CN113108686B - Strain measuring device for spacecraft and measuring method thereof - Google Patents
Strain measuring device for spacecraft and measuring method thereof Download PDFInfo
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- CN113108686B CN113108686B CN202110398632.8A CN202110398632A CN113108686B CN 113108686 B CN113108686 B CN 113108686B CN 202110398632 A CN202110398632 A CN 202110398632A CN 113108686 B CN113108686 B CN 113108686B
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01B—MEASURING LENGTH, THICKNESS OR SIMILAR LINEAR DIMENSIONS; MEASURING ANGLES; MEASURING AREAS; MEASURING IRREGULARITIES OF SURFACES OR CONTOURS
- G01B7/00—Measuring arrangements characterised by the use of electric or magnetic techniques
- G01B7/16—Measuring arrangements characterised by the use of electric or magnetic techniques for measuring the deformation in a solid, e.g. by resistance strain gauge
- G01B7/18—Measuring arrangements characterised by the use of electric or magnetic techniques for measuring the deformation in a solid, e.g. by resistance strain gauge using change in resistance
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Abstract
The invention provides a strain measurement device for a spacecraft and a measurement method thereof, wherein the strain measurement device is configured on a spacecraft platform and is used for sensing, conditioning and transmitting dynamic strain signals on an active section connecting ring, and then the dynamic interface force of an emission active section can be calculated according to the strain signals, so that a basis is provided for the structural design of the spacecraft and the ground mechanical test. The invention mainly solves the engineering problems of compact structure, light weight, high reliability requirement, suitability for high-frequency dynamic strain measurement, realization of online acquisition and the like of a measuring device in the process of identifying the dynamic load of the interface between the spacecraft and the carrier rocket. The validity of the device is also verified by the actual on-orbit application.
Description
Technical Field
The invention relates to the field of aerospace measurement, in particular to a strain measurement device for a spacecraft and a measurement method thereof.
Background
The load characteristics of the spacecraft and the carrying interface are important basis for the structural design of the spacecraft. The user manual of the main rocket at home and abroad specifies the three-way loading condition of the satellite-rocket interface in a quasi-static overload mode. However, this load spectrum, which is usually given in the carrying user manual, is relatively simple, reflecting only the maximum value, the load amplitude being completely decoupled from the frequency.
In recent years, the aerospace technical research institutions at home and abroad recognize that the traditional interface acceleration control can generate serious 'over-test' problems due to the fact that boundary conditions exist between the experimental state and the emission state of a spacecraft. Therefore, a strain measurement device is needed to realize the unification of load amplitude and frequency, and the aerospace applicability of the traditional strain measurement device (CN 2422617) and the three-dimensional strain measurement device (patent number: CN 102636105A) is not obvious. In order to ensure the stability of on-orbit data, the invention provides a spacecraft strain measurement device. Gradually paying attention to the phenomenon of 'over test' is relieved by increasing interface force control (force limit control) on the basis of satellite-arrow interface acceleration control. As the input of force limit control, accurately mastering the satellite-rocket interface load state in the launching stage has very important significance for the lightweight design and the sensitive load dynamic optimization design of the spacecraft structure.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a strain measurement device for a spacecraft and a measurement method thereof.
The invention provides a strain measurement device for a spacecraft, which comprises a temperature self-compensation strain gauge, a power conditioning module, a strain bridge module and an operational amplifier module, wherein:
the power conditioning module receives an external power supply and supplies power to the strain bridge module and the operational amplifier module;
the temperature self-compensating strain gauge is arranged on the active section connecting ring of the spacecraft, detects dynamic strain signals, and outputs the dynamic strain signals through the strain bridge circuit module and the operational amplifier module in sequence.
Preferably, the power conditioning device further comprises a power module, wherein the power module provides direct current to the power conditioning module.
Preferably, the system further comprises an A/D sampling module, wherein the A/D sampling module receives the dynamic strain signal output by the operational amplifier module.
Preferably, the operational amplifier module comprises a differential amplifier, wherein the homodromous input end and the reverse input end of the differential amplifier are connected with the output end of the strain bridge module, and the output end of the differential amplifier is connected with the A/D sampling module.
Preferably, the temperature self-compensating strain gauge comprises a self-compensating strain gauge for measuring three directions of a measuring point of an active segment connecting ring of the spacecraft.
Preferably, the strain bridge module comprises a full-bridge circuit formed by a temperature self-compensation strain gauge and a fixed resistor, the input end of the strain bridge module is connected with the power conditioning module, and the output end of the strain bridge module is connected with the operational amplifier circuit.
Preferably, the temperature self-compensating strain relief module further comprises a shell, and the temperature self-compensating strain relief module, the power supply conditioning module, the strain bridge module and the operational amplifier module are all arranged in the shell.
Preferably, the temperature self-compensating strain relief module, the strain bridge module and the operational amplifier module are all provided with a plurality of groups, and each group corresponds to one measuring point on the spacecraft active section connecting ring.
Preferably, the temperature self-compensating strain relief is connected to the strain bridge module by a three-wire connection.
The invention also provides a measuring method based on the strain measuring device for the spacecraft, which comprises the following steps of:
and a signal acquisition step: the method comprises the steps that an acquisition instruction is received, and a strain measurement device acquires a dynamic strain signal for an active section connecting ring of a spacecraft;
and a signal processing step: the acquired dynamic strain signals are stored after being conditioned and converted.
Compared with the prior art, the invention has the following beneficial effects:
1. according to the strain measuring device, the temperature self-compensation three-way strain gauge is adopted on the spacecraft for the first time, and each strain gauge measuring circuit adopts a three-wire system connection method, so that the accuracy of high-frequency dynamic strain measurement is improved.
2. The method can accurately and efficiently acquire the dynamic strain signals on the connecting ring of the spacecraft and the carrier rocket, and acquire the dynamic interface force of the launching active section through calculation, thereby providing basis for the structural design of the spacecraft and the ground mechanical test.
3. The strain measuring device has compact structure, light weight and high reliability, and can realize on-line acquisition and forwarding. The validity of the device is also verified by the actual on-orbit application.
4. The method is used for sensing, conditioning and transmitting the dynamic strain signals on the spacecraft active section connecting ring, and according to the strain signals, the dynamic interface force of the launching active section spacecraft and the carrier rocket can be calculated, so that a basis is provided for the structural design of the spacecraft and the ground mechanical test.
Drawings
Other features, objects and advantages of the present invention will become more apparent upon reading of the detailed description of non-limiting embodiments, given with reference to the accompanying drawings in which:
FIG. 1 is a schematic diagram of a three-way strain relief.
FIG. 2 is a strain gage layout at the connection ring of a spacecraft and a launch vehicle.
Fig. 3 is a block diagram of a strain measurement system information flow.
Fig. 4 is a schematic diagram of the interior of the strain gauge.
Fig. 5 is an equivalent circuit schematic diagram of the three-wire system connection method.
Fig. 6 is a schematic diagram of a strain monitoring control box.
Fig. 7 is a schematic diagram of the working principle of the strain monitoring control box.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the present invention, but are not intended to limit the invention in any way. It should be noted that variations and modifications could be made by those skilled in the art without departing from the inventive concept. These are all within the scope of the present invention.
As shown in fig. 1 to 7, the invention provides a strain measurement device for a spacecraft and a measurement method thereof, wherein the strain measurement device is configured on a spacecraft platform and is used for sensing, conditioning and transmitting dynamic strain signals on an active section connecting ring, and calculating dynamic interface force of an emission active section according to the strain signals, so that basis is provided for structural design and ground mechanical test of the spacecraft. The invention mainly solves the engineering problems of compact structure, light weight, high reliability requirement, suitability for high-frequency dynamic strain measurement, realization of online measurement and the like of a measuring device in the interface dynamic load identification process.
The invention provides a strain measurement device for a spacecraft, which comprises a temperature self-compensating strain gauge, a power management module, a strain bridge module, an operational amplifier module, a single machine shell and accessory accessories.
The strain in three directions is measured by using a temperature self-compensating three-way strain gauge (0 DEG, 45 DEG and 90 DEG) at each measuring point to calculate the longitudinal strain, the circumferential strain and the shear strain of the point, and the self-compensating strain gauge used in the invention can be applied to a strain gauge with a linear expansion coefficient of 23 multiplied by 10 -6 m·m -1 ·℃ -1 The method for selecting the resistance temperature coefficient of the strain gauge sensitive gate material of the connecting ring between the spacecraft and the carrier rocket, which is made of the aluminum alloy material, comprises the following steps:
the temperature change has a significant effect on all properties of the strain gauge, the most important of which is the false output of the strain gauge due to the temperature change, commonly referred to as false output, i.e. thermal output, expressed as:
ε T ＝[α R /K+(α s -α g )]ΔT s
wherein: epsilon T Is the heat output of the strain gauge, alpha R The temperature coefficient of resistance of the strain gage sensitive gate material, K is the sensitivity coefficient of the strain gage, alpha s Is the linear expansion coefficient of the test piece material, alpha g The linear expansion coefficient T of the strain gage sensitive gate material s Is the temperature of the test piece.
Let epsilon T =0, the temperature coefficient of resistance of the self-compensating strain gauge sensitive gate material can be calculated: alpha R ＝K(α g -α s ) Each strain gauge measuring circuit adopts a three-wire system connection method so as to avoid the influence of the use of long wires on bridge balance.
As can be seen from fig. 5, the bridge output port introduces a wire resistance r. The bridge can be considered approximately as an open circuit output, considering that the output of the bridge will be connected to an instrumentation differential amplifier with a very large input resistance (ideally infinity), and therefore its effect on the bridge is negligible. The total resistance of the 1 st and 2 nd bridge arms (numbered clockwise from the strain gauge) both comprise the resistance r of the equal length wire, at which time the bridge remains balanced without external forces. When an external force acts on the bridge, the error relative to an ideal bridge is specifically analyzed as follows:
the total resistance of the strain gauge is R (1+δ) assuming that the resistance change of the strain gauge due to an external force is Δr=δ·r (δ++ε, ε is the measured strain). The ideal bridge output without regard to wire resistance is:
the bridge output in the three-wire system connection method is as follows:
the error between the two is:
the strain gauge with a large nominal resistance value can reduce errors caused by long wires, and in addition, since the resistance r of the connected long wires is smaller and is in the same temperature field, the resistance changes of the wires caused by temperature change can be considered to be synchronous, so that the balance of the bridge is not influenced.
The strain measurement device is used for sensing, collecting and conditioning dynamic strain signals on the spacecraft active section connecting ring. The specific measurement process is that a system starts working according to program control instructions 5 minutes before the spacecraft is launched, a dynamic strain gauge senses and collects strain signals on a satellite rocket connecting ring of a satellite active section, analog signals are transmitted to a dynamic strain front-mounted conditioning box to be subjected to signal conditioning and conversion, the converted signals are transmitted to a strain monitoring control box to be received and stored, the whole active section is collected for about 30 minutes, the collection is stopped after the orbit separation is completed, and then complete strain data of the whole launching section is obtained through on-satellite data downloading. The strain monitoring control box performs A/D conversion and storage on the regulated strain signals, and performs data transmission after obtaining satellite comprehensive electronic instructions so as to download the satellite comprehensive electronic instructions to the ground station through the antenna.
The foregoing describes specific embodiments of the present invention. It is to be understood that the invention is not limited to the particular embodiments described above, and that various changes or modifications may be made by those skilled in the art within the scope of the appended claims without affecting the spirit of the invention. The embodiments of the present application and features in the embodiments may be combined with each other arbitrarily without conflict.
Claims (6)
1. The utility model provides a strain measurement device for spacecraft, its characterized in that includes temperature self-compensating strain relief, power conditioning module, strain bridge module and fortune and put the module, wherein:
the power conditioning module receives an external power supply and supplies power to the strain bridge module and the operational amplifier module;
the temperature self-compensating strain gauge is arranged on the active section connecting ring of the spacecraft, and is used for detecting a dynamic strain signal and outputting the dynamic strain signal sequentially through the strain bridge circuit module and the operational amplifier module;
the temperature self-compensating strain gauge comprises a self-compensating strain gauge for measuring three directions of a measuring point of an active section connecting ring of the spacecraft;
the strain bridge circuit module comprises a full-bridge circuit formed by temperature self-compensation strain relief and fixed resistors, the input end of the strain bridge circuit module is connected with the power supply conditioning module, and the output end of the strain bridge circuit module is connected with the operational amplifier circuit;
the temperature self-compensating strain gauge is connected into the strain bridge circuit module through a three-wire system connection method;
the resistance change of the strain gauge under the action of external force is delta R=delta.R, delta is delta-oc epsilon, epsilon is the measured strain, the total resistance of the strain gauge is R (1+delta), and the ideal bridge output without considering the resistance of the lead is as follows:
the bridge output in the three-wire system connection method is as follows:
the error between the two is:
the temperature self-compensation strain relief module, the strain bridge circuit module and the operational amplifier module are all provided with a plurality of groups, and each group corresponds to one measuring point on the spacecraft active section connecting ring.
2. The strain measurement device for a spacecraft of claim 1, further comprising a power module that provides direct current to the power conditioning module.
3. The strain measurement device for a spacecraft of claim 1, further comprising an a/D sampling module that receives the dynamic strain signal output by the op-amp module.
4. A strain measurement device for a spacecraft as claimed in claim 3, wherein the op-amp module comprises a differential amplifier, the co-directional input and the counter-directional input of the differential amplifier being connected to the output of the strain bridge module, the output of the differential amplifier being connected to the a/D sampling module.
5. The strain measurement device for a spacecraft of claim 1, further comprising a housing, wherein the temperature self-compensating strain relief, the power conditioning module, the strain bridge module, and the op-amp module are all mounted within the housing.
6. A measurement method based on a strain measurement device for a spacecraft according to claims 1-5, characterized by the steps of:
and a signal acquisition step: the method comprises the steps that an acquisition instruction is received, and a strain measurement device acquires a dynamic strain signal for an active section connecting ring of a spacecraft;
and a signal processing step: the acquired dynamic strain signals are stored after being conditioned and converted.
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| CN101487748B (en) * | 2006-06-30 | 2012-05-09 | 合肥工业大学 | Shearing force measurement bridge circuit in bending-twisting combined test apparatus |
| CN100494934C (en) * | 2006-06-30 | 2009-06-03 | 合肥工业大学 | Bending and twisting combined test device and its use in measuring internal force |
| CN102095596A (en) * | 2011-01-11 | 2011-06-15 | 中南大学 | Real-time temperature compensation method of bridge fatigue life gauge |
| CN105277111B (en) * | 2014-07-04 | 2018-03-16 | 北京强度环境研究所 | Satellite and the rocket locking device strain monitoring system |
| CN104296897B (en) * | 2014-09-12 | 2016-08-17 | 上海卫星工程研究所 | The satellite and the rocket six degree of freedom interfacial force computational methods of ring strain measurement are connected based on the satellite and the rocket |
| CN106323156B (en) * | 2016-08-06 | 2018-12-18 | 太原理工大学 | Frequency hopping spread spectrum communication means based on the adjustable wireless strain sensing device of bridge |
| CN212253981U (en) * | 2020-05-26 | 2020-12-29 | 杭州知愚科技有限公司 | Temperature test and compensation system of strain sensor |
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