CN109407551B - Method for carrying out semi-physical simulation test on joint control section of carrier rocket - Google Patents
Method for carrying out semi-physical simulation test on joint control section of carrier rocket Download PDFInfo
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Abstract
The invention discloses a method for carrying out a semi-physical simulation test on a joint control section of a carrier rocket, which comprises the following steps: time synchronization processing is carried out on the linear section control system and the nonlinear section control system, so that when the linear section control system controls a servo system of the rocket body model, the nonlinear section control system also controls an attitude control spray pipe of the rocket body model; carrying out simulation calculation on the rocket body model, wherein the simulation calculation comprises the following steps: the method comprises the steps that interference torque generated by an attitude control spray pipe is superposed into a linear control model and is resolved to obtain a first attitude angle error, and interference torque generated by a servo system is superposed into a nonlinear control model and is settled to obtain a second attitude angle error; and combining and superposing the first attitude angle error and the second attitude angle error according to the proportional coefficients of the control moments respectively generated by the servo system and the attitude control spray pipe to obtain the attitude angle error at the current moment. The invention realizes the purpose of carrying out semi-physical simulation test on the carrier rocket in the joint control section.
Description
Technical Field
The invention relates to the technical field of semi-physical simulation, in particular to a method for carrying out semi-physical simulation test on a joint control section of a carrier rocket.
Background
The carrier rocket is generally a three-stage rocket, and the flight process of the carrier rocket can be divided into the following flight sections according to the change of characteristics of the carrier rocket in the flight process: a first-stage, a second-stage and a third-stage active section and a gliding section; in order to meet the requirements of different flight sections, different attitude control systems are correspondingly adopted to carry out attitude control on the carrier rocket.
Specifically, the attitude control system of the launch vehicle can be divided into attitude control systems of a primary, a secondary and a tertiary active section (linear section) and a gliding section (nonlinear section) according to the change of characteristics and the difference of actuating mechanisms in the flight process of the launch vehicle. The attitude control systems of the first-stage, second-stage and third-stage active sections effectively and correspondingly eliminate the interference influence generated by the carrier rocket in each flight section by adopting a control method of a swing engine, and quickly and correctly realize the control of the attitude of the carrier rocket. For the attitude control system of the nonlinear section, a method for controlling the attitude by adopting an attitude control nozzle (a typical nonlinear switch control method) is adopted to eliminate the nonlinear interference generated by the carrier rocket in the flight section of the gliding section.
As shown in fig. 1, the control method of the swing engine includes the following processes: the attitude motion state of the carrier rocket is measured in real time through an attitude angle or attitude angular velocity measuring device in a linear section attitude control system to obtain attitude angle and attitude angular velocity signals, the measured attitude angle and attitude angular velocity signals are subjected to signal synthesis in a signal synthesis and synthesis device in the linear section attitude control system, and are subjected to signal synthesis through a correction network or an attitude controller in the linear section attitude control system to generate attitude control signals, and the attitude control signals are directly transmitted to a control device (servo system) in the linear section attitude control system to drive an engine of the servo system to swing to generate control torque and perform attitude control on the carrier rocket.
As shown in fig. 2, the non-linear control of the launch vehicle generally uses a non-linear switch control method, and the nozzle of the launch vehicle is controlled to work in a positive-on-off-negative-on manner. Taking the pitching channel as an example, the attitude dynamics equation of the attitude control nozzle control section (nonlinear control section) is as follows:
in the formula: i is a symbol for controlling the switch,
in order to correct the angular deviation of the pitch attitude,
for pitch attitude angular rate, alpha0As static amplification factor, alpha, of the attitude angle channel1And the dynamic amplification factor of the attitude angular rate channel.
Formula 1 describes the dynamic characteristics of the posture of the rocket body during the control of the posture control spray pipe, and a satellite-rocket separation section, a discharge section and a gliding section, wherein the shaking of the propellant is considered during the axial overload period, and only the rigid body motion is considered during the non-axial overload period. In conclusion, when the carrier rocket is in a nonlinear flight section, the attitude control nozzle can be applied to control the attitude of the carrier rocket.
Research shows that after the carrier rocket executes the flight process of the three-stage active section, the three-stage engine of the carrier rocket is shut down, and in a time period of 5s after the shut-down of the carrier rocket is finished, namely the attitude motion of the carrier rocket is subjected to combined attitude control by a servo system of a control mechanism of a linear section attitude control system and an attitude control spray pipe controlled by a nonlinear section attitude control system, and the control stage is called as a combined control section.
As shown in FIG. 3, a timing diagram of the combined control section is provided, wherein tk31 represents three-stage engine shutdown time, T- "0" -III represents three-stage active section control system ending control time, and T-ZK-T1 represents three-stage gliding section control path closing time.
And after the flight section of the three-stage active section is finished, the three-stage engine is shut down, and after the three-stage engine is shut down for 5s, the control of the three-stage active section control system is finished, namely the servo mechanism returns to zero. And (3) the three-stage engine is shut down, and after the engine is shut down for 1s, the three-stage gliding section control path is communicated, so that the carrier rocket is subjected to attitude control by combining a servo system of a linear section control system control mechanism and an attitude control spray pipe of a nonlinear section control system control mechanism in 4s from T-ZK-T1 to T- '0' -III, and the control effect of the combined control section and the caused attitude disturbance cannot be subjected to simulation test or examination by adopting the conventional sectional semi-physical simulation test method for the flight section of the carrier rocket combined control section.
Disclosure of Invention
The invention aims to provide a method for performing a semi-physical simulation test on a joint control section of a carrier rocket, which achieves the purpose of performing the semi-physical simulation test on the joint control section under the simultaneous action of two execution mechanisms by adopting a mode of linearly superposing the effect of attitude angle control on the basis of the configuration of a semi-physical simulation test system of the conventional carrier rocket control system.
In order to achieve the above purpose, the invention is realized by the following technical scheme:
a method for carrying out a semi-physical simulation test on a joint control section of a carrier rocket comprises the following processes: performing time synchronization processing on a linear section control system and a nonlinear section control system, so that when the linear section control system controls a servo system of an arrow body model, the nonlinear section control system also controls an attitude control spray pipe of the arrow body model; performing simulation calculation on the rocket body model, wherein the simulation calculation comprises the following steps: the interference torque generated by the attitude control spray pipe is superposed into the linear control model and is resolved to obtain a first attitude angle error, and the interference torque generated by the servo system is superposed into the nonlinear control model and is settled to obtain a second attitude angle error; and combining and superposing the first attitude angle error and the second attitude angle error according to the proportional coefficients of the control moments respectively generated by the servo system and the attitude control spray pipe to obtain the attitude angle error at the current moment.
Further, the linear control model is a first linear control kinetic equation:
in the formula, ωX1、ωy1、ωz1The attitude angular velocity of the rocket model calculated for the solution of the linear kinetic equation,
ψ1、γ1attitude angle deviations in pitch, yaw and rolling directions which are respectively calculated by solving a linear dynamic equation,
δψ、δγthe engine pivot angle, J, in the servo system controlling the pitch, yaw and roll directions, respectivelycIs the moment of inertia of the rocket model, d30、b3、b30To control moment coefficient, M'rX、M′rY、M′rZThe disturbance moments in pitching, yawing and rolling directions are respectively;
the interference torque generated by the attitude control spray pipe is as follows:
in the formula, M1、M2、M3Interference moments generated by the attitude control spray pipe in the pitching direction, the yawing direction and the rolling direction are respectively;
Kψ、Kγrespectively control the pitching direction, the yawing direction and the rolling direction,
b3ψ、d3γcontrol moment coefficients in pitching, yawing and rolling directions are respectively;
the second linear control dynamic equation obtained after the disturbance moment generated by the attitude control spray pipe is superposed is as follows:
and calculating by using the second linear control braking mechanical equation to obtain the first attitude angle error.
Further, the nonlinear control model is a first nonlinear control kinetic equation
In the formula, ωX2、ωy2、ωz2The attitude angular velocity of the rocket model calculated for the solution of the nonlinear dynamical equation,
ψ2、γ2attitude angle deviations in pitch, yaw and rolling directions which are respectively calculated by solving a nonlinear dynamical equation,
Kψ、Kγthe attitude control spray pipe switching signals for controlling pitching, yawing and rolling directions are respectively determined by the working conditions of the attitude control engine, namely +1 is taken when the positive attitude control spray pipe works, -1 is taken when the negative attitude control spray pipe works, and 0, M 'is taken when the attitude control spray pipe does not work'rX1、M′rY1、M′rZ1Respectively are disturbance moments in pitching, yawing and rolling directions,
b3ψ、d3γcontrol moment coefficients in pitching, yawing and rolling directions are respectively; the disturbance torque generated by the servo system is as follows:
and (3) superposing the disturbance torque generated by the servo system to obtain a second nonlinear control dynamic equation:
and calculating by using the second nonlinear control dynamics equation to obtain the second attitude angle error.
Further, the step of linearly superimposing the first and second attitude angle errors further comprises:
at the same simulation moment, the pitching direction control moment generated by the swing angle of the engine is as
The pitching direction control moment generated by the attitude control spray pipe is as follows
The two control modes control the proportional coefficient of the moment to be
The sum of the pitch attitude angle deviations generated by the two kinetic equations is
Namely the pitching attitude angle deviation value of the arrow model entering the measurement equation;
at the same simulation moment, the yaw direction control moment generated by the swing angle of the engine is equal to
The yaw direction control moment generated by the attitude control spray pipe is Kψb3ψThe two control modes control the proportional coefficient of the torque to be
The sum of the yaw attitude angle deviations generated by the two kinetic equations is
Psi is the yaw attitude angle deviation value of the rocket model entering the measurement equation;
at the same simulation moment, the rolling direction control moment generated by the swing angle of the engine is as follows
The rolling direction control moment generated by the attitude control spray pipe is d3γKγThe two control modes control the proportional coefficient of the torque to be
The rolling attitude angle deviation superposition value generated by the two kinetic equations is
And gamma is the rolling attitude angle deviation value of the arrow body model entering the measurement equation.
The invention has the following technical effects:
according to the method, time synchronization processing is carried out on the linear section control system and the nonlinear section control system, so that when the linear section control system controls the servo system of the rocket body model, the nonlinear section control system also controls the attitude control spray pipe of the rocket body model; carrying out simulation calculation on the rocket body model, wherein the simulation calculation comprises the following steps: the method comprises the steps that interference torque generated by an attitude control spray pipe is superposed into a linear control model and is resolved to obtain a first attitude angle error, and interference torque generated by a servo system is superposed into a nonlinear control model and is settled to obtain a second attitude angle error; and combining and superposing the first attitude angle error and the second attitude angle error according to the proportional coefficients of the control moments respectively generated by the servo system and the attitude control spray pipe to obtain the attitude angle error at the current moment, thereby realizing the purpose of performing a semi-physical simulation test on the carrier rocket in the joint control section.
Drawings
FIG. 1 is a block diagram of the main components of a linear section attitude control system of a launch vehicle in the prior art;
FIG. 2 is a block diagram of the main components of a prior art attitude control system for a nonlinear section of a launch vehicle;
FIG. 3 is a timing diagram of a joint control segment in the prior art;
FIG. 4 is a schematic diagram illustrating a schematic principle of a method for performing a semi-physical simulation test on a joint control section of a launch vehicle according to an embodiment of the present invention;
fig. 5 is a schematic flow chart of a method for performing a semi-physical simulation test on a joint control section of a launch vehicle provided in an embodiment of the present invention.
Detailed Description
Research shows that in a combined control section for jointly carrying out attitude control on a carrier rocket in a 5s time period after the shutdown of an engine in a flight active section of the carrier rocket, namely a servo system controlled by a linear section control system and an attitude control spray pipe controlled by a nonlinear section control system, a sectional semi-physical simulation experiment system cannot carry out semi-physical simulation experiment on the carrier rocket, namely the control effect of the combined control section and the caused attitude disturbance cannot be subjected to simulation experiment or examination.
Aiming at the problems, the invention provides a method for carrying out a semi-physical simulation test on a joint control section of a carrier rocket, which realizes the semi-physical simulation of the joint control section by adopting a mode of carrying out linear superposition on the effect of attitude angle control.
The present invention will now be further described by way of the following detailed description of a preferred embodiment thereof, taken in conjunction with the accompanying drawings.
Referring to fig. 1 and fig. 2, a method for performing a semi-physical simulation test on a joint control segment of a launch vehicle according to this embodiment includes the following steps: and (3) a closed-loop semi-physical simulation system is built, and when the carrier rocket is subjected to attitude control jointly by a servo system controlled by a linear section control system and an attitude control spray pipe controlled by a nonlinear section control system, an rocket body model of the carrier rocket is built and is subjected to parallel independent calculation.
Specifically, first, the linear section control system and the nonlinear section control system are subjected to time synchronization processing, so that when the linear section control system controls the servo system, the nonlinear section control system also controls the attitude control nozzle.
And (3) superposing interference items generated by the attitude control spray pipe into the linear control model, and superposing interference generated by the servo mechanism into the nonlinear control model.
Further, the disturbance torque generated by the servo system is superposed to the first nonlinear control dynamics equation to be calculated to obtain a second nonlinear control dynamics equation, and the second nonlinear control dynamics equation is used for calculation to obtain the first attitude angle error.
And superposing the interference torque generated by the attitude control spray pipe to the first linear control dynamic equation to calculate to obtain a second linear control dynamic equation, and calculating by using the second linear control dynamic equation to obtain a second attitude angle error.
Further, the first linear control dynamics equation is:
in the formula (I), the compound is shown in the specification,
δψ、δγthe engine yaw angle, omega, in a servo system for controlling the pitch, yaw and roll directions, respectivelyX1、ωy1、ωz1Angular velocity of attitude of rocket model calculated for solution of linear kinetic equation, JcIs the moment of inertia of the rocket model, d30、b3、b30In order to control the torque coefficient, the torque coefficient is controlled,
ψ1、γ1attitude angle deviations, M ', in pitch, yaw, and roll directions, respectively, solved by a linear kinetic equation'rX、M′rY、M′rZThe disturbance moments in pitching, yawing and rolling directions are respectively.
The attitude control spray pipe corresponds to linear control dynamics, and the generated disturbance moment is as follows:
in the formula, M1、M2、M3The interference moment generated by the attitude control spray pipe,
Kψ、Kγthe attitude control spray pipe switching signals for controlling pitching, yawing and rolling directions are respectively determined by the working conditions of the attitude control engine, namely, the positive attitude control spray pipe is taken as +1 when working, the negative attitude control spray pipe is taken as-1 when working, and the attitude control spray pipe is taken as 0 when not working,
b3ψ、d3γthe control moment coefficients of pitching, yawing and rolling directions are respectively.
The second linear control dynamic equation obtained after the disturbance moment generated by the attitude control spray pipe is superposed is as follows:
calculating by using the formula (4) to obtain a first attitude angle error as follows:
ψ1、γ1。
the first nonlinear control dynamics equation is:
in the formula, ωX2、ωy2、ωz2The attitude angular velocity of the rocket model calculated for the solution of the nonlinear dynamical equation,
ψ2、γ2attitude angle deviations in pitch, yaw and rolling directions which are respectively calculated by solving a nonlinear dynamical equation,
Kψ、Kγthe attitude control spray pipe switching signals for controlling pitching, yawing and rolling directions are respectively determined by the working conditions of the attitude control engine, namely +1 is taken when the positive attitude control spray pipe works, -1 is taken when the negative attitude control spray pipe works and 0 is taken when the attitude control spray pipe does not work. Attitude control engine layouts may be different for different rockets, but the operating conditions are similar. M'rX1、M′rY1、M′rZ1The disturbance moments in pitching, yawing and rolling directions are respectively. The disturbance torque generated by the servo system (specifically the engine swing angle in the servo system) on the nonlinear control dynamics is as follows:
and superposing the disturbance torque generated by the swing angle of the engine to obtain a second nonlinear control dynamic equation as follows:
and calculating by using the formula (7) to obtain the second attitude angle error as follows:
ψ2、γ2。
and the attitude angle error obtained by superposing and combining the first attitude angle error and the second attitude angle error according to the proportion of control moments respectively generated by the attitude control spray pipe and the servo system is used as the input value of the measurement system at the current simulation moment. And then the measuring system outputs the new attitude angle error data, and the simulation is finished.
And linearly superposing the first attitude angle error and the second attitude angle error according to the control moment generated by the two control modes, wherein the specific superposition mode is as follows:
at the same simulation moment (same simulation period), the control moment of the pitching direction generated by the swing angle of the engine is equal to
The pitching direction control moment generated by the attitude control spray pipe is as follows
The two control modes control the proportional coefficient of the moment to be
The sum of the pitch attitude angle deviations generated by the two kinetic equations is
Namely the pitch attitude angle deviation value of the arrow model entering the measurement equation.
At the same simulation moment, the yaw direction control moment generated by the swing angle of the engine is equal to
The yaw direction control moment generated by the attitude control spray pipe is Kψb3ψThe two control modes control the proportional coefficient of the torque to be
The yaw attitude angle deviation generated by the two kinetic equationsA superimposed value of
Psi is the yaw attitude angle deviation value of the rocket model entering the measurement equation.
At the same simulation moment, the rolling direction control moment generated by the swing angle of the engine is as follows
The rolling direction control moment generated by the attitude control spray pipe is d3γKγThe two control modes control the proportional coefficient of the torque to be
The rolling attitude angle deviation superposition value generated by the two kinetic equations is
And gamma is the rolling attitude angle deviation value of the arrow body model entering the measurement equation.
After the first attitude angle error and the second attitude angle error are linearly superposed, the attitude angle error enters an arrow body attitude measurement system, and is output as a measurement value of the attitude angle error:
in the formula (I), the compound is shown in the specification,
and G(s) is a transfer function of the measuring link.
After the measured value of the attitude angle error enters an arrow body attitude control system, linear control controlled by a servo system and nonlinear control using an attitude control spray pipe as an actuating mechanism are respectively carried out, and the linear control output of the servo system is as follows:
in the formula, a0、a0γStatic amplification factors, G, for pitch (yaw) and roll channels, respectivelySF(s) is the transfer function of the servo system.
In summary, when the carrier rocket is subjected to attitude control by combining two actuating mechanisms, namely a servo system controlled by a linear section control system and an attitude control spray pipe controlled by a nonlinear section control system, interference terms generated by the two actuating mechanisms are simultaneously introduced into an rocket body resolving model, and the rocket body model of the carrier rocket is resolved in parallel and independently to obtain two attitude angle errors respectively. And (3) superposing and combining the two attitude angle errors according to the proportion of control moments respectively generated by the attitude control spray pipe and the servo system to obtain a new attitude angle error, taking the new attitude angle error as an input value of the measurement system at the current simulation moment, then outputting and processing the data of the new attitude angle error by the measurement system, ending the simulation of the simulation period, and entering the next simulation period cycle. Therefore, the problem that the original semi-physical simulation scheme cannot realize the semi-physical simulation of the combined control section is solved.
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.
Claims (2)
1. A method for carrying out a semi-physical simulation test on a joint control section of a carrier rocket is characterized by comprising the following processes: performing time synchronization processing on a linear section control system and a nonlinear section control system, so that when the linear section control system controls a servo system of an arrow body model, the nonlinear section control system also controls an attitude control spray pipe of the arrow body model;
performing simulation calculation on the rocket body model, wherein the simulation calculation comprises the following steps:
the interference torque generated by the attitude control spray pipe is superposed into the linear control model and is solved to obtain a first attitude angle error, and the interference torque generated by the servo system is superposed into the nonlinear control model and is solved to obtain a second attitude angle error;
combining and superposing the first attitude angle error and the second attitude angle error according to proportional coefficients of control moments respectively generated by the servo system and the attitude control spray pipe to obtain an attitude angle error at the current moment;
the linear control model is a first linear control kinetic equation:
in the formula, ωX1、ωy1、ωz1The attitude angular velocity of the rocket model calculated for the solution of the linear kinetic equation,
ψ1、γ1attitude angle deviations in pitch, yaw and rolling directions which are respectively calculated by solving a linear dynamic equation,
δψ、δγthe engine pivot angle, J, in the servo system controlling the pitch, yaw and roll directions, respectivelycIs the moment of inertia of the rocket model, d30、b3、b30To control moment coefficient, M'rX、M′rY、M′rZThe disturbance moments in pitching, yawing and rolling directions are respectively;
the interference torque generated by the attitude control spray pipe is as follows:
in the formula, M1、M2、M3Interference moments generated by the attitude control spray pipe in the pitching direction, the yawing direction and the rolling direction are respectively;
Kψ、Kγrespectively control the pitching direction, the yawing direction and the rolling direction,
b3ψ、d3γcontrol moment coefficients in pitching, yawing and rolling directions are respectively;
the second linear control dynamic equation obtained after the disturbance moment generated by the attitude control spray pipe is superposed is as follows:
calculating by using the second linear control braking mechanical equation to obtain the first attitude angle error;
the nonlinear control model is a first nonlinear control kinetic equation:
in the formula, ωX2、ωy2、ωz2The attitude angular velocity of the rocket model calculated for the solution of the nonlinear dynamical equation,
ψ2、γ2attitude angle deviations in pitch, yaw and rolling directions which are respectively calculated by solving a nonlinear dynamical equation,
Kψ、Kγrespectively controlling pitching, yawing and rolling directions by attitude controlDetermining the working condition of the engine, namely taking +1 when the positive attitude control spray pipe works, taking-1 when the negative attitude control spray pipe works, and taking 0, M 'when the attitude control spray pipe does not work'rX1、M′rY1、M′rZ1Respectively are disturbance moments in pitching, yawing and rolling directions,
b3ψ、d3γcontrol moment coefficients in pitching, yawing and rolling directions are respectively;
the disturbance torque generated by the servo system is as follows:
and (3) superposing the disturbance torque generated by the servo system to obtain a second nonlinear control dynamic equation:
and calculating by using the second nonlinear control dynamics equation to obtain the second attitude angle error.
2. The method for semi-physical simulation testing of a launch vehicle joint control section of claim 1, wherein the step of linearly superimposing the first and second attitude angle errors further comprises:
at the same simulation moment, the pitching direction control moment generated by the swing angle of the engine is as
The pitching direction control moment generated by the attitude control spray pipe is as follows
The two control modes control the proportional coefficient of the moment to be
The sum of the pitch attitude angle deviations generated by the two kinetic equations is
Namely the pitching attitude angle deviation value of the arrow model entering the measurement equation;
at the same simulation moment, the yaw direction control moment generated by the swing angle of the engine is equal to
The yaw direction control moment generated by the attitude control spray pipe is Kψb3ψThe two control modes control the proportional coefficient of the torque to be
The sum of the yaw attitude angle deviations generated by the two kinetic equations is
Psi is the yaw attitude angle deviation value of the rocket model entering the measurement equation;
at the same simulation moment, the rolling direction control moment generated by the swing angle of the engine is as follows
The rolling direction control moment generated by the attitude control spray pipe is d3γKγThe two control modes control the proportional coefficient of the torque to be
The rolling attitude angle deviation superposition value generated by the two kinetic equations is
And gamma is the rolling attitude angle deviation value of the arrow body model entering the measurement equation.
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Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH06129798A (en) * | 1992-10-19 | 1994-05-13 | Mitsubishi Heavy Ind Ltd | Objective missile |
| CN102589350A (en) * | 2012-01-09 | 2012-07-18 | 林德福 | Semi-physical simulation system for developing laser terminal guidance ammunition round |
| JP2015137991A (en) * | 2014-01-24 | 2015-07-30 | セイコーエプソン株式会社 | Functional elements, sensor device, electronic apparatus and movable body |
| CN104898635A (en) * | 2014-10-27 | 2015-09-09 | 中国运载火箭技术研究院 | High thrust liquid rocket fault reconfiguration control method |
| CN104898680A (en) * | 2015-05-04 | 2015-09-09 | 湖北航天技术研究院总体设计所 | Solid carrier rocket attitude control method based on solid variable-jet-direction jet engine |
| CN106200668A (en) * | 2016-09-12 | 2016-12-07 | 上海航天控制技术研究所 | Outer loop energy resource system and test method thereof for semi-physical simulation |
| CN106444430A (en) * | 2016-11-09 | 2017-02-22 | 上海宇航系统工程研究所 | Control system and method for sublevel reentry of carrier rocket, and simulation system and method |
| CN107807626A (en) * | 2017-09-27 | 2018-03-16 | 上海航天控制技术研究所 | One kind can autonomous configuration flight control system based on Embedded Multi-task |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20050137724A1 (en) * | 2003-10-10 | 2005-06-23 | Georgia Tech Research Corporation | Adaptive observer and related method |
-
2018
- 2018-12-14 CN CN201811529697.6A patent/CN109407551B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH06129798A (en) * | 1992-10-19 | 1994-05-13 | Mitsubishi Heavy Ind Ltd | Objective missile |
| CN102589350A (en) * | 2012-01-09 | 2012-07-18 | 林德福 | Semi-physical simulation system for developing laser terminal guidance ammunition round |
| JP2015137991A (en) * | 2014-01-24 | 2015-07-30 | セイコーエプソン株式会社 | Functional elements, sensor device, electronic apparatus and movable body |
| CN104898635A (en) * | 2014-10-27 | 2015-09-09 | 中国运载火箭技术研究院 | High thrust liquid rocket fault reconfiguration control method |
| CN104898680A (en) * | 2015-05-04 | 2015-09-09 | 湖北航天技术研究院总体设计所 | Solid carrier rocket attitude control method based on solid variable-jet-direction jet engine |
| CN106200668A (en) * | 2016-09-12 | 2016-12-07 | 上海航天控制技术研究所 | Outer loop energy resource system and test method thereof for semi-physical simulation |
| CN106444430A (en) * | 2016-11-09 | 2017-02-22 | 上海宇航系统工程研究所 | Control system and method for sublevel reentry of carrier rocket, and simulation system and method |
| CN107807626A (en) * | 2017-09-27 | 2018-03-16 | 上海航天控制技术研究所 | One kind can autonomous configuration flight control system based on Embedded Multi-task |
Non-Patent Citations (4)
| Title |
|---|
| "Computed Torque Control of an aerospace craft using nonlinear inverse model and rotation matrix";Jong Tai Jang等;《2015 15th International Conference on Control, Automation and Systems (ICCAS)》;20151231;全文 * |
| "Development of hardware-in-the-loop simultion for rocket guidance system";Yulyan Wahyu Hadi等;《2015 International Conference on Electrical Engineering and Informatics (ICEEI)》;20151231;全文 * |
| "基于最优制导的运载火箭姿态控制方法研究";周静等;《上海航天》;20161231;第33卷;全文 * |
| "运载火箭电动伺服与发动机间隙补偿控制方法";贺从园等;《上海航天》;20161231;第33卷;全文 * |
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