approximate point angle of the equivalent pulse point; r_orbIs the geocentric distance of the target track.

3. The solid launch vehicle orbital correction method of claim 2, wherein the four-stage apparent displacement increment R_M4Four stage apparent velocity increment W_M4Calculated according to the following formula (iv):

in the formula, Isp₄Is the average specific impulse of a four-stage engine, m₄₀Starting mass at four ignition moments, m_4pIs four-level propellant charge m_4fThe mass of the arrow body left after the four-stage engine is shut down, T_s4For four engine operating times, k_m4The coefficients are four-level incremental correction coefficients.

4. The solid launch vehicle inbound correction method of claim 2 wherein said taxi attitude command is determined according to the steps of:

the rocket slides to four-stage equivalent pulse point time t under the action of gravity according to Kepler orbit recursion_impSpeed position state quantity r_sub，v_sub；

r_sub＝[r_xei，r_yei，r_zei]；

v_sub＝[v_xei，v_yei，v_zei]；

According to the state quantity r_sub，v_subNumber of tracks a_sub、e_subAnd the number of tracks of the target track a_orb、e_orb、i_orbDetermining the number of tracks of the transition track;

calculating the state quantity r of the transition orbit according to the orbit number of the transition orbit_orb，v_orb；

Calculating the required velocity increment v under the geocentric inertial system according to the following formula (five)_pa：

v_pa＝v_orb-v_sub(ii) a (V)

Increment the required speed v under the earth center inertia system_paConversion to required velocity increment v in launch train_x'、v_y'、v_z'；

Wherein v is_paxIs v is_paComponent in the x-direction, v_payIs v is_paComponent in the y-direction, v_pazIs v is_paComponent in the z direction, A₀For the launch azimuth of the rocket, B₀Geographical latitude of rocket launch point, E₀The earth center distance of the rocket launching point is L, and the direction cosine matrix is represented by L;

calculating a four-level attitude angle command according to the following formula set (six):

ψ_cmd＝-arcsinv_z′

γ_cmd0 (six);

wherein the content of the first and second substances,ψ_cmd、γ_cmdrespectively representing the pitching, yawing and rolling direction command attitude angles.

5. The solid launch vehicle orbital correction method of claim 1 wherein the predicted orbital number is determined according to the method of:

in the four-stage active stage flight process, at regular intervals of time delta t₄Obtaining the actual value of the current positionApparent velocity delta Δ W_M4And actual apparent displacement increment Δ R_M4；

According to the internal ballistic curve and the actual apparent speed increment delta W of the engine_M4Actual apparent displacement increment Δ R_M4Calculating to obtain the apparent speed increment and apparent displacement increment of the whole course₄Predicted apparent displacement delta ofPredicting apparent velocity delta

Based on predicted apparent displacement incrementsPredicting apparent velocity deltaAnd calculating the number of predicted orbit entering tracks after shutdown in the current state.

6. The solid launch vehicle in-orbit correction method of claim 5, wherein the compensation value for the four-stage attitude angle command is calculated according to the following method:

calculating compensation value of four-stage attitude angle instruction according to predicted number of in-orbit orbits and target orbit numberAnd delta phi_cmd。

7. The solid launch vehicle in-orbit correction method of claim 6, wherein the four-stage attitude angle command is updated with the offset value until the four-stage engine is shut down due to exhaustion according to the following method:

updating the four-level attitude angle command according to the following formula (seven):

γ_cmd0 (seven).

8. The method of correcting for solid launch vehicle orbital transfer of claim 2 wherein the state quantity of velocity position [ r ] in the centroidal inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]Calculated according to the following method:

acquiring the speed position state quantity [ x, y, z, vx, vy, vz ] of the rocket in a launching system;

the velocity position state quantity [ x, y, z, vx, vy, vz of the transmitting system]Converted into a speed position state quantity [ r ] of the earth-center fixed connection system_xef,r_yef,r_zef,v_xef,v_yef,v_zef]；

The velocity position state quantity [ r ] under the earth-core solid connection is expressed by the following formula (eight)_xef,r_yef,r_zef,v_xef,v_yef,v_zef]Converted into a speed position state quantity [ r ] under the geocentric inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]

[x_ef，y_ef，z_ef]Represents a position parameter of the earth-centered solid-connected system, [ x ]_ei，y_ei，z_ei]Showing groundPosition parameters under the condition of the cardiac inertia system.

Technical Field

The disclosure relates generally to the technical field of rockets, and in particular relates to a solid carrier rocket in-orbit correction method.

Background

The carrier rocket is generally composed of multiple stages, and the carrier rocket continuously overcomes the earth gravity to reach the state of height and speed required by orbit entering through the thrust action of boosting engines at all stages, so as to enter the corresponding earth orbit. In the process, the guidance system calculates and obtains a command attitude angle required by the rocket at present by comparing the current position speed state of the rocket with the position speed state required by the orbit entering, performs attitude adjustment by a control system executing mechanism, further sends time sequence commands of ignition, shutdown, separation and the like according to the states of different flight moments, further guides the rocket to fly to reach the position and speed state required by the orbit entering according to the command attitude angle, and finally realizes the accurate orbit entering of the satellite.

For the purpose of accurate tracking, the speed at the moment of tracking needs to be accurately controlled. In the process of the carrier rocket in orbit flight, the used boosting engines are mainly divided into two types, namely a solid engine and a liquid engine. Compared with a liquid engine, the solid engine has the characteristics of simple engine structure, short service cycle and the like, but cannot be accurately shut down like the liquid engine due to the characteristics of the solid engine.

For the rocket with a liquid engine at the stage of entering the orbit, iterative calculation can be directly carried out when the rocket approaches the state of entering the orbit through a certain guidance algorithm, and the engine is directly shut down after the condition is met, so that the accurate control of the speed of entering the orbit is achieved.

For a rocket with a solid engine at an entry stage, the current main solution is to start a thrust termination device by the thrust termination device similar to the shutdown of a liquid engine after the entry condition is met, so that the engine is shut down, and the aim of accurately controlling the entry speed is fulfilled. The other main solution is to add a set of liquid rail-controlled engine on the basis of the solid engine at the final stage, and after the final stage is shut down, the liquid rail-controlled engine is turned on to correct the rail-entering speed until the rail-entering speed requirement is met.

Due to the characteristics of the solid engine, after ignition, the charges in the solid engine can be continuously combusted and then are ejected out of the spray pipe to generate thrust, and the magnitude and the working process of the thrust can be described by an inner ballistic curve, wherein the inner ballistic curve expresses the process that the engine thrust changes along with the ignition time. Generally, the internal ballistic curve of a solid rocket engine can be obtained through design analysis and trial runs. According to the prior design result, the inner ballistic curve is generally influenced by temperature to cause the change of thrust and working time, and the inner ballistic deviation caused by the temperature is one of the main sources of the rocket body energy deviation in the actual flying process and is also the main reason of generating speed deviation at the last stage of the solid rocket.

The two solutions of adding a thrust termination device and adding a liquid rail control engine to the four-stage engine which are possible at present have the following main disadvantages:

the thrust termination device is added, a reverse spray pipe and initiating explosive devices need to be installed at the top of the engine shell, the original structure of the shell is affected, and the reliability of a product is reduced. Meanwhile, for the engine shell which is known by winding carbon fiber and other materials, the installation process is difficult to realize;

the addition of the liquid rail-controlled engine requires a series of additional equipment such as a gas cylinder, a storage tank, a conduit, a thrust chamber and the like at four stages, so that the complexity and the risk of the product are increased;

in addition to the increase of the cost of the two devices, the mass of the two devices directly occupies the whole transport capacity of the carrier rocket in a ratio of 1:1, which has a great influence on the benefit of the carrier rocket, and the phase change increases the cost.

Disclosure of Invention

In view of the above-mentioned drawbacks and deficiencies of the prior art, it is desirable to provide a method and system for reducing the speed error of the engine during shutdown to improve the accuracy of the rail entry without shutdown and end speed correction capability.

In a first aspect, the present application provides a method for solid launch vehicle in-orbit correction, the method comprising:

when the three-stage engine is shut down, acquiring the speed position state quantity of the rocket body;

determining a sliding attitude instruction and four-stage ignition time according to the speed position state quantity and the track parameter information; the track parameter information is track parameter information of four-level in-track;

carrying out attitude adjustment on the rocket in a sliding attitude adjusting section according to the sliding attitude instruction until the rocket is adjusted to a four-level dead axle attitude; the sliding attitude adjusting section is a section from the shutdown of the three-stage engine to the ignition of the four-stage transmitter;

when the fourth-stage ignition time is reached, a fourth-stage ignition instruction is sent out, so that the rocket enters a fourth-stage active section;

according to the internal trajectory curve and the engine state parameters of the engine, calculating the whole course of the four-stage active section at intervals delta t₄The predicted number of in-track tracks;

calculating to obtain a compensation value of the four-level attitude angle instruction by predicting the number of the orbit entering tracks and the number of the target tracks;

and updating the four-stage attitude angle command according to the compensation value until the four-stage engine is exhausted and shut down.

According to the technical scheme provided by the embodiment of the application, the four-stage ignition time is determined according to the following steps:

the speed position state quantity [ r ] under the earth center inertia system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]Track converted into calculation time of guidance instructionRoot number a_sub、e_sub；

Calculating the apogee height R of the current sliding orbit of the rocket according to the following formula (I)_ap；

R_ap＝a_sub(1+e_sub) (ii) a (A)

Judgment of R_apWhen the track height is less than the target track, the equivalent pulse time t is determined according to the following formula group (II)_impAnd four-stage ignition time t_ig：

t_imp＝t_{i_ap}

R_ap＝a_sub(1+e_sub)

t_ig＝t_{i_ap}-t_core4

t_core4＝t₄-R_M4/W_M4(ii) a (II)

Wherein E_impIs the approximate point angle of the equivalent pulse point; r_orbIs the geocentric distance of the target track; t is t_core4Is a core time of four levels, t₄For four engine operating times, R_M4In four-level apparent displacement increments, W_M4Is a four-stage apparent speed increment; e is a deviation from a point angle;

judgment of R_apWhen the track height is larger than or equal to the target track, determining the equivalent pulse time t according to the following formula (III)_impAnd four-stage ignition time t_ig：

t_ig＝t_imp-t_core4

t_core4＝t₄-R_M4/W_M4(ii) a (III)

Wherein E_impIs the approximate point angle of the equivalent pulse point; r_orbIs the geocentric distance of the target track.

If t is determined according to formula set (three)_impAnd t determined according to formula set (two)_{i_ap}Satisfy t_{i_ap}-t_imp≤R_M4/W_M4If so, let t_imp＝t_{i_ap}-R_M4/W_M4。

According to the technical scheme provided by the embodiment of the application, the four-level apparent displacement increment R_M4Four stage apparent velocity increment W_M4Calculated according to the following formula (iv):

in the formula, Isp₄Is the average specific impulse of a four-stage engine, m₄₀Starting mass at four ignition moments, m_4pIs four-level propellant charge m_4fThe mass of the arrow body left after the four-stage engine is shut down, T_s4For four engine operating times, k_m4The coefficients are four-level incremental correction coefficients.

According to the technical scheme provided by the embodiment of the application, the sliding attitude command is determined according to the following steps:

the rocket slides to four-stage equivalent pulse point time t under the action of gravity according to Kepler orbit recursion_impState quantity r of_sub，v_sub；

r_sub＝[r_xei，r_yei，r_zei]；

v_sub＝[v_xei，v_yei，v_zei]；

According to the state quantity r_sub，v_subNumber of tracks a_sub、e_subAnd the number of tracks of the target track a_orb、e_orb、i_orbDetermining the number of tracks of the transition track;

calculating the state quantity r of the transition orbit according to the orbit number of the transition orbit_orb，v_orb；

Calculating the required velocity increment v under the geocentric inertial system according to the following formula (five)_pa：

v_pa＝v_orb-v_sub(ii) a (V)

Increment the required speed v under the earth center inertia system_paConversion to required velocity increment v in launch train_x'、v_y'、v_z'；

Wherein v is_paxIs v is_paComponent in the x-direction, v_payIs v is_paComponent in the y-direction, v_pazIs v is_paComponent in the z direction, A₀For the launch azimuth of the rocket, B₀Geographical latitude of rocket launch point, E₀The earth center distance of the rocket launching point is L, and the direction cosine matrix is represented by L;

calculating a four-level attitude angle command according to the following formula set (six):

ψ_cmd＝-arcsinv′_z

γ_cmd0 (six)

Wherein the content of the first and second substances,ψ_cmd、γ_cmdrespectively representing the pitching, yawing and rolling direction command attitude angles.

According to the technical scheme provided by the embodiment of the application, the predicted number of the track-in tracks is determined according to the following method:

in the four-stage active stage flight process, at regular intervals of time delta t₄Obtaining the actual apparent velocity delta W generated by the current position_M4And actual apparent displacement increment Δ R_M4；

According to the internal ballistic curve and the actual apparent speed increment delta W of the engine_M4Actual apparent displacement increment Δ R_M4Calculating to obtain the apparent speed increment and apparent displacement increment of the whole course₄Predicted apparent displacement delta ofPredicting apparent velocity delta

Based on predicted apparent displacement incrementsPredicting apparent velocity deltaAnd calculating the number of predicted orbit entering tracks after shutdown in the current state.

According to the technical scheme provided by the embodiment of the application, the compensation value of the four-level attitude angle instruction is obtained by calculation according to the following method:

calculating compensation value of four-stage attitude angle instruction according to predicted number of in-orbit orbits and target orbit numberAnd delta phi_cmd。

According to the technical scheme provided by the embodiment of the application, the four-stage attitude angle command is updated through the compensation value according to the following method until the four-stage engine runs out and is shut down:

updating the four-level attitude angle command according to the following formula (seven):

γ_cmd0 (seven).

According to the technical scheme provided by the embodiment of the application, the speed position state quantity [ r ] under the geocentric inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]Calculated according to the following method:

acquiring the speed position state quantity [ x, y, z, vx, vy, vz ] of the rocket in a launching system;

the velocity position state quantity [ x, y, z, vx, vy, vz of the transmitting system]Converted into a speed position state quantity [ r ] of the earth-center fixed connection system_xef,r_yef,r_zef,v_xef,v_yef,v_zef]；

The state quantity [ r ] under the earth-heart solid connection is determined by the following formula (eight)_xef,r_yef,r_zef,v_xef,v_yef,v_zef]Converted into a state quantity [ r ] under the geocentric inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]

[x_ef，y_ef，z_ef]Represents a position parameter of the earth-centered solid-connected system, [ x ]_ei，y_ei，z_ei]The position parameters under the condition of geocentric inertial system are shown.

The beneficial effect of this application is: before the solid rocket engine enters the orbit, namely after the three-stage solid engine is shut down, the four-stage guidance instruction calculation work is started, at the moment, the rocket body is accelerated by three times of flight in the active section, has very high height and speed and slides to the height of the preset track, the state and the instruction are updated at intervals in the four-stage active section process, the energy deviation brought by the engine is eliminated, and therefore the precision is improved, the solid rocket engine can overcome the defect that the solid rocket engine cannot be shut down accurately like a liquid engine, the orbit entering precision is improved, a thrust termination device is not needed, and the orbit entering cost is reduced.

Drawings

Other features, objects and advantages of the present application will become more apparent upon reading of the following detailed description of non-limiting embodiments thereof, made with reference to the accompanying drawings in which:

FIG. 1 is a flow chart of example 1 of the present application;

FIG. 2 is a rocket state process diagram of embodiment 1 of the present application;

Detailed Description

The present application will be described in further detail with reference to the following drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the relevant invention and not restrictive of the invention. It should be noted that, for convenience of description, only the portions related to the present invention are shown in the drawings.

It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present application will be described in detail below with reference to the embodiments with reference to the attached drawings.

Example 1

The embodiment provides a solid carrier rocket in-orbit correction method, which comprises the following steps:

s10, acquiring the speed position state quantity of the rocket body when the three-stage engine is shut down;

s20, determining a sliding attitude command and four-stage ignition time according to the speed position state quantity and the track parameter information; the track parameter information is track parameter information of four-level in-track;

s30, adjusting the posture of the rocket in the sliding posture adjusting section according to the sliding posture instruction until the posture is adjusted to a four-stage dead axle posture; the sliding attitude adjusting section is a section from the shutdown of the three-stage engine to the ignition of the four-stage transmitter;

s40, when the four-stage ignition time is reached, a four-stage ignition instruction is sent out, so that the rocket enters a four-stage active section;

s50, calculating the whole process of the four-stage active section at each interval time delta t according to the internal ballistic curve of the engine and the engine state parameters₄The predicted number of in-track tracks;

s60, calculating to obtain a compensation value of the four-level attitude angle instruction through predicting the number of the orbit entering and the number of the target orbit;

and S70, updating the four-stage attitude angle command according to the compensation value until the four-stage engine is exhausted and shut down.

As shown in fig. 2, the starting point of the calculation of the present invention is after the three-stage solid engine is shut down, when the arrow body is accelerated through three active flight steps, and has a high altitude and speed, and slides to the altitude of the predetermined orbit. And at the moment, four-stage guidance instruction calculation is started, so that the solid rocket engine is shut down as the liquid rocket engine.

The four-level guidance instruction calculation in fig. 2 is the calculation of the four-level attitude angle instruction in the present application; the four-stage active segment engine state identification in fig. 2 refers to obtaining engine state parameters.

Wherein the four stage ignition timing is determined according to the following steps:

s21, determining the state quantity [ r ] of velocity position based on the inertia system of the earth' S center_xei,r_yei,r_zei,v_xei,v_yei,v_zei]And calculating to obtain the track number a of the guidance instruction at the calculation time_sub、e_sux；

Interconversion of state of the earth's center inertial system and orbital radical into prior art method, a_subDenotes the semi-major axis of the track, e_subRepresenting the eccentricity of the track.

Under the inertial coordinate system, the absolute state parameter of the rocket is [ r_a,v_a]＝[r_xei,r_yei,r_zei,v_xei,v_yei,v_zei]. For the representation of the track elements, it needs to be converted into intermediate quantities, indirectly describing:

orbital moment vector:

H＝r_a×v_a

lifting point vector:

N＝[0 0 0]^T×H＝[-H_X -H_Y 0]^T

eccentricity vector:

by these three vector expressions, the track element can be solved. However, the vector expression does not intuitively reflect the physical meaning of the track elements, and needs to be described in detail in a scalar form.

The orbital moment scalar equation is:

H＝||H||₂

the expression for orbital energy is:

according to the orbit momentum moment and the orbit energy, and then according to three vector expressions, the specific forms of six elements of the orbit can be obtained:

the semimajor axis is:

the half drift diameter is as follows:

the eccentricity ratio is:

e＝||e||₂

the inclination angle of the track is as follows:

the ascending right ascension:

the argument of the perigee is:

the approach point angle (rocket current orbit position t) is:

in the above, the track six-element specific expression has been expressed in detail. Under the geocentric inertial coordinate system, six elements of the track and a Cartesian state parameter r_a,v_aThe relationship (c) is expressed in detail, and the constraint of the terminal orbit element is used for converting into the constraint of the state parameter.

Velocity position state quantity [ r ] under the geocentric inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]Calculated according to the following method:

acquiring the speed position state quantity [ x, y, z, vx, vy, vz ] of the rocket in a launching system;

velocity of the transmitting systemDegree position state quantity [ x, y, z, vx, vy, vz]Converted into a speed position state quantity [ r ] of the earth-center fixed connection system_xef,r_yef,r_zef,v_xef,v_yef,v_zef]；

The velocity position state quantity [ r ] under the earth-core solid connection is expressed by the following formula (eight)_xef,r_yef,r_zef,v_xef,v_yef,v_zef]Converted into a speed position state quantity [ r ] under the geocentric inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]

[x_ef，y_ef，z_ef]Represents a position parameter of the earth-centered solid-connected system, [ x ]_ei，y_ei，z_ei]The position parameters under the condition of geocentric inertial system are shown.

Specifically, the method comprises the following steps:

according to the transmitting azimuth angle A₀

Longitude λ of transmission point₀

Elevation H of transmitting point₀

Geographic latitude B of transmitting point₀

Latitude of geocentric of emission point

Included angle between ground plumb line and center radius

Center-to-center distance of emission points

The emission point centroid radial is in the emission coordinate system component:

then the position state quantity [ x, y, z, vx, vy, vz ] can be determined according to the velocity of the transmitting system]Obtaining the speed position state quantity [ r ] of the earth-core solid connection system_xef,r_yef,r_zef,v_xef,v_yef,v_zef]

In the formula, a conversion matrix

Further fix the state quantity [ r ] of the earth core under the system_xef,r_yef,r_zef,v_xef,v_yef,v_zef]Converted into a state quantity [ r ] under the geocentric inertial system_xei,r_yei,r_zei,v_xei,v_yei,v_zei]The part adopts a simplified processing method, and is considered to be

S22, calculating the apogee height R of the current sliding orbit of the rocket according to the following formula (I)_ap；

R_ap＝a_sub(1+e_sub) (ii) a (A)

S23, judgment of R_apSmaller than target trackThe equivalent pulse time t is determined according to the following formula group (two)_impAnd four-stage ignition time t_ig：

t_imp＝t_{i_ap}

R_ap＝a_sub(1+e_sub)

t_ig＝t_{i_ap}-t_core4

t_core4＝t₄-R_M4/W_M4(ii) a (II)

Wherein E_impIs the approximate point angle of the equivalent pulse point; r_orbIs the geocentric distance of the target track; t is t_core4Is a core time of four levels, t₄For four engine operating times, R_M4In four-level apparent displacement increments, W_M4The four-stage apparent velocity increment is shown, and E is a deviation from a point angle;

judgment of R_apWhen the track height is larger than or equal to the target track, determining the equivalent pulse time t according to the following formula (III)_impAnd four-stage ignition time t_ig：

t_ig＝t_imp-t_core4

t_core4＝t₄-R_M4/W_M4(ii) a (III)

Wherein E_impIs the approximate point angle of the equivalent pulse point; r_orbIs the geocentric distance of the target track.

If t is determined according to formula set (three)_impAnd t determined according to formula set (two)_{i_ap}Satisfy t_{i_ap}-t_imp≤R_M4/W_M4If so, let t_imp＝t_{i_ap}-R_M4/W_M4。

Wherein the quaternary apparent displacement increment R_M4Four stage apparent velocity increment W_M4Calculated according to the following formula (iv):

in the formula, Isp₄Is the average specific impulse of a four-stage engine, m₄₀Starting mass at four ignition moments, m_4pIs four-level propellant charge m_4fThe mass of the arrow body left after the four-stage engine is shut down, T_s4For four engine operating times, k_m4The coefficients are four-level incremental correction coefficients.

Wherein the glide attitude command is determined according to the following steps:

s24, according to the Kepler orbit, the rocket slides to the four-stage equivalent pulse point time t under the action of gravity_impSpeed position state quantity r_sub，v_sub. Three position quantities and three speed quantities of the four-level equivalent pulse point:

r_sub＝[r_xei，r_yei，r_zei]；

v_sub＝[v_xei，v_yei，v_zei]；

s25, according to the state quantity r_sub，v_subNumber of tracks a_sub、e_subAnd the number of tracks of the target track a_orb、e_orb、i_orbDetermining the number of tracks of the transition track;

number of four-stage track a_orb、e_orb、i_orbAs known, for the rest orbit number, the geometric relationship of the orbit parameters can be known

u＝ω+f

The number of the remaining tracks of the transition track is calculated as follows

ω_4.inj＝u_4.inj-f_4.inj；

S26, calculating the state quantity r of the transition orbit according to the orbit number of the transition orbit_orb，v_orb；

The algorithm is the inverse operation of the state quantity calculation orbit root algorithm introduced above, and belongs to the known algorithm

Inputting: semimajor axis a, eccentricity e, orbit inclination angle i, ascension angle omega of ascending intersection point, argument omega of perigee and true perigee angle theta, mu is the gravity constant of the earth.

And (3) outputting: the state quantity of the earth inertia system comprises the position: x is the number of_EI、y_EI、z_EISpeed: v. of_EIx、v_EIy、v_EIz。

(1) Module for calculating specific angular momentum

(2) Calculating a position vector in a near focus coordinate system

(3) Calculating a velocity vector in a near-focus coordinate system

(4) Calculating a transformation matrix from a near-focus coordinate system to a geocentric equatorial coordinate system

(5) Calculating a position vector under a geostationary system

(6) Calculating velocity vector under geostationary system

S27, calculating the required velocity increment v under the geocentric inertial system according to the following formula (V)_pa：

v_pa＝v_orb-v_sub(ii) a (V)

S28, increasing the required speed v under the geocentric inertia system_paConversion to required velocity increment v in launch train_x'、v_y'、v_z'；

Wherein v is_paxIs v is_paComponent in the x-direction, v_payIs v is_paComponent in the y-direction, v_pazIs v is_paComponent in the z direction, A₀For the launch azimuth of the rocket, B₀Geographical latitude of rocket launch point, E₀Is the geocentric distance of the rocket launching point; l represents a directional cosine matrix;

and S29, calculating a four-stage attitude angle command according to the following formula group (six):

ψ_cmd＝-arcsinv′_z

γ_cmd0 (six);

wherein the content of the first and second substances,ψ_cmd、γ_cmdrespectively representing the pitching, yawing and rolling direction command attitude angles.

The track parameter information in the present embodiment therefore includes other parameters than the speed position state quantities involved in steps S21-S29.

According to the technical scheme provided by the embodiment of the application, the predicted number of the track-in tracks is determined according to the following method:

in the four-stage active stage flight process, at regular intervals of time delta t₄Obtaining the actual apparent velocity delta W generated by the current position_M4And actual apparent displacement increment Δ R_M4(ii) a Engine state parameterThe number is then referred to as the actual apparent speed increment Δ W of the current position of the engine_M4And actual apparent displacement increment Δ R_M4；

According to the internal ballistic curve and the actual apparent speed increment delta W of the engine_M4Actual apparent displacement increment Δ R_M4Calculating to obtain the apparent speed increment and apparent displacement increment of the whole course₄Predicted apparent displacement delta ofPredicting apparent velocity delta

Predicting apparent displacement incrementPredicting apparent velocity deltaAnd four-stage apparent displacement increment R_M4Four stage apparent velocity increment W_M4The calculation methods of (1) are identical, only the time of each parameter is replaced by the interval time delta t₄The latter time is not described herein;

based on predicted apparent displacement incrementsPredicting apparent velocity deltaCalculating the number of predicted orbit entering tracks after shutdown in the current state; the algorithm for predicting the number of the track-in tracks is consistent with the algorithm for predicting the number of the track-in tracks, and is not described herein.

In this embodiment, the compensation value of the four-level attitude angle command is obtained by calculation according to the following method:

calculating compensation value of four-stage attitude angle instruction according to predicted number of in-orbit orbits and target orbit numberAnd delta phi_cmd. The calculation method is the same as formula (six), and is not described herein.

According to the technical scheme provided by the embodiment of the application, the four-stage attitude angle command is updated through the compensation value according to the following method until the four-stage engine runs out and is shut down:

updating the four-level attitude angle command according to the following formula (seven):

γ_cmd0 (seven).

The above description is only a preferred embodiment of the application and is illustrative of the principles of the technology employed. It will be appreciated by a person skilled in the art that the scope of the invention as referred to in the present application is not limited to the embodiments with a specific combination of the above-mentioned features, but also covers other embodiments with any combination of the above-mentioned features or their equivalents without departing from the inventive concept. For example, the above features may be replaced with (but not limited to) features having similar functions disclosed in the present application.