CN112325711A - Carrier rocket orbit height keeping control method based on balanced flight theory - Google Patents

Carrier rocket orbit height keeping control method based on balanced flight theory Download PDF

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CN112325711A
CN112325711A CN202011162810.9A CN202011162810A CN112325711A CN 112325711 A CN112325711 A CN 112325711A CN 202011162810 A CN202011162810 A CN 202011162810A CN 112325711 A CN112325711 A CN 112325711A
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孟云鹤
连一君
罗宗富
陈琪锋
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Shenzhen 303 Defense Technology Co ltd
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
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Abstract

The invention discloses a carrier rocket orbit height keeping control method based on a balance flight theory, which comprises the following steps of firstly, carrying out autonomous guidance control on a rocket based on the balance theory (including modes of balance flight, quasi-balance flight and the like); secondly, acquiring an acceleration radial additional control quantity; and finally, acquiring an acceleration dip angle additional control quantity according to the acceleration radial additional control quantity, and then acquiring a thrust acceleration dip angle considering the track height maintaining control. The control method has the characteristics of high precision and small calculation amount, can meet the engineering application requirements of the rocket-borne computer on real-time calculation, and is high in practicability.

Description

Carrier rocket orbit height keeping control method based on balanced flight theory
Technical Field
The invention relates to the technical field of aerospace, in particular to a carrier rocket orbit height keeping control method based on a balanced flight theory.
Background
The normal flight procedure of the carrier rocket is strictly time-dependent, and the precise time sequence design is carried out according to the factors of the flight environment, rocket body structure, fuel consumption, power characteristics, effective load and the like of the rocket and by comprehensively considering the target orbit of the carrier mission, so that the time sequence of the action logic of each executing mechanism is controlled, and the flight procedure needs to be revised every time the launching mission is executed. Because of the need of satisfying many engineering constraints, this calculation process is very complicated, the calculated amount is very large, and the real-time requirement cannot be satisfied, and usually, the firing data is injected before the rocket is fired.
The rocket inevitably encounters various faults in the process of executing tasks, and the occurrence time and the fault mode are uncertain. When the rocket has sudden thrust failure, the rocket flies according to a preset program to cause task failure, for example, in the flight of a Long-March five-number remote second carrier rocket, the thrust of a core first-stage engine is instantaneously and greatly reduced, so that the rocket cannot reach the preset flying speed and height, and finally, a second-stage rocket and a satellite enter in the Western Pacific again to launch a task and lose profits. If the flight program can be adjusted timely, the loss of the mission can be avoided, for example, the Tuxing 5 carrier rocket carries the Apollo 13 airship, the second-stage main engine of the rocket shuts down 132 seconds in advance due to reasons, the other 4 engines work for 34 seconds in a compensatory way, and the airship smoothly enters the lunar orbit. At present, most of the existing rockets in China do not have the capabilities of real-time fault detection, fault-tolerant processing and redundancy reconstruction, and once major abnormalities such as power system faults occur in the flight process, the coping strategies cannot be executed autonomously, so that the development of the intelligent rocket technology is very urgent.
The traditional carrier rocket guidance method adopts a perturbation guidance or track tracking mode, namely a standard trajectory is designed offline in advance, when the carrier rocket actually flies, a guidance control system controls the actual flight trajectory of the carrier rocket to perturb near the standard trajectory, so that the actual flight trajectory is attached to the standard trajectory as far as possible. However, the guidance method is low in fault tolerance, when the guidance method encounters a thrust abnormal fault, the performance of the carrier rocket is reduced, and sufficient power cannot be generated to continue tracking program ballistic flight, so that the actual flight trajectory greatly deviates from the standard ballistic trajectory, and even serious consequences such as rocket instability and the like may occur.
Therefore, based on the design margin and the orbit entering capability of the carrier rocket in the aerospace system, it is of great significance to research one of the key points (orbit height maintenance control problem) of autonomous intelligent fault handling of the carrier rocket aiming at unexpected situations that the carrier rocket encounters non-fatal engine faults and the like with abnormal thrust.
Disclosure of Invention
The invention provides a carrier rocket orbit height keeping control method based on a balanced flight theory, which has the following specific technical scheme:
a carrier rocket orbit height keeping control method based on a balanced flight theory comprises the following steps:
the method comprises the following steps of firstly, carrying out autonomous guidance control on the rocket, specifically:
in the balanced flight process, the optimal guidance law theta is obtained through the expression 3)PThe maneuvering orbital transfer time T from the elliptical trajectory to the circular orbit of the equilibrium flight is obtained by expression 6):
Figure BDA0002744785150000021
Figure BDA0002744785150000022
wherein: r represents the geocentric distance; μ represents an earth gravity coefficient; n is the flight angular velocity of the target circular orbit,
Figure BDA0002744785150000023
omega is the angular velocity of the range,
Figure BDA0002744785150000024
t represents time; a iscThrust acceleration of the rocket;
in the quasi-equilibrium flight process, the transition time delta T from quasi-equilibrium flight to equilibrium flight is obtained through an expression 18), and the optimal guidance law theta is obtained through an expression 21)P：
Figure BDA0002744785150000025
Figure BDA0002744785150000026
Wherein: r is0Representing an initial ground center distance of quasi-equilibrium flight; v. ofθIs the circumferential velocity component at the moment of failure,
Figure BDA0002744785150000027
step two, acquiring an acceleration radial additional control quantity delta a through an expression 24)c_r：
Figure BDA0002744785150000031
Wherein: k is a radical of1And k2For the control parameters of the height feedback control law, take
Figure BDA0002744785150000032
Figure BDA0002744785150000033
Is a natural frequency; xi is a damping ratio; Δ r is the deviation of the actual ground center distance and the target ground center distance;
Figure BDA0002744785150000034
is the rate of change of the deviation of the actual ground center distance from the target ground center distance;
step three, acquiring the acceleration inclination through an expression 30)Angle additional control quantity delta thetaPAnd then the thrust acceleration inclination angle theta considering the track height holding control is obtainedP+ΔΘP：
Figure BDA0002744785150000035
Preferably, in the above technical solution, the autonomous guidance of the rocket in the step one is specifically:
step a1, carrying out dynamic analysis on rocket motion, identifying the power system fault of the rocket, and acquiring thrust acceleration a under the fault modecA function of time of;
step a2, thrust acceleration a to rocketcThe judgment is specifically as follows:
if the thrust acceleration a of the rocketcExpression 12 is satisfied), the rocket enters a balanced flight state:
Figure BDA0002744785150000036
wherein: g0Is the acceleration of the gravity of the circular orbit,
Figure BDA0002744785150000037
if the thrust acceleration a of the rocketcExpression 15) is satisfied but expression 12 is not satisfied), the rocket enters a quasi-balanced flight state;
Figure BDA0002744785150000038
wherein: Δ h is the height margin;
if the thrust acceleration a of the rocketcExpression 15 is not satisfied), the rocket enters the atmosphere to crash;
step a3, carrying out autonomous guidance according to the condition that the rocket is in balanced flight or quasi-balanced flight.
Preferably, in the above technical solution, in the step three:
introducing height control will bring about a change in the tilt angle of the control acceleration, resulting in expression 29):
ac sin(ΘP+ΔΘP)＝ac sinΘP+Δac_r 29)；
from expression 29), consider Δ ΘPFor small quantities, expression 30) is derived from the equivalent change of the trigonometric function, for calculating the acceleration tilt angle additional control amount in consideration of the track height maintenance control.
The invention provides a carrier rocket orbit height keeping control method based on a balanced flight theory, which has the specific scheme that: firstly, carrying out autonomous guidance control on the rocket based on a balance theory (comprising modes such as balance flight and quasi-balance flight); secondly, acquiring an acceleration radial additional control quantity; and finally, acquiring an acceleration dip angle additional control quantity according to the acceleration radial additional control quantity, and then acquiring a thrust acceleration dip angle considering the track height maintaining control. The control method has the characteristics of high precision and small calculation amount, can meet the engineering application requirements of the rocket-borne computer on real-time calculation, and is high in practicability.
In addition to the objects, features and advantages described above, other objects, features and advantages of the present invention are also provided. The present invention will be described in further detail below with reference to the drawings.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the invention and, together with the description, serve to explain the invention and not to limit the invention. In the drawings:
FIG. 1 is a schematic view of a transient ellipse in ballistic flight;
FIG. 2 is a schematic diagram of the mechanical analysis of a rocket in flight outside the atmosphere according to the present invention;
FIG. 3 is a schematic view of a force analysis of a continuous thrust rocket according to the present invention;
FIG. 4(a) is a guidance law diagram for balanced flight;
FIG. 4(b) is a guidance law diagram for quasi-equilibrium flight;
fig. 5 is a comparison graph of simulation results of the balanced flight guidance law and the balanced flight guidance law of the present invention in consideration of the track height maintenance control.
Detailed Description
Embodiments of the invention will be described in detail below with reference to the drawings, but the invention can be implemented in many different ways, which are defined and covered by the claims.
Example (b):
an autonomous control method used in a rocket thrust failure mode, namely a carrier rocket orbit height maintaining control method based on a balanced flight theory, specifically comprises the following steps:
firstly, carrying out autonomous guidance control on a rocket:
1. the details of the analysis of the rocket flight motion under the thrust failure mode are as follows:
according to the theory of projectile ballistics of the carrier rocket, the rocket trajectory is a portion of an elliptical orbit of the geocenter, as shown in figure 1. Due to different mission missions, the trajectories of a carrier rocket and a ballistic missile are also different, the instantaneous elliptical orbit near point of the carrier rocket orbit point is generally above the safe height, and the elliptical orbit near points of the last-stage main engine shutdown point of the ballistic missile intersect with the earth.
In the air layer flight section before the orbit, the rocket motion is generally expressed as the lifting process of the flight height and the speed, but in consideration of the engineering constraints such as the actual thrust-weight ratio of the rocket and the like, the two lifting processes do not require synchronization but are emphasized to achieve better efficiency. For example, in the section before the point of entry, a strategy of continuously increasing the speed and making the altitude change more stable is often adopted (i.e., the speed increment is mainly reflected in the circumferential direction, and the radial speed is basically kept constant), and at this time, the motion of the rocket is in the stage of the far point of the instantaneous elliptical orbit and continuously increasing the near point, such as the near point lifting in fig. 1. If the rocket has insufficient thrust, the height of the trajectory near the place cannot be timely lifted, the intersection of the trajectory and the earth cannot be avoided, and according to the elliptic trajectory theory, the height of the trajectory in subsequent rocket flight continuously drops, and finally the task fails.
The rocket motion under the thrust failure mode is subjected to mechanical analysis, a local horizontal coordinate system (the original point is the rocket center of mass, and the three axes are respectively in the radial direction, the circumferential direction and the ballistic surface normal direction) is established, the rocket stress is concentrated in the ballistic plane outside the atmosphere, and the stress conditions of the rocket body in the radial direction and the circumferential direction are shown in figure 2, wherein: the radial force is the radial component of the gravity, the centrifugal force and the thrust, and the resultant force determines the motion of the rocket in the radial direction, namely the height direction, wherein the centrifugal force is related to the circumferential speed and the orbit height. The circumferential force is the circumferential component of the thrust, the effect is to change the circumferential velocity, and at the same time, the circumferential velocity change will directly affect the centrifugal force.
If the thrust value is large enough, the rapid acceleration in the circumferential direction can be realized on the basis of ensuring the radial three-force balance, and the rocket has higher maneuvering orbital transfer efficiency; if the thrust is too small, the radial three-force balance cannot be supported, and at the moment, if the speed cannot be effectively increased in the living height range, the centrifugal force is improved, and finally the rocket can be crashed.
The magnitude of the engine thrust for the radial balance portion is a major loss of capacity because it is not translated into a speed increment (understandable against impulse effects); therefore, the more the thrust is reduced when the rocket fails, the longer the flight time is, and the larger the ratio of the capacity loss is; it can be extended from this that if the radial component of the thrust is used to ensure the radial force balance, the circumferential component of the thrust is accelerated, and the thrust direction is adjusted in real time as the circumferential velocity increases and the centrifugal acceleration increases, so that the radial force is always balanced, the minimum capacity loss is realized at the acceleration level, and the optimal ballistic adjustment is also realized. The flight dynamics mechanism analysis not only can explain the basic principle that thrust fault mission loss occurs in the air layer flight section of the rocket before the rocket enters the orbit, but also provides an idea for fault disposal strategy research.
2. The dynamic modeling of the rocket balance flight process specifically comprises the following steps:
in the process that the rocket flies out of the atmospheric layer and runs towards the target orbit, a rocket stress model of continuous thrust is shown in figure 3, the rocket is under the action of the gravity of the earth and the thrust of an engine, and the centrifugal force also needs to be considered in a local horizontal coordinate systemThe inertial force acts. In FIG. 3, OEDenotes the geocentric, beta denotes the range angle, r denotes the geocentric distance, acIndicating thrust acceleration (vector), arRepresenting the radial component of thrust acceleration, aθRepresenting the circumferential component of thrust acceleration, ΘPIndicating thrust acceleration inclination (i.e. thrust acceleration a)cAngle to the local horizontal, also called best guidance law), v represents the velocity vector.
The mass change in rocket flight is attributed to the change of acceleration, and the rocket flight dynamics are obtained in the radial direction and the circumferential direction of a local horizontal coordinate system as shown in an expression 1):
Figure BDA0002744785150000061
wherein: t represents time, and μ represents an earth gravity coefficient.
3. Judging the state of the rocket and carrying out autonomous guidance control, wherein the details are as follows:
thrust acceleration a of rocketcThe determination is determined according to the thrust fault mode, can be a determined time function, and can also be measured in real time. This thrust acceleration is also determined over time, which is generally determined by default as a failure mode, and is not a constant, but rather a time-varying quantity, without measurement.
3.1, when the balance flight is satisfied, the radial total acceleration component and the velocity component in the flight process are both 0, namely
Figure BDA0002744785150000062
Figure BDA0002744785150000063
Considering the determination of the magnitude of the continuous thrust acceleration, the direction is adjustable, and the thrust acceleration is decomposed and substituted into an expression 1) to obtain an expression 2):
Figure BDA0002744785150000064
obtaining an expression 3) from an expression 2) to solve the optimal guidance law theta of the balanced flightP：
Figure BDA0002744785150000071
The flight angular velocity of the target circular orbit
Figure BDA0002744785150000072
And range angular velocity
Figure BDA0002744785150000073
Substituting expression 2), and combining the two equations (specifically, squaring the left and right sides of the two equations in expression 2), and then adding the left and right sides to obtain expression 4):
Figure BDA0002744785150000074
solving an ordinary differential equation based on expression 4) to obtain expression 5):
Figure BDA0002744785150000075
further integrating the two sides of the expression 5) to obtain an expression 6) for calculating the maneuvering orbital transfer time T from the elliptical trajectory to the circular orbit of the balanced flight:
Figure BDA0002744785150000076
by solving the quantitative integral expression 6), the maneuvering orbital transfer time from the elliptical trajectory to the circular orbit meeting the balanced flight condition can be obtained, and the flight process from the acceleration of the elliptical trajectory to the target circular orbit can be analyzed.
Such as: the ratio of the thrust acceleration to the gravitational acceleration of the circular orbit is set as
Figure BDA0002744785150000077
When ω < n, transforming denominator of integral term in expression 6) to obtain expression 7):
Figure BDA0002744785150000081
transforming expression 6) to obtain expression 8):
Figure BDA0002744785150000082
wherein:
Figure BDA0002744785150000083
the ratio of the range angular velocity omega to the flight angular velocity n of the target circular orbit; k is the ratio of the gravitational acceleration and the thrust acceleration of the circular orbit,
Figure BDA0002744785150000084
taking integral intermediate transformation variables tau and alpha, order
Figure BDA0002744785150000085
α＝τ2，τ0Is taken as the value of t at the moment t. Integrating expression 8) to obtain expression 9):
Figure BDA0002744785150000091
wherein: t is time; EllipticF is a first type of incomplete elliptic integral, or can be further expanded into expression 10):
Figure BDA0002744785150000092
the solution of the first kind of incomplete elliptic integral can refer to the prior art, can give a high-order approximate solution to meet the requirement of quick calculation, or utilizesThe numerical integration method performs the calculation. In the case of the expression 10) can be solved analytically, ω (T) (i.e. the range angular velocity at time T) is obtained actually, and T ∈ [0, T ∈ [ T ] []As a function of, and thus the thrust acceleration tilt angle thetaP(optimal guidance law) according to expression 3) can be solved quickly. Therefore, the balance flight process can be theoretically analyzed, the time and the fuel consumption of the whole process can be rapidly calculated, and the speed loss or the thrust efficiency of the continuous thrust orbital transfer process can also be calculated.
It should be noted that: the balanced flight state only represents temporary safety and does not represent long-term danger relief, if a fuel leakage condition exists, the circumferential acceleration time is not long enough, the circumferential speed constraint required by a circular orbit can not be reached when fuel is exhausted, and the rocket still has difficulty in entering the safe orbit.
In the solving process of expression 10), the condition to be satisfied is expression 11):
Figure BDA0002744785150000093
when the rocket is in an acceleration section, and omega is less than or equal to n, the condition required to be met by balanced flight can be obtained as expression 12):
Figure BDA0002744785150000094
according to the balanced flight condition, whether the rocket is in a dangerous state or not or whether the rocket has self-rescue capability or not can be judged based on the thrust acceleration level of the rocket after the fault.
3.2, because of the existence of the thrust acceleration circumferential component, the circumferential velocity component will be increased, bringing the effect that the centrifugal acceleration will be increased continuously, thereby reducing the thrust acceleration component required for achieving the radial force balance continuously, so the circumferential direction and the radial direction are dynamic processes of mutual coupling and mutual conversion, and therefore the balanced flight condition expression 12) still has a certain margin.
If it is not
Figure BDA0002744785150000101
The centrifugal acceleration can be increased according to the height level of the rocket, namely whether the rocket can be accelerated rapidly along the circumferential direction within the range of slightly reducing the allowable height, so that the further descending of the rocket is restrained, and the judgment can be carried out through time estimation.
The thrust acceleration is completely concentrated in the circumferential direction, and the radial negative acceleration, namely theta, is not consideredPWhen the thrust acceleration is 0, the effect of the thrust acceleration on the circumferential acceleration is the best, so the radial centrifugal acceleration is accelerated fastest, the centrifugal acceleration and the gravity are balanced through the transition time delta T, the rocket reaches the survival altitude, and the geocentric vector is rL，rLR- Δ h, Δ h is a height margin, and can be approximated using expression 13):
Figure BDA0002744785150000102
wherein: v. ofθIs the circumferential velocity component at the moment of failure,
Figure BDA0002744785150000103
then the condition for achieving rocket landing to rise is expression 14):
Figure BDA0002744785150000104
i.e. the sum of centrifugal acceleration and thrust acceleration is greater than gravitational acceleration, avθFor the circumferential velocity increase in this process, Δ vθ≈acΔT。
If the thrust acceleration a of the rocketcExpression 15 is not satisfied), the rocket enters the atmosphere to crash.
Figure BDA0002744785150000105
3.3 thrust acceleration a if rocketcSatisfy the requirement ofExpression 15) but not expression 12), the rocket enters a quasi-equilibrium flight state.
In fact, during the flight, the acceleration in the circumferential direction will make the speed increase fastest, the centrifugal acceleration increase fastest, but the initial radial negative acceleration value in this case is also the largest, and is also fast under the altitude; conversely, if the entire thrust is applied in the radial direction, the negative acceleration in the radial direction is minimal, the height decrease is slow, but the centrifugal acceleration cannot be increased. Between the two extremes mentioned there is a compromise between a limited range of altitude reduction and a fast implementation of the equilibrium flight regime, which requires optimization of the thrust acceleration tilt angle and minimum burn-up if the transition time from quasi-equilibrium flight to equilibrium flight is minimized.
Expressing the centrifugal acceleration by the circumferential velocity component, expression 1) is transformed into expression 16):
Figure BDA0002744785150000111
after the time of delta T, the rocket can reach new balance to make r0Representing the initial ground center distance of quasi-equilibrium flight, and the ground center distance (r, also referred to herein as equilibrium flight initial ground center distance) is not changed, then r ═ r0Expression 17 is derived from expression 16):
Figure BDA0002744785150000112
for expression 17), the second order small quantity a is omittedc 2cos2ΘP·ΔT2Obtaining a transition time Δ T expression 18 from quasi-equilibrium flight to equilibrium flight):
Figure BDA0002744785150000113
the shortest transition time Delta T from quasi-equilibrium flight to equilibrium flight is
Figure BDA0002744785150000114
Yielding expression 19):
Figure BDA0002744785150000121
in conjunction with expression 20), expression 21) can be employed to calculate the optimal guidance law ΘP：
Figure BDA0002744785150000122
Figure BDA0002744785150000123
By comparing expression 3) with expression 21), it can be found that satisfaction is satisfied
Figure BDA0002744785150000124
Balanced flight conditions of the condition, and satisfaction
Figure BDA0002744785150000125
The guidance law in the case of the quasi-equilibrium flight state of the condition is shown in fig. 4(a) and 4 (b). Under the quasi-equilibrium flight condition, the optimal direction of the thrust acceleration is not along the radial direction or the circumferential direction, but the different actions of the two directions are still considered, so that the shortest flight time of the radial force equilibrium can be realized. Since the balanced flight mode represents the most effective utilization mode of the thrust, the expressions 3) and 21) are also the optimal guidance schemes for the extraatmospheric flight section before the rocket enters the orbit.
Secondly, acquiring an acceleration radial additional control quantity delta ac_r：
In order to meet the requirement of precision of the rocket orbit entering point, the height control of the orbit entering point is required, and the realization of the height control can be reflected to the thrust acceleration inclination angle by designing a guidance law for controlling the acceleration.
The radial equation for orbital motion can be written as expression 22):
Figure BDA0002744785150000131
altitude control can be viewed as an additional control based on radial balance flight control as in expression 23):
Figure BDA00027447851500001311
wherein: Δ r is the deviation of the actual ground center distance and the target ground center distance; Δ ac_rAdding a control quantity to the acceleration in a radial direction;
designing the control law of the height feedback as expression 24):
Figure BDA0002744785150000132
wherein: k is a radical of1And k2Control parameters of a height feedback control law; here, Δ r is Δ H, and Δ H is a deviation of the actual height from the track-keeping target height.
Substituting expression 24) into expression 23), expression 25) can be obtained:
Figure BDA0002744785150000133
wherein:
Figure BDA0002744785150000134
is the rate of change of Δ r;
Figure BDA0002744785150000135
is composed of
Figure BDA0002744785150000136
The rate of change of (c);
expression 25) is naturally converged due to ω → n approaching 0 during the rocket flight approaching equilibrium, and therefore, control law design can be performed based on the left side of expression 25).
Obtaining the characteristic equation of the above expression 25) as expression 26):
λ2+k2λ+k1＝0 26)；
in order to make the system converge as soon as possible, the parameter k should be designed2And k1。
Root of the characteristic equation of expression 25) is expression 27):
Figure BDA0002744785150000137
for the control system to be stable, the conditional expression 28) should be satisfied:
k2＞0，k1is greater than 0, and
Figure BDA0002744785150000138
alternatively, if the natural frequency and damping ratio of the second-order system are expressed, there are
Figure BDA0002744785150000139
Figure BDA00027447851500001310
Is a natural frequency; ξ is the damping ratio.
Therefore, the natural frequency and the damping ratio are designed, and control law design can be achieved.
Thirdly, acquiring an acceleration dip angle additional control quantity and acquiring a thrust acceleration dip angle considering track height maintaining control:
introducing height control will bring about a change in the tilt angle of the control acceleration, resulting in expression 29):
acsin(ΘP+ΔΘP)＝acsinΘP+Δac_r 29)；
from expression 29), consider Δ ΘPFor small quantities, expression 30 is derived from the equivalent change of the trigonometric function), the acceleration tilt angle additional control quantity considering the track height maintenance control is calculated:
Figure BDA0002744785150000141
a thrust acceleration inclination angle theta in consideration of the track height maintenance control is obtainedP+ΔΘP。
Simulation analysis is carried out on a power flight section outside the atmospheric layer of a certain type of two-stage carrier rocket, the failure mode is the thrust descent failure of the engine without fuel leakage, the failure time is set to be 350s, and the residual thrust proportion of the engine is 55%. The result obtained by the balance flight guidance law simulation is shown as a solid line in fig. 5, and the flying height of the circular orbit reaches 245.1 km; design a set of k1And k2(the natural frequency is preferably 50 here, and the damping ratio is 0.9), and the track holding altitude is set to 200km, the result of the simulation of the balanced flight guidance law with track holding (the solution of the present invention) is shown by the dotted line in fig. 5, and the circular orbit flight altitude can be controlled to 205.8 km. Therefore, the track keeping control method is effective, the target of controlling the height of the track is achieved, and high precision is achieved.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims (3)

1. A carrier rocket orbit height keeping control method based on a balanced flight theory is characterized by comprising the following steps:
the method comprises the following steps of firstly, carrying out autonomous guidance control on the rocket, specifically:
in the balanced flight process, the optimal guidance law theta is obtained through the expression 3)PThe maneuvering orbital transfer time T from the elliptical trajectory to the circular orbit of the equilibrium flight is obtained by expression 6):
Figure FDA0002744785140000011
Figure FDA0002744785140000012
wherein: r represents the geocentric distance; μ represents an earth gravity coefficient; n is the flight angular velocity of the target circular orbit,
Figure FDA0002744785140000013
omega is the angular velocity of the range,
Figure FDA0002744785140000014
t represents time; a iscThrust acceleration of the rocket;
in the quasi-equilibrium flight process, the transition time delta T from quasi-equilibrium flight to equilibrium flight is obtained through an expression 18), and the optimal guidance law theta is obtained through an expression 21)P：
Figure FDA0002744785140000015
Figure FDA0002744785140000016
Wherein: r is0Representing an initial ground center distance of quasi-equilibrium flight; v. ofθIs the circumferential velocity component at the moment of failure,
Figure FDA0002744785140000017
step two, acquiring an acceleration radial additional control quantity delta a through an expression 24)c_r：
Figure FDA0002744785140000018
Wherein: k is a radical of1And k2Is a heightThe control parameters of the feedback control law are taken
Figure FDA0002744785140000021
Figure FDA0002744785140000022
Is a natural frequency; xi is a damping ratio; Δ r is the deviation of the actual ground center distance and the target ground center distance;
Figure FDA0002744785140000023
is the rate of change of the deviation of the actual ground center distance from the target ground center distance;
step three, acquiring the additional control quantity delta theta of the acceleration inclination angle through an expression 30)PAnd then the thrust acceleration inclination angle theta considering the track height holding control is obtainedP+ΔΘP：
Figure FDA0002744785140000024
2. The balance flight theory-based carrier rocket orbit altitude maintaining control method according to claim 1, wherein the autonomous guidance of the rocket in the first step is specifically as follows:
step a1, carrying out dynamic analysis on rocket motion, identifying the power system fault of the rocket, and acquiring thrust acceleration a under the fault modecA function of time of;
step a2, thrust acceleration a to rocketcThe judgment is specifically as follows:
if the thrust acceleration a of the rocketcExpression 12 is satisfied), the rocket enters a balanced flight state:
Figure FDA0002744785140000025
wherein: g0Is the acceleration of the gravity of the circular orbit,
Figure FDA0002744785140000026
if the thrust acceleration a of the rocketcExpression 15) is satisfied but expression 12 is not satisfied), the rocket enters a quasi-balanced flight state;
Figure FDA0002744785140000027
wherein: Δ h is the height margin;
if the thrust acceleration a of the rocketcExpression 15 is not satisfied), the rocket enters the atmosphere to crash;
step a3, carrying out autonomous guidance according to the condition that the rocket is in balanced flight or quasi-balanced flight.
3. The balance-flight-theory-based launch vehicle orbit altitude maintaining control method according to claim 2, characterized in that in the third step:
introducing height control will bring about a change in the tilt angle of the control acceleration, resulting in expression 29):
acsin(ΘP+ΔΘP)＝acsinΘP+Δac_r 29)；
from expression 29), consider Δ ΘPFor small quantities, expression 30) is derived from the equivalent change of the trigonometric function, for calculating the acceleration tilt angle additional control amount in consideration of the track height maintenance control.
CN202011162810.9A 2020-10-27 2020-10-27 Carrier rocket orbit height keeping control method based on balanced flight theory Pending CN112325711A (en)

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CN113836645A (en) * 2020-10-27 2021-12-24 深圳三零三防务科技有限公司 Method for reconstructing online flight program of carrier rocket and keeping and controlling track height

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113836645A (en) * 2020-10-27 2021-12-24 深圳三零三防务科技有限公司 Method for reconstructing online flight program of carrier rocket and keeping and controlling track height
CN113836645B (en) * 2020-10-27 2024-02-02 深圳三零三防务科技有限公司 On-line flight program reconstruction and orbit height maintenance control method for carrier rocket

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