CN111221268A - Semi-physical simulation test method under condition of large-range change of attitude angle - Google Patents

Semi-physical simulation test method under condition of large-range change of attitude angle Download PDF

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Publication number
CN111221268A
CN111221268A CN202010076529.7A CN202010076529A CN111221268A CN 111221268 A CN111221268 A CN 111221268A CN 202010076529 A CN202010076529 A CN 202010076529A CN 111221268 A CN111221268 A CN 111221268A
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China
Prior art keywords
degrees
attitude
angle
guided rocket
attitude angle
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CN202010076529.7A
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CN111221268B (en
Inventor
肖堃
王娇
陈炜
高晓波
王鹏
张欣
刘梦焱
刘琦
张翔
朱克炜
方莉
李毅
周群凯
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Xian Institute of Modern Control Technology
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Xian Institute of Modern Control Technology
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B17/00—Systems involving the use of models or simulators of said systems
    • G05B17/02—Systems involving the use of models or simulators of said systems electric
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Engineering & Computer Science (AREA)
  • Automation & Control Theory (AREA)
  • Management, Administration, Business Operations System, And Electronic Commerce (AREA)
  • Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)

Abstract

The invention discloses a semi-physical simulation test method under the condition of wide-range change of an attitude angle, which is characterized in that when the attitude angle of a guided rocket is changed in a wide range and exceeds the working angle range of a three-axis turntable, the switching of a mark position is carried out, and the switching is carried out from three-component injection to six-component injection.

Description

Semi-physical simulation test method under condition of large-range change of attitude angle
Technical Field
The invention belongs to the technical field of system simulation, and particularly relates to a semi-physical simulation test method under the condition of large-range change of an attitude angle.
Background
Semi-physical simulation, also called hardware-in-loop simulation, is a simulation in which a part of components of a simulated system and the part of components and simulation equipment interface equipment are embedded in a simulation system, and the simulation is usually real-time simulation in which the time advance of the simulation system is strictly the same as that of a real system.
When a semi-physical simulation test is carried out, the inertial navigation device is fixed on a three-axis rotary table, the three-axis rotary table can move in yaw, rolling and pitching directions, and the simulator sends an instruction to the rotary table through an implementation network system so that the rotary table can move relatively according to a missile control model.
The traditional simulation test method is that three attitude angles of pitching, yawing and rolling of a guided missile are simulated in the trajectory flight simulation of the simulation test through a three-axis turntable, and an inertial integrated navigation device receives velocity increments of X, Y, Z injected into a computer in three directions so as to calculate the acceleration and position information of a guided rocket.
However, in the traditional simulation test, when the attitude angle of the simulated missile changes within a large range of (-75 degrees to-90 degrees or 75 degrees to 90 degrees), the use boundary of a three-axis turntable is exceeded, the attitude angle is limited, an ambiguous point is generated by the attitude angle, and the trajectory simulation cannot be completed.
Disclosure of Invention
In view of the above, the invention provides a semi-physical simulation test method under the condition of wide-range change of the attitude angle, which can complete a simulation test when the attitude angle of the guided rocket changes in a wide range.
The technical scheme for realizing the invention is as follows:
a semi-physical simulation test method under the condition of wide-range change of attitude angles is characterized in that when a pitch angle of a guided rocket is more than or equal to-75 degrees and less than or equal to 75 degrees, the attitude motion of a triaxial rotary table is utilized to simulate a pitching attitude angle, a yawing attitude angle, a rolling attitude angle and corresponding attitude angular velocities of the guided rocket; injecting speed increments of the guided rocket in X, Y, Z three directions into an inertial navigation device, and resolving the acceleration and position information of the guided rocket by the inertial navigation device;
when the pitch angle of the guided rocket is smaller than-75 degrees within the range of-90 degrees or smaller than 75 degrees and smaller than 90 degrees within the range of 75 degrees, the speed increment of the guided rocket in the X, Y, Z three directions, the pitch angle speed increment, the yaw angle speed increment and the roll angle speed increment are all injected into an inertial navigation system device, and the inertial navigation device calculates to obtain the acceleration and position information of the guided rocket.
Further, the simulation computer sends a state mark to a set address of the real-time network according to the attitude angle of the three-axis turntable, and the injection computer reads the state mark and executes a corresponding test mode.
Has the advantages that:
when the use boundary of the three-axis rotary table is limited, the injection computer identifies the zone bit of the set address of the real-time network, controls the switching of the injection amount of the inertial navigation device, and realizes the halt of the rotary table but the continuation of the simulation test, thereby completing the semi-physical simulation test when the attitude angle of the guided rocket changes in a large range.
Drawings
FIG. 1 shows a simulation system constructed by the method of the present invention.
FIG. 2 is a flow chart of the method of the present invention.
Detailed Description
The invention is described in detail below by way of example with reference to the accompanying drawings.
The invention provides a semi-physical simulation test method under the condition of wide-range change of an attitude angle, which mainly comprises the following steps of (1) as shown in a figure 2;
1. and (3) building a simulation system as shown in figure 1, connecting all parts of the guided rocket normally, and preparing for a simulation test. The inertial navigation device initial state flag bit Bz is 0.
2. BZ is a status flag bit, in the invention, BZ is a status flag written into the real-time network set address by the simulation computer, and after the injection computer reads the flag bit on the real-time network set address, the three-component injection mode or the six-component injection mode is switched.
3. And starting a simulation system to perform a simulation test, wherein the state flag bit Bz of the inertial navigation device is 1 in a variation interval from-75 degrees to 75 degrees of the pitch angle of the pilot rocket. The pitching, yawing and rolling attitude angles and attitude angular velocities of the guided rocket are simulated through the attitude motion of the rotary table, and the simulation computer outputs velocity increment (namely a) of the guided rocket X, Y, Z in three directionsx、ay、az) And the acceleration and position information of the guided rocket is obtained by resolving the inertial navigation device according to the attitude motion of the three-axis turntable and the received speed increments.
4. When the attitude angle required by the flight trajectory of the guided rocket exceeds the using boundary of the rotary table (namely the pitch angle of the guided rocket is more than or equal to-90 degrees and less than-75 degrees or the pitch angle of the guided rocket is more than or equal to 75 degrees and less than or equal to 90 degrees), the rotary table stops running and the boundary angle is kept unchanged, at the moment, the injection computer reads the BZ mark on the preset address of the real-time network, at the moment, the state mark bit Bz of the inertial navigation device is 2, the injection computer is switched to a six-component injection mode from a three-component injection mode, and besides the injection of the speed increment in three directions of X, Y, Z, the pitch angle speed increment, the yaw angle speed increment and the roll angle speed increment (x、ωy、ωz) The attitude angle is realized by angular rate integration, and the inertial navigation device calculates the acceleration and position information of the guided rocket according to the injection data, so that the simulation test is completed.
According to the invention, when the attitude angle of the pilot rocket is changed in a large range, the zone bit is switched from three-component injection to six-component injection. And finishing the simulation test when the attitude angle of the guided rocket changes in a large range.
In summary, the above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention. 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 (2)

1. A semi-physical simulation test method under the condition of wide-range change of attitude angles is characterized in that when the pitch angle of a guided rocket is more than or equal to-75 degrees and less than or equal to 75 degrees, the attitude motion of a three-axis turntable is utilized to simulate the pitch attitude angle, the yaw attitude angle, the roll attitude angle and the corresponding attitude angular velocity of the guided rocket; injecting speed increments of the guided rocket in X, Y, Z three directions into an inertial navigation device, and resolving the acceleration and position information of the guided rocket by the inertial navigation device;
when the pitch angle of the guided rocket is smaller than-75 degrees within the range of-90 degrees or smaller than 75 degrees and smaller than 90 degrees within the range of 75 degrees, the speed increment of the guided rocket in the X, Y, Z three directions, the pitch angle speed increment, the yaw angle speed increment and the roll angle speed increment are all injected into an inertial navigation system device, and the inertial navigation device calculates to obtain the acceleration and position information of the guided rocket.
2. The method as claimed in claim 1, wherein the simulation computer sends the status flag to the address of the real-time network according to the attitude angle of the three-axis turntable, and the injection computer reads the status flag and executes the corresponding test mode.
CN202010076529.7A 2020-01-23 2020-01-23 Semi-physical simulation test method under condition of large-range change of attitude angle Active CN111221268B (en)

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114046806A (en) * 2021-12-03 2022-02-15 九江精密测试技术研究所 High-precision attitude simulation slewing device

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101363699A (en) * 2007-08-06 2009-02-11 孙再庸 Spinning stability fire-fighting rocket bomb, launching set thereof and system
JP2012106515A (en) * 2010-11-15 2012-06-07 Mitsubishi Space Software Kk Rocket guidance computing device, rocket guidance system, rocket, rocket guidance computing program, and rocket guidance computing method of rocket guiding computing device
GB201403426D0 (en) * 2014-02-27 2014-04-16 Atlantic Inertial Systems Ltd Inertial navigation system
CN110262282A (en) * 2019-05-27 2019-09-20 南京理工大学 Rocket grade, which is settled in an area, controls three-axle table semi-matter simulating system and method

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101363699A (en) * 2007-08-06 2009-02-11 孙再庸 Spinning stability fire-fighting rocket bomb, launching set thereof and system
JP2012106515A (en) * 2010-11-15 2012-06-07 Mitsubishi Space Software Kk Rocket guidance computing device, rocket guidance system, rocket, rocket guidance computing program, and rocket guidance computing method of rocket guiding computing device
GB201403426D0 (en) * 2014-02-27 2014-04-16 Atlantic Inertial Systems Ltd Inertial navigation system
CN110262282A (en) * 2019-05-27 2019-09-20 南京理工大学 Rocket grade, which is settled in an area, controls three-axle table semi-matter simulating system and method

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
杨永亮;吴小役;李颖;王斌翊;: "制导火箭弹飞行控制半实物仿真系统设计" *
王晨;董景新;高宗耀;杨栓虎;陈静;: "火箭弹大动态单轴平台惯导系统姿态算法" *

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114046806A (en) * 2021-12-03 2022-02-15 九江精密测试技术研究所 High-precision attitude simulation slewing device

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