CN104973268A - High-frequency micro-vibration isolation device of spacecraft control moment gyroscope - Google Patents
High-frequency micro-vibration isolation device of spacecraft control moment gyroscope Download PDFInfo
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Abstract
The invention provides a high-frequency micro-vibration isolation device of a spacecraft control moment gyroscope. The high-frequency micro-vibration isolation device is of a central symmetry structure and comprises an installation plate (1), a bottom plate (3) and vibration isolation units arranged between the installation plate (1) and the bottom plate (3). Each vibration isolation unit comprises a radial damping unit (4), an axial damping unit (5) and an elastic supporting unit (6). According to the novel micro-vibration isolation device, vibration isolation performance is good at CMG work rotation speed, six-degree-of-freedom micro vibration transferred to a star body can be effectively reduced, and posture stability and pointing accuracy of the star body can be improved.
Description
Technical field
The invention belongs to passive vibration isolation technical field, be specifically related to a kind of high frequency micro vibration spacer assembly of Spacecraft Control moment gyro.
Background technology
High Performance spacecraft is an important directions of contemporary space flight development, be implementation goal accurately identify with accurately locate, the Core equipment of the space mission such as spatial information highly effective and safe transmits, survey of deep space, in national defence army building, prevent and reduce natural disasters, the numerous areas such as resource environment all has eager application demand and wide application prospect.
High Performance spacecraft platform needs to have very high pointing accuracy and degree of stability, and a little deviation may will cause serious impact to High Performance spacecraft.Such as, for the earth observation satellite in 500km orbit altitude, the visual field offset that the angular oscillation of 2.06arcsec causes is up to 50m.Therefore, High Performance spacecraft is very responsive to small disturbance, needs a pulsation-free working environment when it performs space tasks in-orbit.
Micro-vibration is the Important Disturbed Factors affecting High Performance spacecraft pointing accuracy and degree of stability.When micro-vibration refers to that spacecraft in orbit, on star, rotatable parts High Rotation Speed, driver train step motion, thruster ignition operation, large-scale flexible part pass in and out the flutter that a kind of amplitude is less, frequency the is higher response that the risk factors such as shade produce.The vibrational energy of micro-vibration is more weak, can not destroy spacecraft structure, therefore, does not cause too much concern before this.But along with the fast development of High Performance spacecraft, the harm of micro-vibration more and more highlights, have a strong impact on pointing accuracy and the degree of stability of spacecraft platform.
Control moment gyroscope (Control Moment Gyroscope, be called for short CMG) is the important attitude control actuator of a class of spacecraft, and the spatial attitude adjustment that it utilizes momentum compensation principle to be spacecraft provides applied moment.But in the course of the work, the factor such as quiet, the unbalance dynamic characteristic of the high speed rotor of such actuating unit and the design defect of bearing, become the main source of micro-vibration on spacecraft, larger restrictive function has been caused to the high precision of spacecraft, high stability growth requirement.
As shown in Figure 1, for certain CMG in prior art is accompanied with the control torque output curve diagram of micro-obstacle of vibration, as can be seen from Figure 1, micro-vibration makes control torque curve produce small fluctuation, but for high-precision satellite equipment, this disturbance can to destroy relevant space mission.Therefore, micro-vibration that CMG produces is suppressed very necessary.
In the micro-vibration suppressing method of the numerous CMG of space engineering, between disturbing source and spacecraft body, place isolation mounting is a kind of efficient, practical technological means, as shown in Figure 2, is the schematic diagram adopting isolation mounting to suppress the micro-obstacle of vibration of CMG in prior art.
So far, there is large quantifier elimination for micro-vibration isolation mounting both at home and abroad.Typically comprise folding arm beam configuration isolation mounting of the viscous damping fluid isolation mounting of U.S.'s Hubble, Stewart configuration platform isolation mounting and India D.Kamesh proposition etc.As shown in Figure 3, be the roughly configuration picture of typical isolation mounting.
But existing micro-vibration isolation mounting, generally has the deficiencies such as the lower and high-frequency vibration isolation limited efficiency of configuration complexity, reliability and stability, thus constrains the use of micro-vibration isolation mounting.
Summary of the invention
For the defect that prior art exists, the invention provides a kind of high frequency micro vibration spacer assembly of Spacecraft Control moment gyro, can effectively solve the problem.
The technical solution used in the present invention is as follows:
The invention provides a kind of high frequency micro vibration spacer assembly of Spacecraft Control moment gyro, symmetrical structure centered by this high frequency micro vibration spacer assembly, comprising: adapter plate (1), base plate (3) and the vibration isolation unit be arranged between described adapter plate (1) and described base plate (3);
Described adapter plate (1) is cross structure, comprising: be positioned at the substrate (1-5) at center and be centrosymmetric outward extending 1st cantilever (1-1), the 2nd cantilever (1-2), the 3rd cantilever (1-3) and the 4th cantilever (1-4) from described substrate (1-5); Described substrate (1-5) offers mounting hole (1-6), by described mounting hole (1-6), is connected by described adapter plate (1) with vibration isolation object;
The below mode that is centrosymmetric of described 1st cantilever (1-1), described 2nd cantilever (1-2), described 3rd cantilever (1-3) and described 4th cantilever (1-4) arranges the 1st vibration isolation unit (2-1), the 2nd vibration isolation unit (2-2), the 3rd vibration isolation unit (2-3) and the 4th vibration isolation unit (2-4) respectively;
Wherein, for arbitrary i-th vibration isolation unit, it is arranged at the below of the i-th cantilever (1-i), wherein, and i=1,2,3 or 4;
Then: described i-th vibration isolation unit comprises: radial damping unit (4), axial damping unit (5) and elastic support unit (6);
Described radial damping unit (4) is arranged along the radial direction of described i-th cantilever (1-i), for providing the freedom of motion of horizontal direction, comprises radial damping layer (4-1) and radial constraint plate (4-2); Wherein, described radial constraint plate (4-2) comprises radial constraint substrate (4-2-1) and radial attaching parts (4-2-2) that is one-body molded with it and that be positioned at below it; Wherein, described radial attaching parts (4-2-2) divides into outer end and inner end;
The end face of described radial damping layer (4-1) is connected with the bottom surface of described i-th cantilever (1-i), and the bottom surface of described radial damping layer (4-1) is connected with described radial constraint substrate (4-2-1);
Described axial damping unit (5) is arranged along the axial direction of described i-th cantilever (1-i), for providing the freedom of motion of vertical direction, comprise left axial constraint plate (5-1), right axial constraint plate (5-2), left axial damping layer (5-3), right axial damping layer (5-4) and axial shear plate (5-5);
By from one end to the other side direction, end to endly set gradually described left axial constraint plate (5-1), described left axial damping layer (5-3), described axial shear plate (5-5), described right axial damping layer (5-4) and described right axial constraint plate (5-2);
And, the top of described axial shear plate (5-5) is integrally formed is provided with axial connecting part (5-5-1), described axial connecting part (5-5-1) for being fixedly connected with the inner end of described radial attaching parts (4-2-2), and then described axial damping unit (5) is fixedly installed to described radial damping unit (4) below and near center;
The bottom of described left axial constraint plate (5-1) is provided with left axial constraint attaching parts (5-1-1), the bottom of described right axial constraint plate (5-2) is provided with right axial constraint attaching parts (5-2-1); By described left axial constraint attaching parts (5-1-1) and described right axial constraint attaching parts (5-2-1), described axial damping unit (5) is fixed to described base plate (3);
Described elastic support unit (6) for providing the rigidity of vertical direction, and ensures that spacer assembly has certain torsional stiffness and roll stiffness, avoids the normal attitude control torque affecting control moment gyroscope to export; The top of described elastic support unit (6) is fixed to the outer end of described radial attaching parts (4-2-2), and the bottom of described elastic support unit (6) is fixed to described base plate (3);
The upper surface of described base plate (3) is used for being connected with described vibration isolation unit, provides the support of described vibration isolation unit; The lower surface of described base plate (3) is used for being connected and fixed with protected object.
Preferably, described adapter plate (1) is spring plate.
Preferably, for described radial damping unit (4), the described radial damping layer (4-1) that it comprises is viscoelastic material;
For described axial damping unit (5), its described left axial damping layer (5-3) comprised and described right axial damping layer (5-4) are viscoelastic material.
Preferably, between described radial damping unit (4) and described i-th cantilever (1-i), between described radial damping unit (4) and described radial constraint substrate (4-2-1), between described left axial damping layer (5-3) and described left axial constraint plate (5-1), between described left axial damping layer (5-3) and described axial shear plate (5-5), between described right axial damping layer (5-4) and described axial shear plate (5-5), integrated vulcanizing forming technique is all adopted to carry out between described right axial damping layer (5-4) and described right axial constraint plate (5-2) bonding.
Preferably, described radial attaching parts (4-2-2) divides into outer end and inner end, and, offer outer connecting bore in the outer end of described radial attaching parts (4-2-2); Interconnecting hole is offered at the inner end of described radial attaching parts (4-2-2);
The described axial connecting part (5-5-1) of described axial damping unit (5) offers axial connecting bore; Described axial connecting bore and described interconnecting hole are bolted fixing;
The top of described elastic support unit (6) is provided with elastic support connecting bore, and described elastic support connecting bore and described outer connecting bore are bolted fixing.
Preferably, described elastic support unit (6) is folding arm beam.
The high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention has the following advantages: have that configuration is simple, reliability and stability is high and high-frequency vibration isolation is an effective advantage.
Accompanying drawing explanation
Fig. 1 is the control torque output curve diagram that in prior art, certain CMG is accompanied with micro-obstacle of vibration;
Fig. 2 is the schematic diagram adopting isolation mounting to suppress the micro-obstacle of vibration of CMG in prior art;
Fig. 3 is the roughly configuration picture of typical isolation mounting;
Fig. 4 is transmissibility schematic diagram provided by the invention;
Fig. 5 is the transport diagram of curves of passive vibration isolation provided by the invention;
Fig. 6 is the overall perspective view of the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention;
Fig. 7 is the schematic diagram after marking in detail in figure 6;
Fig. 8 is the lateral plan of the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention;
Fig. 9 is the birds-eye view of the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention;
Figure 10 is the structural representation of adapter plate;
Figure 11 is the assembled state schematic diagram of any one vibration isolation unit;
Figure 12 is the structural representation of radial damping unit;
Figure 13 is the structural representation of axial damping unit;
Figure 14 is the schematic diagram of each bonding plane in radial damping unit;
Figure 15 is each bonding plane schematic diagram of axial damping unit;
Figure 16 is damping unit principle schematic;
Figure 17 is elastic support Unit Design constructional drawing;
Figure 18 is the structural representation of base plate;
Figure 19 is base plate and vibration isolation unit schematic diagram with bolts;
Figure 20 is that power/moment transport calculates schematic diagram;
Figure 21 is the F obtained after Frequency Response Analysis
xdirection transport correlation curve figure;
Figure 22 is the F obtained after Frequency Response Analysis
ydirection transport correlation curve figure;
Figure 23 is the F obtained after Frequency Response Analysis
zdirection transport correlation curve figure;
Figure 24 is the M obtained after Frequency Response Analysis
xdirection transport correlation curve figure;
Figure 25 is the M obtained after Frequency Response Analysis
ydirection transport correlation curve figure;
Figure 26 is the M obtained after Frequency Response Analysis
zdirection transport correlation curve figure;
Figure 27 is the composition and working principle figure of measuring system;
Figure 28 be under working speed 6000rpm disturbance force at F
xdirection time-domain curve figure;
Figure 29 be under working speed 6000rpm disturbance force at F
xdirection frequency curve figure;
Figure 30 be under working speed 6000rpm disturbance force at F
ydirection time-domain curve figure;
Figure 31 be under working speed 6000rpm disturbance force at F
ydirection frequency curve figure;
Figure 32 be under working speed 6000rpm disturbance force at F
zdirection time-domain curve figure;
Figure 33 be under working speed 6000rpm disturbance force at F
zdirection frequency curve figure;
Figure 34 be under working speed 6000rpm disturbance force at M
xdirection time-domain curve figure;
Figure 35 be under working speed 6000rpm disturbance force at M
xdirection frequency curve figure;
Figure 36 be under working speed 6000rpm disturbance force at M
ydirection time-domain curve figure;
Figure 37 be under working speed 6000rpm disturbance force at M
ydirection frequency curve figure;
Figure 38 be under working speed 6000rpm disturbance force at M
zdirection time-domain curve figure;
Figure 39 be under working speed 6000rpm disturbance force at M
zdirection frequency curve figure.
Detailed description of the invention
Below in conjunction with accompanying drawing, the present invention is described in detail:
The invention provides a kind of high frequency micro vibration spacer assembly of Spacecraft Control moment gyro, adopt passive vibration isolation technology, it has that configuration is simple, reliability and stability are high, high-frequency vibration isolation is effective and without the need to advantages such as energy inputs.
Its general principles is:
Adopt the transmissibility T in frequency domain, namely by the Input Forces of isolation mounting and the ratio of power output, evaluate the vibration isolating effect of isolation mounting.
Concrete, as shown in Figure 4, be transmissibility schematic diagram provided by the invention; For isolation mounting, transmissibility T=f
out/ f
in; Wherein, f
infor the input disturbance power of isolation mounting; f
outfor the output disturbance power of isolation mounting.Visible, when transmissibility T is less than 1, when being namely less than input disturbance power by the output disturbance power of isolation mounting, think that disturbance is effectively isolated, isolation mounting has played effect.
As shown in Figure 5, be the transport diagram of curves of passive vibration isolation.In the present invention, the initial frequency point of effective vibration isolation scope is defined as effective isolation frequency f
e, namely as the frequency>=f of disturbance input
etime, transmissibility T≤1 must be had.For single degree of freedom passive vibration isolation system,
f
nfor the natural frequency of vibrating isolation system.Therefore, in passive vibration isolation system, the effective isolation frequency of vibrating isolation system reduces with the reduction of natural frequency, and effective vibration isolation scope expands with the reduction of natural frequency.
Be it can also be seen that by Fig. 5, the dampingratioζ of vibrating isolation system on the impact of transport curve shape obviously, is mainly reflected in the rate of decay after transport peak of curve size and vibration isolation.Damping ratio is larger, and transport peak value is less, and the rate of decay after vibration isolation is relatively slack-off.
Based on above-mentioned principle, the novel spacer assembly of the micro-vibration of Spacecraft Control moment gyro that this patent proposes is elasticity-Visco elastical complex structure, with reference to figure 6, it is the overall perspective view of the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention; With reference to figure 7, it is the schematic diagram after marking in detail in figure 6; With reference to figure 8, it is the lateral plan of the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention; With reference to figure 9, it is the birds-eye view of the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention.
Composition graphs 6-Fig. 9, the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention, symmetrical structure centered by this high frequency micro vibration spacer assembly, comprising: adapter plate 1, base plate 3 and the vibration isolation unit be arranged between adapter plate 1 and base plate 3.Below respectively adapter plate 1 provided by the invention, base plate 3 and vibration isolation unit are described in detail respectively:
(1) adapter plate
Adapter plate is the most top layer part of isolation mounting, and it is the spring plate through quenching, and quenching technical mainly reduces the rigidity of steel plate.Adapter plate thickness is even, as shown in Figure 10, for the structural representation of adapter plate, as seen from Figure 10, adapter plate is cross structure, comprising: be positioned at the substrate 1-5 at center and be centrosymmetric outward extending 1st cantilever 1-1, the 2nd cantilever 1-2, the 3rd cantilever 1-3 and the 4th cantilever 1-4 from substrate 1-5; Wherein, substrate 1-5 offers mounting hole 1-6, by mounting hole 1-6, can mode with bolts, and be connected with vibration isolation object by adapter plate 1, vibration isolation object is control moment gyroscope.The effect of adapter plate is mainly reflected in: for vibration isolation object provides erecting stage, and connects other parts of vibration isolation object and isolation mounting.
(2) vibration isolation unit
In the present invention, be provided with 4 vibration isolation units altogether, be respectively the 1st vibration isolation unit 2-1, the 2nd vibration isolation unit 2-2, the 3rd vibration isolation unit 2-3 and the 4th vibration isolation unit 2-4, further, the below mode of being centrosymmetric that 4 vibration isolation units are arranged at the 1st cantilever 1-1, the 2nd cantilever 1-2, the 3rd cantilever 1-3 and the 4th cantilever 1-4 is respectively arranged.
For arbitrary i-th vibration isolation unit, it is arranged at the below of the i-th cantilever, wherein, and i=1,2,3 or 4;
Then: the i-th vibration isolation unit comprises: radial damping unit 4, axial damping unit 5 and elastic support unit 6.As shown in figure 11, be the assembled state schematic diagram of any one vibration isolation unit.As a kind of example, as can be seen from Figure 11, as can be seen from Figure 11, in the present invention, elastic support unit is in parallel with axial damping unit, and common support plays radial damping unit, and, mode with bolts between elastic support unit and radial damping unit; Mode with bolts between the radial damping unit of axial damping unit.
Introduce structure and the principle of radial damping unit 4, axial damping unit 5 and elastic support unit 6 below respectively:
(2.1) radial damping unit
As shown in figure 12, be the structural representation of radial damping unit; Radial damping unit 4 is arranged along the radial direction of the i-th cantilever 1-i, for providing the freedom of motion of horizontal direction, comprises radial damping layer 4-1 and radial constraint plate 4-2; Wherein, radial constraint plate 4-2 comprises radial constraint substrate 4-2-1 and radial attaching parts 4-2-2 that is one-body molded with it and that be positioned at below it; Wherein, radial attaching parts 4-2-2 divides into outer end and inner end; By radial attaching parts 4-2-2, adapter shaft arranged side by side is to damping unit 5 and elastic support unit 6.
(2.2) axial damping unit
As shown in figure 13, be the structural representation of axial damping unit; Axial damping unit 5 is arranged along the axial direction of the i-th cantilever 1-i, for providing the freedom of motion of vertical direction, comprises left axial constraint plate 5-1, right axial constraint plate 5-2, left axial damping layer 5-3, right axial damping layer 5-4 and axial shear plate 5-5;
By from one end to the other side direction, end to endly set gradually left axial constraint plate 5-1, left axial damping layer 5-3, axial shear plate 5-5, right axial damping layer 5-4 and right axial constraint plate 5-2;
And, the top of axial shear plate 5-5 is integrally formed is provided with axial connecting part 5-5-1, axial connecting part 5-5-1 is used for being fixedly connected with the inner end of radial attaching parts 4-2-2, such as, mode with bolts, so axial damping unit 5 is fixedly installed to radial damping unit 4 below and near center;
The bottom of left axial constraint plate 5-1 is provided with left axial constraint attaching parts 5-1-1, the bottom of right axial constraint plate 5-2 is provided with right axial constraint attaching parts 5-2-1; By left axial constraint attaching parts 5-1-1 and right axial constraint attaching parts 5-2-1, axial damping unit 5 is fixed to base plate 3.
Below introduce the principle of radial damping unit provided by the invention and axial damping unit performance shock damping action:
Viscoelastic material is a kind of typical damping material, has the feature of viscous liquid and elastic solid, can be regarded as energy savings and waste of power material combines with certain proportion.When subjected to external, vibration mechanical energy image position can store by elastic part like that, can discharge, deformation-recovery after external force removes; Changes mechanical energy is then that thermal energy consumption dissipates by stickiness part.This absorption to energy and dissipation are exactly the shock damping action of viscoelastic material.Viscoelastic damping material obtains the attention of height in space industry, and domestic and international many space flight departments and scholar conduct in-depth research the theoretical analysis of viscoelastic damping material and embody rule.As a kind of efficient, ripe, cheap vibration suppression technology, it is widely used in space engineering.
Radial damping unit provided by the invention and axial damping unit, can be referred to as damping unit, comprises three damping layers altogether, be respectively: radial damping layer 4-1, left axial damping layer 5-3 and right axial damping layer 5-4; Further, these three damping layers are viscoelastic material, thus play damping isolation effect.
In addition, no matter for radial damping unit of the present invention, or for axial damping unit, the present invention is all designed to a kind of special configuration, as shown in figure 16, is damping unit principle schematic; As seen from Figure 16, damping unit is made up of restraint layer 101, damping layer 102 and shear layer 103 3 part, and 104 represent shear deformation scope; Wherein, damping layer is viscoelastic material, and restraint layer and shear layer are metallic material, and what this device adopted is common iron, because itself and viscoelastic material adhesive property are good.Under external excitation, when shear layer and restraint layer generation relative motion, drive viscoelastic material to produce shear deformation, viscoelastic material can effectively absorb, dissipate vibrational energy, reduce the transmission of micro-vibrational energy, for isolation mounting provides structural damping.
For the present invention, for radial damping unit, it radial damping layer 4-1 comprised adopts viscoelastic material, namely bears the effect of damping layer; And the cantilever be arranged at respectively above and below radial damping layer 4-1 and radial constraint plate, then bear the effect of shear layer and restraint layer respectively.Therefore, the configuration of Figure 16 is met.
For axial damping unit 5, by from one end to the other side direction, end to endly set gradually left axial constraint plate 5-1, left axial damping layer 5-3, axial shear plate 5-5, right axial damping layer 5-4 and right axial constraint plate 5-2; As can be seen from this kind of structure, comprise the configuration of two Figure 16 altogether, that is: (1) left axial damping layer 5-3 bears damping layer effect; Lay respectively at the left axial constraint plate of left axial damping layer 5-3 both sides and axial shear plate, bear the effect of restraint layer and shear layer respectively.(2) right axial damping layer 5-4 bears damping layer effect; Lay respectively at the axial shear plate of right axial damping layer 5-4 both sides and right axial constraint plate, bear the effect of shear layer and restraint layer respectively.
This shows, in the present invention, damping layer all adopts viscoelastic material, its both sides are fixed with restraint layer and shear layer respectively, restraint layer and shear layer are referred to as metal piece, in specific implementation, it is bonding that viscoelastic material and metal piece can adopt integrated vulcanizing forming technique to carry out, and has higher bind strength between viscoelastic material and metal piece by bonding by Vulcanization.
Concrete, as shown in figure 14, be the schematic diagram of each bonding plane in radial damping unit; As shown in figure 15, be each bonding plane schematic diagram of axial damping unit.In figures 14 and 15, the face of arrow points is bonding plane.Radial damping unit 4 with between the i-th cantilever 1-i, between radial damping unit 4 with radial constraint substrate 4-2-1, between left axial damping layer 5-3 and left axial constraint plate 5-1, between left axial damping layer 5-3 and axial shear plate 5-5, between right axial damping layer 5-4 and axial shear plate 5-5, all to adopt integrated vulcanizing forming technique to carry out between right axial damping layer 5-4 and right axial constraint plate 5-2 bonding.
(2.3) elastic support unit
Elastic support unit mainly provides the supporting role of Low rigidity; for the principle prototype of ground, guarantee that isolation mounting can bear the Action of Gravity Field of CMG, for AEROSPACE APPLICATION exactly; be then the isolation mounting in protection emission process, avoid vibrating and the destruction of overload to it.Therefore, elastic support element stiffness will be applicable to, if rigidity is excessive; micro-vibration overwhelming majority of CMG will be transmitted to spacecraft platform by support unit; isolation mounting does not have vibration isolating effect, and if rigidity is too small, then do not have the support to CMG and protective effect.
As shown in figure 17, the top of elastic support unit 6 is fixed to the outer end of radial attaching parts 4-2-2 to elastic support Unit Design, specifically can mode with bolts; The bottom of elastic support unit 6 is fixed to base plate 3, herein, and also can mode with bolts.Elastic support unit is folding arm beam configuration, and material is spring steel, and it can provide the Low rigidity of vertical direction, by changing width and the thickness of folding arm beam cross-sectional plane, can increase or reduce the rigidity size of elastic support unit.On the other hand, this configuration also can ensure that the torsional stiffness of vibrating isolation system and roll stiffness can not be too low, exports in order to avoid affect CMG gesture stability moment.
(3) base plate
Base plate is the bottom part of isolation mounting, is aluminium alloy plate, plate has many tapped bore and through hole, as shown in figure 18, is the structural representation of base plate; Base plate mainly plays the effect of connection, and wherein tapped bore is the connecting bore of elastic support unit and axial constraint plate, and through hole is the connecting bore of the protected object such as isolation mounting and spacecraft body.As shown in figure 19, base plate and vibration isolation unit schematic diagram with bolts is.
The high frequency micro vibration spacer assembly of Spacecraft Control moment gyro provided by the invention, its vibration isolation mechanism is: because the rigidity of folding arm beam and viscoelastic material is all lower, natural frequency is less, therefore, six-freedom degree all has good passive vibration isolation performance.This isolation mounting is equivalent to a low-pass filter, while guarantee CMG low frequency control torque passes through, can the disturbance of effective filter high-frequency rate, and reduce the impact of micro-vibration on satellite platform.
This isolation mounting specifically has the following advantages:
1, adopt elastic_viscoelastic composite structure design, introduce viscoelastic material and structural damping is provided, play and absorb and the effect of dissipate vibrational energy, while isolation high frequency components, effectively can weaken the disturbance response of low frequency range;
2, propose a kind of brand-new isolation mounting configuration, break through the typical vector free axis method of viscoelastic material in the past and restriction damping layer configuration, make full use of this most actv. oscillation damping and energy dissipating form of viscoelastic material shear deformation;
3, this isolation mounting is passive vibration isolation, technology is simple and reliable, volume and quality little, on spacecraft original structure impact little, without the need to extra power supply, be applicable to the application of spacecraft;
4, this isolation mounting configuration has higher application expansion, and expection can be applicable to micro-isolating technique of other high-speed moving parts such as spacecraft flywheel.
L-G simulation test example
For CMG, the main frequency of its micro-vibrational perturbation is identical with the rotating speed of CMG high speed rotor, and the rotating speed of CMG is generally in 6000 revs/min, i.e. 100Hz, and therefore the analysis & verification of this device anti-vibration performance is also concerned about the frequency domain near 100Hz.
(1) simulation calculation
Set up the finite element dynamics of micro-vibration isolating device and CMG.Wherein folding arm girder construction uses beam element modeling, and other parts all use solid element modeling, to improve model design accuracy.Base plate is thought and is belonged to spacecraft body, can not embody in finite element model.
Consider the frequency sex change of viscoelastic material dynamic property (Storage modulus and dissipation factor) in dynamics analysis, by STRUCTURAL SENSITIVITY ANALYSIS INDESIGN and Parameters Optimal Design, optimize the scantling of structure of isolation mounting further, improve anti-vibration performance.Modal strain energy method of iteration is utilized to calculate the natural frequency of vibrating isolation system, as shown in table 1.Because the 7th rank natural frequency is greater than 400Hz, not within care frequency range herein, therefore do not list in table.
6 rank natural frequencys before the vibrating isolation system of table 1 simulation calculation
In finite element model, using the entrance point of the rotor node of CMG as disturbance, using the output point of the restraint joint of isolation mounting lower end as disturbance.Input disturbance power F
inwith distrubing moment M
in, by the frequency response analysis of model, the support reaction F of output node can be obtained
outwith branch counter moment M
out, can obtain further by after isolation mounting, the transmissibility T of disturbance
f=F
out/ F
in, the transmissibility T of moment
m=M
out/ M
in, as shown in figure 20, for power/moment transport calculates schematic diagram.In fig. 20, P1 represents rotor node, i.e. entrance point; P2 represents outside frame node; P3 represents output node.
Design of Structural parameters is carried out to isolation mounting, on rotor node, inputs six-freedom degree direction (F successively subsequently
x, F
y, F
z, M
x, M
y, M
z) unit force/moment; Wherein, F
x, F
yand F
zrepresent the power in x, y and z direction respectively; M
x, M
yand M
zrepresent the moment in x, y and z direction respectively; Elect analysis frequency domain as 0-150Hz.The power obtained after Frequency Response Analysis/moment transport curve is as shown in Figure 21-Figure 26, and wherein, Figure 21 is the F obtained after Frequency Response Analysis
xdirection transport correlation curve figure; Figure 22 is the F obtained after Frequency Response Analysis
ydirection transport correlation curve figure; Figure 23 is the F obtained after Frequency Response Analysis
zdirection transport correlation curve figure; Figure 24 is the M obtained after Frequency Response Analysis
xdirection transport correlation curve figure; Figure 25 is the M obtained after Frequency Response Analysis
ydirection transport correlation curve figure; Figure 26 is the M obtained after Frequency Response Analysis
zdirection transport correlation curve figure.Numerical value unit is decibel (dB), and effective isolation frequency in each degree of freedom direction is as shown in table 2.Also provide a comparison the transport with or without considering viscoelastic material damping in Figure 21-Figure 26, wherein dotted line is the shock damping action not considering viscoelastic material.
The effective isolation frequency on rear six direction optimized by table 2
From transport diagram of curves and limited isolation frequency table, after optimizing, maximum effective isolation frequency of isolation mounting is 28.97Hz, namely the disturbance that frequency is greater than 28.97Hz effectively can be isolated by this isolation mounting, and micro-oscillation frequency of single frame CMG provided by the invention is generally in 100-150Hz, therefore, the novel isolation mounting that the present invention proposes can effectively isolate the high frequency micro vibration that CMG six degree of freedom produces.
From the comparing result in Figure 21-Figure 26, the natural frequency of the structural damping of viscoelastic material not influential system, but greatly reduce the peak value of response of system to disturbance in resonance range, effectively reduce the infringement of disturbance to structure, the castering action of damping to vibration isolating effect is fairly obvious.
(2) experimental verification
Draw the schedule drawing of this micro-vibration isolating device, and complete processing in kind and assembling, build and obtain ground microvibration measuring system, experimental system on land is made up of Kistler 9253B12 test desk, optics vibration-isolating platform and Data collecting and analysis system, and the composition and working principle of measuring system as shown in figure 27.Wherein, Kistler Table is widely used in disturbance force and the distrubing moment of measuring flywheel, number passes the aerospace components such as pedestal, have measurement range large, measure that frequency domain is wide, survey precision advantages of higher.
With the disturbance input of the counteraction flyback simulation CMG of high speed rotating.Confirmatory experiment comprises two kinds of operating modes, operating mode 1: be that flywheel and Kistler Table platform are directly rigidly connected by bolt, operating mode 2: be that flywheel is connected with Kistler Table platform indirectly via after isolation mounting.
Start flywheel, its rotating speed is progressively adjusted to working speed (6000rpm), measure the force and moment under two kinds of connection operating modes, obtain the time domain disturbance delivery curve on six degree of freedom, by Fourler transform, obtain the disturbance delivery curve in frequency domain, concrete, as shown in figure 28, for disturbance force under working speed 6000rpm is at F
xdirection time-domain curve figure; As shown in figure 29, for disturbance force under working speed 6000rpm is at F
xdirection frequency curve figure; As shown in figure 30, for disturbance force under working speed 6000rpm is at F
ydirection time-domain curve figure; As shown in figure 31, for disturbance force under working speed 6000rpm is at F
ydirection frequency curve figure; As shown in figure 32, for disturbance force under working speed 6000rpm is at F
zdirection time-domain curve figure; As shown in figure 33, for disturbance force under working speed 6000rpm is at F
zdirection frequency curve figure; As shown in figure 34, for disturbance force under working speed 6000rpm is at M
xdirection time-domain curve figure; As shown in figure 35, for disturbance force under working speed 6000rpm is at M
xdirection frequency curve figure;
As shown in figure 36, for disturbance force under working speed 6000rpm is at M
ydirection time-domain curve figure; As shown in figure 37, for disturbance force under working speed 6000rpm is at M
ydirection frequency curve figure; As shown in figure 38, for disturbance force under working speed 6000rpm is at M
zdirection time-domain curve figure; As shown in figure 39, for disturbance force under working speed 6000rpm is at M
zdirection frequency curve figure; Table 3 enumerates out the disturbance output valve of two kinds of operating modes under frequency of operation (100Hz).
Disturbance output valve contrast under table 3 frequency of operation (100Hz)
Disturbance result of a measurement under comparative analysis two kinds of operating modes is known:
(1), in time domain, the isolation mounting disturbance connected under operating mode exports specific rigidity connection operating mode and is significantly reduced, all fairly obvious on six-freedom degree.
In the frequency domain of (2) 0-350Hz, the isolation mounting disturbance connected under operating mode exports and is all less than the operating mode that is rigidly connected, and wide band disturbance can effectively be isolated.
In sum, simulation calculation and experimental verification all show, novel micro-vibration isolating device that the present invention proposes has good anti-vibration performance under CMG working speed, can effectively reduce the transmission of the micro-vibration of six-freedom degree to celestial body, contributes to the attitude stability and the pointing accuracy that improve celestial body.
The above is only the preferred embodiment of the present invention; it should be pointed out that for those skilled in the art, under the premise without departing from the principles of the invention; can also make some improvements and modifications, these improvements and modifications also should look protection scope of the present invention.
Claims (6)
1. the high frequency micro vibration spacer assembly of a Spacecraft Control moment gyro, it is characterized in that, symmetrical structure centered by this high frequency micro vibration spacer assembly, comprising: adapter plate (1), base plate (3) and the vibration isolation unit be arranged between described adapter plate (1) and described base plate (3);
Described adapter plate (1) is cross structure, comprising: be positioned at the substrate (1-5) at center and be centrosymmetric outward extending 1st cantilever (1-1), the 2nd cantilever (1-2), the 3rd cantilever (1-3) and the 4th cantilever (1-4) from described substrate (1-5); Described substrate (1-5) offers mounting hole (1-6), by described mounting hole (1-6), is connected by described adapter plate (1) with vibration isolation object;
The below mode that is centrosymmetric of described 1st cantilever (1-1), described 2nd cantilever (1-2), described 3rd cantilever (1-3) and described 4th cantilever (1-4) arranges the 1st vibration isolation unit (2-1), the 2nd vibration isolation unit (2-2), the 3rd vibration isolation unit (2-3) and the 4th vibration isolation unit (2-4) respectively;
Wherein, for arbitrary i-th vibration isolation unit, it is arranged at the below of the i-th cantilever (1-i), wherein, and i=1,2,3 or 4;
Then: described i-th vibration isolation unit comprises: radial damping unit (4), axial damping unit (5) and elastic support unit (6);
Described radial damping unit (4) is arranged along the radial direction of described i-th cantilever (1-i), for providing the freedom of motion of horizontal direction, comprises radial damping layer (4-1) and radial constraint plate (4-2); Wherein, described radial constraint plate (4-2) comprises radial constraint substrate (4-2-1) and radial attaching parts (4-2-2) that is one-body molded with it and that be positioned at below it; Wherein, described radial attaching parts (4-2-2) divides into outer end and inner end;
The end face of described radial damping layer (4-1) is connected with the bottom surface of described i-th cantilever (1-i), and the bottom surface of described radial damping layer (4-1) is connected with described radial constraint substrate (4-2-1);
Described axial damping unit (5) is arranged along the axial direction of described i-th cantilever (1-i), for providing the freedom of motion of vertical direction, comprise left axial constraint plate (5-1), right axial constraint plate (5-2), left axial damping layer (5-3), right axial damping layer (5-4) and axial shear plate (5-5);
By from one end to the other side direction, end to endly set gradually described left axial constraint plate (5-1), described left axial damping layer (5-3), described axial shear plate (5-5), described right axial damping layer (5-4) and described right axial constraint plate (5-2);
And, the top of described axial shear plate (5-5) is integrally formed is provided with axial connecting part (5-5-1), described axial connecting part (5-5-1) for being fixedly connected with the inner end of described radial attaching parts (4-2-2), and then described axial damping unit (5) is fixedly installed to described radial damping unit (4) below and near center;
The bottom of described left axial constraint plate (5-1) is provided with left axial constraint attaching parts (5-1-1), the bottom of described right axial constraint plate (5-2) is provided with right axial constraint attaching parts (5-2-1); By described left axial constraint attaching parts (5-1-1) and described right axial constraint attaching parts (5-2-1), described axial damping unit (5) is fixed to described base plate (3);
Described elastic support unit (6) for providing the rigidity of vertical direction, and ensures that spacer assembly has certain torsional stiffness and roll stiffness, avoids the normal attitude control torque affecting control moment gyroscope to export; The top of described elastic support unit (6) is fixed to the outer end of described radial attaching parts (4-2-2), and the bottom of described elastic support unit (6) is fixed to described base plate (3);
The upper surface of described base plate (3) is used for being connected with described vibration isolation unit, provides the support of described vibration isolation unit; The lower surface of described base plate (3) is used for being connected and fixed with protected object.
2. the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro according to claim 1, is characterized in that, described adapter plate (1) is spring plate.
3. the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro according to claim 1, is characterized in that, for described radial damping unit (4), the described radial damping layer (4-1) that it comprises is viscoelastic material;
For described axial damping unit (5), its described left axial damping layer (5-3) comprised and described right axial damping layer (5-4) are viscoelastic material.
4. the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro according to claim 3, it is characterized in that, between described radial damping unit (4) and described i-th cantilever (1-i), between described radial damping unit (4) and described radial constraint substrate (4-2-1), between described left axial damping layer (5-3) and described left axial constraint plate (5-1), between described left axial damping layer (5-3) and described axial shear plate (5-5), between described right axial damping layer (5-4) and described axial shear plate (5-5), integrated vulcanizing forming technique is all adopted to carry out between described right axial damping layer (5-4) and described right axial constraint plate (5-2) bonding.
5. the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro according to claim 1, it is characterized in that, described radial attaching parts (4-2-2) divides into outer end and inner end, further, outer connecting bore is offered in the outer end of described radial attaching parts (4-2-2); Interconnecting hole is offered at the inner end of described radial attaching parts (4-2-2);
The described axial connecting part (5-5-1) of described axial damping unit (5) offers axial connecting bore; Described axial connecting bore and described interconnecting hole are bolted fixing;
The top of described elastic support unit (6) is provided with elastic support connecting bore, and described elastic support connecting bore and described outer connecting bore are bolted fixing.
6. the high frequency micro vibration spacer assembly of Spacecraft Control moment gyro according to claim 1, is characterized in that, described elastic support unit (6) is folding arm beam.
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