CN1101751C - Composite honeycomb sandwich structure - Google Patents
Composite honeycomb sandwich structure Download PDFInfo
- Publication number
- CN1101751C CN1101751C CN97192447A CN97192447A CN1101751C CN 1101751 C CN1101751 C CN 1101751C CN 97192447 A CN97192447 A CN 97192447A CN 97192447 A CN97192447 A CN 97192447A CN 1101751 C CN1101751 C CN 1101751C
- Authority
- CN
- China
- Prior art keywords
- core
- layer
- lamination
- resin
- barrier layer
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Lifetime
Links
- 239000002131 composite material Substances 0.000 title claims abstract description 30
- 239000010410 layer Substances 0.000 claims abstract description 165
- 229920005989 resin Polymers 0.000 claims abstract description 76
- 239000011347 resin Substances 0.000 claims abstract description 76
- 230000004888 barrier function Effects 0.000 claims abstract description 49
- 239000012790 adhesive layer Substances 0.000 claims abstract description 39
- 238000000034 method Methods 0.000 claims abstract description 31
- 239000000853 adhesive Substances 0.000 claims abstract description 18
- 230000001070 adhesive effect Effects 0.000 claims abstract description 18
- 238000003475 lamination Methods 0.000 claims description 44
- 238000013467 fragmentation Methods 0.000 claims description 18
- 238000006062 fragmentation reaction Methods 0.000 claims description 18
- 238000010276 construction Methods 0.000 claims description 14
- 241000256844 Apis mellifera Species 0.000 claims description 12
- 239000000835 fiber Substances 0.000 claims description 10
- 239000004744 fabric Substances 0.000 claims description 9
- 239000011159 matrix material Substances 0.000 claims description 7
- 239000004642 Polyimide Substances 0.000 claims description 4
- 229920001721 polyimide Polymers 0.000 claims description 4
- 238000007711 solidification Methods 0.000 claims description 4
- 230000008023 solidification Effects 0.000 claims description 4
- 150000001875 compounds Chemical class 0.000 claims description 3
- 241000256837 Apidae Species 0.000 claims 1
- 208000035484 Cellulite Diseases 0.000 claims 1
- 206010049752 Peau d'orange Diseases 0.000 claims 1
- 238000010438 heat treatment Methods 0.000 abstract description 2
- 239000011162 core material Substances 0.000 description 102
- 238000004519 manufacturing process Methods 0.000 description 9
- 238000013461 design Methods 0.000 description 8
- 238000012545 processing Methods 0.000 description 8
- 238000001723 curing Methods 0.000 description 7
- 239000000463 material Substances 0.000 description 6
- 230000008569 process Effects 0.000 description 6
- 230000006378 damage Effects 0.000 description 4
- 239000011229 interlayer Substances 0.000 description 4
- 229920000049 Carbon (fiber) Polymers 0.000 description 3
- 238000004026 adhesive bonding Methods 0.000 description 3
- 239000011230 binding agent Substances 0.000 description 3
- 239000004917 carbon fiber Substances 0.000 description 3
- 230000008859 change Effects 0.000 description 3
- 238000013007 heat curing Methods 0.000 description 3
- 230000006872 improvement Effects 0.000 description 3
- 238000003825 pressing Methods 0.000 description 3
- 230000003014 reinforcing effect Effects 0.000 description 3
- 239000007787 solid Substances 0.000 description 3
- KAKZBPTYRLMSJV-UHFFFAOYSA-N Butadiene Chemical compound C=CC=C KAKZBPTYRLMSJV-UHFFFAOYSA-N 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 2
- 241000317929 Chalcididae Species 0.000 description 2
- 238000010521 absorption reaction Methods 0.000 description 2
- 239000002390 adhesive tape Substances 0.000 description 2
- 229910052799 carbon Inorganic materials 0.000 description 2
- 238000007598 dipping method Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000008030 elimination Effects 0.000 description 2
- 238000003379 elimination reaction Methods 0.000 description 2
- 238000005516 engineering process Methods 0.000 description 2
- 239000003365 glass fiber Substances 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 229910052751 metal Inorganic materials 0.000 description 2
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 2
- 239000000203 mixture Substances 0.000 description 2
- 239000005011 phenolic resin Substances 0.000 description 2
- 229920001568 phenolic resin Polymers 0.000 description 2
- 238000012360 testing method Methods 0.000 description 2
- 239000002023 wood Substances 0.000 description 2
- XQUPVDVFXZDTLT-UHFFFAOYSA-N 1-[4-[[4-(2,5-dioxopyrrol-1-yl)phenyl]methyl]phenyl]pyrrole-2,5-dione Chemical compound O=C1C=CC(=O)N1C(C=C1)=CC=C1CC1=CC=C(N2C(C=CC2=O)=O)C=C1 XQUPVDVFXZDTLT-UHFFFAOYSA-N 0.000 description 1
- QUTGXAIWZAMYEM-UHFFFAOYSA-N 2-cyclopentyloxyethanamine Chemical compound NCCOC1CCCC1 QUTGXAIWZAMYEM-UHFFFAOYSA-N 0.000 description 1
- 241000446313 Lamella Species 0.000 description 1
- 208000037656 Respiratory Sounds Diseases 0.000 description 1
- 208000034189 Sclerosis Diseases 0.000 description 1
- 239000006096 absorbing agent Substances 0.000 description 1
- 239000004411 aluminium Substances 0.000 description 1
- 229910052782 aluminium Inorganic materials 0.000 description 1
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 description 1
- 230000001413 cellular effect Effects 0.000 description 1
- 239000003795 chemical substances by application Substances 0.000 description 1
- 239000012141 concentrate Substances 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 238000005520 cutting process Methods 0.000 description 1
- 230000001066 destructive effect Effects 0.000 description 1
- 238000001514 detection method Methods 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 230000005518 electrochemistry Effects 0.000 description 1
- 238000001125 extrusion Methods 0.000 description 1
- 239000011521 glass Substances 0.000 description 1
- 239000003292 glue Substances 0.000 description 1
- 238000005470 impregnation Methods 0.000 description 1
- 208000020442 loss of weight Diseases 0.000 description 1
- 150000002739 metals Chemical class 0.000 description 1
- 239000002245 particle Substances 0.000 description 1
- 230000002093 peripheral effect Effects 0.000 description 1
- 229920003192 poly(bis maleimide) Polymers 0.000 description 1
- 229920003223 poly(pyromellitimide-1,4-diphenyl ether) Polymers 0.000 description 1
- 238000005381 potential energy Methods 0.000 description 1
- 230000002265 prevention Effects 0.000 description 1
- 230000000644 propagated effect Effects 0.000 description 1
- WHSIWXIKBLQGIZ-UHFFFAOYSA-N pyrrole-2,5-dione Chemical compound O=C1NC(=O)C=C1.O=C1NC(=O)C=C1.O=C1NC(=O)C=C1 WHSIWXIKBLQGIZ-UHFFFAOYSA-N 0.000 description 1
- -1 reduce cost Substances 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 230000002787 reinforcement Effects 0.000 description 1
- 239000012779 reinforcing material Substances 0.000 description 1
- 230000008439 repair process Effects 0.000 description 1
- 238000012958 reprocessing Methods 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 230000000717 retained effect Effects 0.000 description 1
- 238000002791 soaking Methods 0.000 description 1
- 239000000758 substrate Substances 0.000 description 1
- 239000002344 surface layer Substances 0.000 description 1
- 229920001187 thermosetting polymer Polymers 0.000 description 1
- 239000002699 waste material Substances 0.000 description 1
- 208000016261 weight loss Diseases 0.000 description 1
- 239000013585 weight reducing agent Substances 0.000 description 1
Images
Classifications
-
- G—PHYSICS
- G10—MUSICAL INSTRUMENTS; ACOUSTICS
- G10K—SOUND-PRODUCING DEVICES; METHODS OR DEVICES FOR PROTECTING AGAINST, OR FOR DAMPING, NOISE OR OTHER ACOUSTIC WAVES IN GENERAL; ACOUSTICS NOT OTHERWISE PROVIDED FOR
- G10K11/00—Methods or devices for transmitting, conducting or directing sound in general; Methods or devices for protecting against, or for damping, noise or other acoustic waves in general
- G10K11/16—Methods or devices for protecting against, or for damping, noise or other acoustic waves in general
- G10K11/172—Methods or devices for protecting against, or for damping, noise or other acoustic waves in general using resonance effects
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B32—LAYERED PRODUCTS
- B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B32B3/00—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form
- B32B3/10—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form characterised by a discontinuous layer, i.e. formed of separate pieces of material
- B32B3/12—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form characterised by a discontinuous layer, i.e. formed of separate pieces of material characterised by a layer of regularly- arranged cells, e.g. a honeycomb structure
-
- 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
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Landscapes
- Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- Acoustics & Sound (AREA)
- Multimedia (AREA)
- Laminated Bodies (AREA)
- Moulding By Coating Moulds (AREA)
- Casting Or Compression Moulding Of Plastics Or The Like (AREA)
Abstract
Resin flow into the cells of honeycomb in sandwich structure is eliminated by using an unsupported film adhesive (108), a barrier layer (110), and a scrim supported adhesive layer (112) between the composite laminate (102) and the core (106). The invention produces superior panels with lighter weights, improved mechanical properties, and more predictable structural performance by keeping resin in the laminate rather than losing it to the core cells. Core crush and ply wrinkling in composite honeycomb sandwich structure are reduced by preventing slipping of tiedown plies relative to the mandrel and to one another during autoclave curing. The invention produces superior panels with lighter weights, improved mechanical properties, and more predictable structural performance. The method involves applying a film adhesive to the tiedown plies in the margin of the part outside the net trim line. During heating of the autoclave and prior to the application of high pressure to the composite structure, the film adhesive cures to form a strong bond between the plies and to the mandrel. When pressure is applied, the tiedown plies are locked together and to the mandrel to prevent slippage between any layers in the panel.
Description
Technical field
The present invention relates to composite honeycomb sandwich structure and manufacture method, especially be subjected to the fabric lamella of resin dipping, its formed extexine is bonded on the apparent surface of the honeycomb core with intermediate barrier layers, does not allow resin flow to honeycomb core from the top layer with this.
Background technology
Current space flight honeycomb core sandwich panel (has the composite laminate top layer, stick with glue agent by autoclaving (autoclaveprocessing) and be cured to core) be used widely, because this panel provides high tenacity-weight ratio (that is, than toughness) and strength-to-weight ratio (that is specific strength).Typical honeycomb core sandwich panel is at United States Patent (USP) the 5th, 284, and 702,4,622,091 and 4,353, make description in No. 947.(the 38th international SAMPE seminar in " processing of cellular composite construction and characteristic research " literary composition of people such as Alteneder, 10-13 day in May, 1993 (PCL Internal No.200-1/93-AWA)) common issue with of these panels is discussed, these problems comprise that the nest core damages (being the destruction of core), the mechanical surface fineness of surperficial lamination porosity and difference.
United States Patent (USP) the 5th, 455, No. 861 descriptions are used for the composite lamainated structure of sound absorption (sound is isolated) and other purposes.This laminated construction has seven layers, and they are as follows:
(1) one deck extexine;
The honeycomb of (2) chalcid fly chambers or porous core;
(3) the interior middle interlayer of one deck front side;
(4) medium honeycomb cores in big honeybee chamber;
(5) the interior middle interlayer of one deck rear side;
(6) rear sides, chalcid fly chamber honeycomb or porous core;
(7) one deck endosexine.
Tuning cavity absorber sound absorption in the middle honeycomb core.The performance of this structure, based on the reason of having discussed, it is subjected to the influence of resin flow to the honeybee chamber of honeycomb core man-hour adding, because this flowing will change the resonance characteristics of structure.
Summary of the invention
With a kind of high mobile resin system, resins a large amount of during autoclaving may flow into in-core.This flowing loses resin from lamination, make resin cause the loss of weight in panel, thereby can not obtain required performance, and the design of whole lamination synusia is all needed to consider this flow losses.Resin also can reduce the thickness that solidifies synusia from the loss of lamination synusia, and this reduction is influential to mechanical performance.In order to obtain required performance and corresponding laminated thickness, just need additional synusia, cause the increase of cost and weight.Because the increase of weight directly jeopardizes the performance and the cost of aircraft in present generation aircraft.Because this flowing is a kind of unpredictalbe comparatively speaking and uncontrollable process.So space flight design and manufacturing require to eliminate the mobile of this inflow in-core or few in the extreme.Except resin flowed into the loss in weight of core, we found might be propagated into by the microfissure that the resin that flows causes the outline line of bonding, diminish its mechanical performance.The integrality of the potential energy counter plate of this microfissure has catastrophic threat.Therefore, require to eliminate and thisly to flow or can be controlled at least.
From being laminated to flowing of core, it occurs is because when the treatment temperature that improves, the viscosity of resin descends due to (promptly thinning).So the effort of prior art solution flow field problem concentrates on the viscosity when resin remains on environment temperature when making solidification temperature usually.For example, a kind of way is to change the processing cycle, makes resin solidification in the low pressure steps process that heats up slowly, and resin chain is increased, at the setting up period in this stage, might increase the viscosity that molecular weight keeps resin by the condition of low temperature.The resin of higher molecular weight also has higher viscosity, and institute can stop resin to being harmful to of core to be flowed so that they keep more viscous.Adopt cure cycle stage by stage, more flowing still can occur, and potential microfissure problem exists still.Equally, the hole of surface layer sheet may increase, and exceeds acceptable limit.Also have, a kind of cure cycle of change can increase the time of autoclaving, and processing cost is obviously increased, and has the danger of processing high value parts under complete uncontrolled and insufficient understanding concrete condition.
We utilize high stir-in resin to eliminate in the honeycomb core of resin (matrix) inflow laminated construction, and obtain repeatability and predictability in sandwich panel processing, thereby the structural behaviour of final panel is had confidence.The barrier layer that we adopt intertexture to support, it is arranged between reinforcing fibre resin compounded lamination and the honeycomb core, and this sandwich is low weight than existing panel under identical characteristics.Because resin remains in the lamination (top layer) that structural strength is provided, rather than flow to, can not make the overweight and inefficacy of panel simultaneously the insignificant in-core of structural strength.Usually, we also use the unsupported adhesive layer of the one deck between barrier layer and lamination, so that lamination is bonded on the barrier layer, adopt above-mentioned these layers (they can in conjunction with form a kind of product), can obtain improved performance, resin is retained in the lamination, reduce excessive resin thus, consider when the designer need not design panel that resin flows into in-core yet, and process panel reliably with satisfied structure.
When curing has the panel on the barrier layer (scrim-supported barrier film) that intertexture supports, especially during than the light weight core material, in the slant edge face Qu Jing of honeycomb core regular meeting the core fragmentation appears on probation.To this, we can utilize and be positioned at that core below the barrier layer (and adhesive layer) is contacted fastens the fragmentation that cementing layer (tiedown ply) reduces core.Because when solidifying, fasten cementing layer and can reduce the slip of barrier layer with respect to core.
The present invention controls the slip of core, allows us to utilize the honeycomb core of light density to make cored structure, and the crushing that is unlikely owing to core makes expensive material become waste material.We are by saving time, material and reprocessing or particle utilization, and reduce manufacturing cost by the manufacture process reliability that improves the space flight quality panel of producing toughness with specific intensity.
The additional cementing layer of fastening means and comprises in the final prefabricated panel that three or more fastens cementing layer.In common practical operation, on the outer surface of panel, also may between lamination and gluing barrier layer, exist and fasten cementing layer.The tight cementing layer of every series of strata is from the protruding complete smooth profile that surpasses finished product of parts.Usually, fasten cementing layer and reinforce separately, and often utilize the fabric strip laid type heart.Especially when using the low-density core, that loose cementing layer is fixed to one another, and be fixed to the type heart, this point is very important.The inefficacy of fabric strip can cause in the panel cementing layer wrinkling or core is broken, when the fragmentation of core also appears at fastening cementing layer and being pulled away from the fabric strip that is affixed to the type heart of contact with core accidentally, fastens cementing layer this moment and fastens the cementing layer slip with respect to other.The adhesion strength of individual fibers band is not enough to overcome the power that acts on when being subjected to adding the autoclaving pressurization on the panel inner core.We find how to make and fasten the cementing layer bonding mutually reliably, and are easy and inexpensive.By cementing layer is adhered to one another, power is affacted independent fastening on the layer equably along the whole cementing layer of fastening, reduce the maximum, force that fabric strip is brought when fastening cementing layer and be bonded to core.When composite honeycomb sandwich structure was described, tack coat was used in the various application of fastening cementing layer in the composite construction commonly used comprising.
Thus, the present invention relates to a kind of manufacturing composite construction on the one hand, especially the improvement of composite honeycomb sandwich structure, in sandwich, fasten the parts that are connected when cementing layer is used for carrying out hot-press solidifying under HTHP, fasten cementing layer and be fixed together, make one deck cementing layer not produce motion with respect to another layer for making, we utilize the adhesive that a kind of lower temperature solidifies to be cured, and press the initial period of heat cure which floor adhesive is linked together before pressurization.We are added to the outside of the profile of leveling fully of parts with adhesive, thereby when finished part, can remove adhesive.
On the other hand, the present invention relates to when constituting composite construction,, when high temperature and high pressure carries out autoclaving, will fasten cementing layer and be bonded to another layer especially to honeycomb sandwich construction.Commonly used will fasten cementing layer only to be bonded to the operation of core unsatisfactory, and can stop any one to fasten layer because this stickup must be satisfied, or one deck be with respect to the slippage of another layer.We find that we can be bonded to another layer with fastening cementing layer effectively, utilize only to fasten to add a kind of low-temperature setting adhesive layer between the cementing layer, to reduce to act on the maximum, force on the adhesive tape outside the leveling profile that exposes of parts.In pressing thermal process, the resin temperature of this rete binding agent in being lower than lamination do not melt and solidifies, thereby before the higher temperature of lamination resin flows and curing is increasing the pressure thermal pressure, just makes and fasten cementing layer and combine.The elimination of rete binding agent is fastened cementing layer and is moved with respect to another layer, select for use in the example of two-maleimide (BMI) sandwich panel at us, we select a kind of adhesive for use, it solidifies when temperature is about 250 °F (121 ℃), BMI is solidified at about 375 °F (191 ℃), and regelate is about 440 °F.
Brief description of drawings
Fig. 1 represents typical composite honeycomb sandwich structure.
Fig. 2 be expression have prevention resin that intertexture supports from the top layer partial sectional view to the sandwich mesexine-core interface on the barrier layer that core flows.
Fig. 3 is the phantom of expression prior art honeycomb sandwich construction, utilizes the supported adhesive layer that does not have the barrier layer, and resin flows into in-core.
Fig. 4 is that the expression resin is turned in the top layer, but stops resin to arrive the partial sectional view of the sandwich of core by the barrier layer of the non-support of heaving.
Fig. 5 is that expression is because the sectional drawing of the core fragmentation of core and barrier layer slip causing honeycomb interlayer panel.
Fig. 6 is that another expression utilization is fastened cementing layer to reduce the sectional drawing of core fragmentation.
Fig. 7 is a curve map of typically making the pressure heat cure period of composite honeycomb sandwich panel, is illustrated in the situation of fastening adhesive cures before the cycle of adding high pressure.
Detailed description of the preferred embodiment
As reference system of the present invention, we will introduce typical composite sandwich structural earlier, and then get back to and of the present inventionly will fasten the method that cementing layer bonds together reliably.
A kind of compound honeycomb interlayer panel can reduce, and eliminates or obviously reduces resin from being laminated to flowing of core, only needs a better simply process-cycle, just can be fit to the manufacturing of space structure spare.This sandwich panel 100 (Fig. 1) has outer panels or top layer 102 usually, is glued on the center honeycomb core 106.Top layer 102 comprises the lamination of the organic substrate resin of multilayer reinforcing fibre, and they occur with composite construction form a kind of curing and firm.Core 106 can be a paper wood, synthetic paper wood, and metal, compositions etc. are selected according to being used for.In panel of the present invention; we obtain higher specific strength and than the panel of toughness; fasten adhesive and reduce between core and the surface infringement slip that occurs through regular meeting because we add one deck at least between core 106 and top layer 102, reduce the fragmentation of core when pressing heat cure with this.
In order to stop resin flowing from the composite laminate surface to core, we use the unsupported adhesive layer 108 of one deck (Fig. 2) between top layer 102 and core 106, the adhesive layer 112 that one deck barrier layer 110 and one deck intertexture support flows in the honeycomb cells 114 of core 106 not make resin.
Fig. 3 represents that resin enters the problem of core, and it is the problem that may occur when only not having barrier layer 110 and adhesive layer 108 using adhesive layer 112.The honeybee chamber 114 of honeycomb enters resin 118, and resin 118 moves from lamination, so the resin in the turned letter top layer 102, the turned letter of resin can influence the performance of structure, because it will reduce cementing layer thickness.The resin turned letter will increase gross weight, and this is because the indoor resin 118 of honeybee is discommodity fully.In all situations, uncontrolled resin flows and turned letter can make panel not use.Especially during hot-working, the micro-crack that begins to occur in the honeybee chamber resin 118, and move to the top layer 102 that fiber is strengthened, particularly this crackle appears at the tack line place between top layer 102 and the core 106.
When the adhesive layer that supports without intertexture if use barrier layer 110 attempted to eliminate honeybee chamber resin 118, undesirable heaving as shown in Figure 4 just may be taken place.Heaving of useless resin 120 enters the honeybee chamber 114 of honeybee chamber core 106 to lower convexity, resin is included in heaves in 120, and resins are turned in the top layer 102.Resin flows is heaved in 120 and is influenced structural behaviour, than uncontrollable situation shown in Figure 3, increases the burden of load-bearing.
As shown in Figure 2, have adhesive layer 108, the barrier layer 110 and the supported adhesive layer 112 that interweaves are obstructed resin flows, do not have honeybee chamber resin 118 or resin to heave 120.Yet we find that the barrier layer produces a slide surface between lamination top layer and core, make the core fragmentation through regular meeting during autoclaving.In fact, in 22 of 31 test panels, the fragmentation of experience core.This mortality is unacceptable from cost and conceptual prospect.We adopt tightly is that cementing layer is used to reduce or eliminate the sliding frequency that damages core, because the fragmentation of core is by due to this slip.
For using RIGIDITE
5250-4-W-IW7-GP-CSW, RIGIDITER
5250-4-W-IM7-GP-CSX, and RIGIDITE
Two-superimposed top layer of maleimide (bismaleimide) that 5250-4-W-IM7-GP-PW makes is by Cytec Engineered Materials, and Inc (Cytec) material soaking is handled, and adhesive layer 108 is preferably selected 0.015psf METLBOND for use
The adhesive of 2550U is also provided by Cytec.Adhesive layer provides additional resin, to promote the high-quality combination between lamination and the barrier layer 110.Barrier layer 110 select for use 0.001 inch thick, but have good viscosity, cure cycle can be stood in surface-treated KAPTON polyimides barrier layer, provides impervious rete between top layer-102 and core 106.Intertexture is preferably selected glass fibre for use, " Style 104 " fiber cloth, and adhesive layer 112 is to select 0.06psfMETLBOND for use
The 2550G adhesive can obtain from Cytec.The supported adhesive layer that interweaves stops the barrier layer to blast the core chamber, so resin can be remained in the lamination (being the top layer), cause the adhesive layer thickness maximum of curing, so we obtains best performance under minimum panel weight.
The cementing layer on top layer 102 adopts the prepreg of two-maleimide (bismaleimide) thermosetting resin impregnation of carbon fibers usually, although the present invention also uses other resin system.What can be used to alternative prepreg can be fibre bundle.Adhesive layer 108 should be repaired, so that obtain enough bonding between top layer 102 and the barrier layer 110.Usually, honeycomb core is the phenolic resins honeycomb that the HRP fibrous glass that can be obtained by Hexcel is strengthened.
In sandwich, the adhesive layer of support and barrier layer are metals at core, and as aluminium, the top layer is included in dissimilar material on the electrochemistry, under the situation as carbon fiber, also has erosion-resisting effect.
Other performance of relevant preferred panel proposes in technical article " Development of aBismaleimide/Carbon Honeycomb Sandwich Structure " (Hartz et al.SAMPE, March, 1996), at this as a reference.The improvement on barrier layer introduced in people's articles such as Hartz, fastens the method and the bonding method of the present invention of cementing layer.
Hartz profile plate provides the machinery that is equivalent to solid BMI/ carbon lamination (solidify) when 0.59MPa (85psig) and the marginal zone performance of physics.We be sure of that in our panel, the curing adhesive layer thickness of marginal zone is equivalent to solid lamination at test, and marginal zone 160 (Fig. 5 and 6) satisfies solid lamination non-destructive detection regulation.Marginal zone and panel mechanical performance do not have the intertexture supported adhesive layer, barrier layer, the performance of the sandwich of gluing composite bed above the use that we obtain.The vertical mechanical tension performance of layer of handing over also meets design requirement.
To core carry out preliminary treatment with eliminate to core with about 235 ℃ (455) volatile matters when being heating and curing, this processing is to carry out before the stack sandwich panel, especially is applicable to the phenolic resins core, elimination discharges caused core one lamination unsticking by the gas of core.
The fragmentation 200 (Fig. 5) of core appears at slant edge face district 155, and this situation occurs in when when adding the fusing of thermal pressure and resin slide with respect to panel 102 in barrier layer 110 and core 106, as shown in Figure 5.Barrier layer 110 and core 106 are shifted to the right side, extrusion die in beveled region 155, thus produce core crushing 200.Sinking appears in the place that 160 inner cores move in the marginal zone, top layer 102.
Referring to Fig. 6, improved honeycomb sandwich construction panel is fastened cementing layer 150 along what beveled region 155 comprised that one deck contacts with core 106 at least.This inclined-plane (promptly, often appearing at 160 places, marginal zone) in a corner of in-core appear at usually panel around, but also may appear at the centre position of panel, in composite construction, need reinforcement or the connecting line or the hard spot place of process.
Usually, we adopt in production and have 5 or 8lb/ft
3The carbon fiber of 0/90 fiber orientation commonly used or the individual layer cementing layer 150 of glass fabric are introduced as people such as Hartz in the two-suitable butadiene imidodicarbonic diamide panel of HRP core.The effect of fastening cementing layer 150 is to stop or limits the slip of top layer with respect to core, thereby reduces because the core fragmentation of sliding and causing.Fasten cementing layer 150) during autoclaving, heat, make matrix resin softening, when fusing also is the mobile liquefaction of high attitude resin basically, the fabric of core and certain roughness is fixed.Adopt these panels, we can save 2.5-4lb/ft
3Core because we can use the low-density honeycomb core that does not make core crushing.It is used in fighter plane, can saves weight more than the 25lb every airplane.
As shown in Figure 6, fastening cementing layer is that one deck is narrow, bar on every side, it contacts to core 106 along at least a portion inclined-plane 155, with core 106 have approximately 1 inch overlapping, and reach marginal zone 160 outward, surpass the sideline 165 of parts.Fasten cementing layer 150 or can be positioned at the flat side of beveled region, perhaps be positioned at (see figure 6) on the band angle surface.Key is to fasten cementing layer 150 and adhesive layer to contact with core under the barrier layer 110, and the barrier layer is used for the surface of lamination and core are combined.Cut open at any place of main part rather than in the narrow peripheral region of beveled region and to fasten cementing layer 150, then the circumferential edges at panel forms frame.Thus, fasten cementing layer 150 and can between panel area inner core 106 and surface 102, a gluing interface be arranged.
Traditional practice is, when making Hartz profile plate, we use four complete cover plates to fasten cementing layer 175, so that layer is affixed with core, these cementing layers are shown among Fig. 6.Before introducing Hartz type barrier layer, these traditional cementing layers 175 are used in the production of sandwich panel usually.We also use them always, although we believe that we can need not contact with core and fasten layers all outside the cementing layer 150 except cementing layer fully now on every side.That is, we can use three layers of total cementing layer and need not five layers, as shown in Figure 6.
Fasten cementing layer 150 and 175 and extend outside the profile of leveling fully, paste on each adhesive spots of the reinforcing material type heart to be solidly connected to adhesive tape by marginal zone 160.For further stoping the slip of fastening cementing layer, we add the low adhesive layer 180 that solidifies (being 121 ℃ to the BMI panel promptly) of one deck the fastening of the smooth profile outside of parts between the cementing layer.Adhesive layer 180 can be eliminated the motion of cementing layer with respect to another layer under the situation of autoclaving cycle pressurization.Under the solidification temperature that is lower than the lamination resin, when about 100-150 temperature, solidify.Need increase at us and to fasten cementing layer before pressing thermal pressure and solidify, and the adhesive layer that solidifies makes and fastens cementing layer and be bonded to another layer.Utilize this bonding method to eliminate the relative motion of cementing layer, and the core crushing of eliminating the wrinkling of panel and may occurring.
Originally fasten the method economical with materials, reduce cost, weight reduction because its use " graphical boxes " fasten on every side cementing layer (with traditional compare need not in sheet).Fastening step normally is, (see figure 5) needs cementing layer between the outer surface on top layer and top layer and the adhesive layer divided into.Traditional fastening system need not just will be failed by " graphical boxes " cementing layer, because barrier layer 110 will allow core to slide.Corbett and Smith method need not bonding methods of the present invention, also can fail once in a while.
Core (that is 5-8lb/ft, of the light weight of-maleamide dipping two for adopting
3) and the adhesive system introduced previously, we keep the angle of chamfer to 20 ± 2 °.
Utilize the angle inclined-plane, the cutting area of honeycomb core (inclined-plane) is from the maximum ga(u)ge to the zero thickness, and is tapered with stable gradient.Edge extent in the composite honeycomb sandwich panel is used the inclined-plane, to have the panel structure of inlaying honeycomb and to provide one to seamlessly transit between the marginal zone that lacks any honeycomb core.Method of the present invention allows us to use need not to fasten cementing layer than conventional method and avoid the panel of core fragmentation to have the steeper angle of chamfer.When we selected 20 ° of inclined-planes for use, we believed that may increase angle is suitable for the angle that panel designs requires to any one.
" autoclaving " be meant in the processing procedure and improve the temperature and pressure that is added to panel, with the resin of sclerosis or cured laminate, bonding simultaneously or adhesive curing be laminated to honeycomb core.Fig. 7 represents the cycle that we select for use, and the binding agent that we are used to fasten cementing layer solidifies when about 250 °F (121 ℃), so lamination just solidifies before the pressure thermal pressure that may cause relative motion between the panel internal layer increases.
If the fragmentation of core, the damage of counter plate is normally to the degree that is difficult to repair, so parts are scrapped.Today the advanced person compound resin and the expense of reinforcing fibre will require a kind of method that in fact can eliminate the core fragmentation.Otherwise the expense of processing be can not accept.The panel that uses above-mentioned design is as much as possible near best design, and the emphasis of consideration is not make the core fragmentation.The inventive method can reduce the motion of the fragmentation of core and cementing layer or wrinkling.
We have introduced some examples of selecting for use, but those skilled in the art will easily can expect various changes, changes and improvements, and these changes can be made under the situation that does not depart from the present invention's design.So, can in all known equivalent technologies scopes, explain claims by those of ordinary skills according to specification fully.Provide many examples that the present invention is described, but be not limited to these examples.Thus, be defined by the claims the present invention, if desired, can limit claim by corresponding prior art.
Claims (20)
1. composite honeycomb sandwich structure comprises:
(a) honeycomb core has a plurality of honeybees chamber;
(b) at least one composite laminate has the cementing layer that the fiber that is bonded to core is strengthened matrix resin;
(c) barrier layer, between lamination and core, with bonding lamination and core, and it is indoor from the honeybee that lamination flows into core to eliminate resin;
(d) adhesive layer that between barrier layer and core, has supports fabrics, indoor to eliminate the honeybee that resin flows into or sinking in the barrier layer arrives core.
2. structure as claimed in claim 1, wherein lamination comprises two-maleimide matrix resin.
3. structure as claimed in claim 1, wherein the barrier layer is a kind of cohesible polyimides.
4. structure as claimed in claim 2, wherein cementing layer comprises two-maleimide.
5. structure as claimed in claim 4 also comprises the adhesive layer of non-support between barrier layer and lamination.
6. in making composite construction, will fasten the method that cementing layer bonds together for one kind, comprise the steps:
(a) the compound preform of composite construction shape is placed in lays on the core, preform has at least the lamination of an impregnating resin and at least two to fasten cementing layer;
(b) it is adhered to one another each to be fastened cementing layer, applies cementing layer in the outside of the profile of leveling fully of composite construction.
Wherein solidify under the solidification temperature of adhesive layer resin in being lower than lamination.
7. method as claimed in claim 6, wherein lamination comprises two-maleic acid imide matrix resin.
8. method as claimed in claim 6, wherein preform comprises the barrier layer of being made by the polyimides of the contiguous lamination that reaches the requirement that can bond.
9. method as claimed in claim 7, wherein preform comprises honeycomb core and the adhesive layer between barrier layer and core.
10. method as claimed in claim 9 also comprises in the support interlaced plies that stops between adhesive layer and the core, is absorbed in the core chamber to stop under the barrier layer.
11. method as claimed in claim 9, wherein preform comprises the adhesive layer between barrier layer and the lamination.
12. method as claimed in claim 9, wherein preform comprises the support intertexture layer that stops between adhesive layer and the core, to stop the honeybee that is absorbed in core under the barrier layer indoor.
13. method as claimed in claim 11, wherein preform comprise and adhesive layer and core between the contacted cementing layer of fastening of core.
14. a method that reduces honeycomb core fragmentation in the inclined-plane composite honeycomb sandwich panel, panel has the lamination of the impregnating resin that is glued to honeycomb core, and honeycomb core has the inclined-plane, and this method comprises:
(a) in beveled region, fasten cementing layer and contact, slide to stop between core and the lamination with the honeycomb core of panel;
(b) will fasten the be laid in outer surface of lamination of cementing layer;
(c) will fasten the cementing layer bonding together, and adhere to the laid type heart with being added to the low-temperature setting adhesive of fastening cementing layer in the panels outside of leveling fully,
Wherein, adhesive apply press thermal pressure before, and in lamination the resin fusing and flow before fusing and solidifying.
15. method as claimed in claim 14, wherein lamination comprises and is used to stop resin to flow to the barrier layer of core chamber from the panel of lamination, and has one deck to fasten cementing layer between barrier layer and core.
16. the composite honeycomb sandwich structure with improved anti-cellulite crushing comprises:
(a) honeycomb core has core chamber and inclined-plane on every side;
(b) at least one composite laminate has the cementing layer that the fiber that is glued to core is strengthened matrix resin;
(c) stop adhesive layer, between lamination and core,, and eliminate resin from being laminated to flowing of core chamber with bonding lamination and core;
(d) periphery fasten cementing layer, the inclined-plane of its core below adhesive layer contacts, and to eliminate the slip of barrier layer with respect to core, reduces the fragmentation of core.
17. structure as claimed in claim 16, wherein lamination comprises two-maleimide matrix resin.
18. structure as claimed in claim 16, wherein the barrier layer is cohesible polyimides.
19. structure as claimed in claim 17, wherein adhesive comprises two-maleimide.
20. structure as claimed in claim 19 also comprises the adhesive layer between barrier layer and lamination.
Applications Claiming Priority (6)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US08/587,160 US5604010A (en) | 1996-01-11 | 1996-01-11 | Composite honeycomb sandwich structure |
| US08/587,160 | 1996-01-11 | ||
| US08/616,903 | 1996-03-15 | ||
| US08/616,903 US5895699A (en) | 1996-03-15 | 1996-03-15 | Tiedown ply for reducing core crush in composite honeycomb sandwich structure |
| US08/620,829 | 1996-03-20 | ||
| US08/620,829 US5685940A (en) | 1996-03-20 | 1996-03-20 | Adhering tiedown plies in composite construction |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN1211947A CN1211947A (en) | 1999-03-24 |
| CN1101751C true CN1101751C (en) | 2003-02-19 |
Family
ID=27416496
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN97192447A Expired - Lifetime CN1101751C (en) | 1996-01-11 | 1997-01-06 | Composite honeycomb sandwich structure |
Country Status (5)
| Country | Link |
|---|---|
| EP (1) | EP0883484A1 (en) |
| JP (2) | JP3913275B2 (en) |
| CN (1) | CN1101751C (en) |
| AU (1) | AU1822997A (en) |
| WO (1) | WO1997025198A1 (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104703781A (en) * | 2012-10-04 | 2015-06-10 | 波音公司 | Composite structure having a stabilizing element |
| CN104973232A (en) * | 2014-02-04 | 2015-10-14 | 波音公司 | Bonded tab and machining device |
| CN109159518A (en) * | 2018-08-28 | 2019-01-08 | 丹阳丹金航空材料科技有限公司 | A kind of aviation aircraft composite plate |
Families Citing this family (24)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE69917264T2 (en) | 1998-05-22 | 2005-05-25 | Cytec Technology Corp., Wilmington | PRODUCTS AND METHOD FOR PREVENTING THE COMPRESSION PRESSURE OF A CORE |
| JP4663174B2 (en) * | 2001-08-03 | 2011-03-30 | 富士重工業株式会社 | Method for forming honeycomb sandwich structure composite material |
| ATE406998T1 (en) | 2003-07-08 | 2008-09-15 | Airbus Gmbh | LIGHTWEIGHT STRUCTURE |
| EP1495858B1 (en) | 2003-07-08 | 2019-08-07 | Airbus Operations GmbH | Lightweight material structure made of metal composite material |
| JP2007098819A (en) * | 2005-10-06 | 2007-04-19 | Mitsubishi Rayon Co Ltd | Manufacturing method of sandwich panel |
| FR2908737B1 (en) * | 2006-11-16 | 2009-12-04 | Airbus France | ACOUSTIC COATING FOR AIRCRAFT INCORPORATING A JELLY EFFECT FROST TREATMENT SYSTEM. |
| US8046915B2 (en) * | 2007-12-12 | 2011-11-01 | General Electric Company | Methods for making composite containment casings |
| US8343298B2 (en) * | 2007-12-13 | 2013-01-01 | The Boeing Company | Aircraft structures bonded with adhesive including magnetostrictive material |
| US20090155524A1 (en) * | 2007-12-13 | 2009-06-18 | Rapp Robert A | Composite panel and method of manufacturing the same |
| BE1017910A3 (en) | 2007-12-21 | 2009-11-03 | Sonaca Sociutu Anonyme | METHOD FOR MANUFACTURING A PANEL COMPRISING AT LEAST ONE NANNED BODY AND A FIRST SKIN REALIZED IN COMPOSITE MATERIAL |
| US8491743B2 (en) * | 2009-12-15 | 2013-07-23 | The Boeing Company | Composite ply stabilizing method |
| JP5619485B2 (en) * | 2010-06-22 | 2014-11-05 | 豊和繊維工業株式会社 | Laminated vehicle interior substrate and manufacturing method thereof |
| US8844873B2 (en) | 2011-09-23 | 2014-09-30 | The Boeing Company | Stabilizer torque box assembly and method |
| FR2987307B1 (en) * | 2012-02-29 | 2017-02-10 | Daher Aerospace | METHOD AND DEVICE FOR THE COMPACTION AND CONSOLIDATION OF A HIGH THERMOPLASTIC DIE COMPOSITE PANEL |
| CN102700181B (en) * | 2012-05-15 | 2014-09-03 | 西安交通大学 | Light multifunctional composite structure |
| CN102909899B (en) * | 2012-10-25 | 2015-03-11 | 溧阳二十八所系统装备有限公司 | Production method for waterproof paper honeycomb compound sandwich board |
| CN103161391A (en) * | 2013-03-07 | 2013-06-19 | 苏州市江诚人防设备有限公司 | Inorganic protective door |
| WO2015013012A1 (en) | 2013-07-26 | 2015-01-29 | Learjet Inc. | Composite material incorporating water ingress barrier |
| JP5992946B2 (en) | 2014-03-25 | 2016-09-14 | 富士重工業株式会社 | Honeycomb structure and method for manufacturing honeycomb structure |
| CN105539809A (en) * | 2014-10-28 | 2016-05-04 | 哈尔滨飞机工业集团有限责任公司 | Sectional honeycomb sandwich structure |
| EA201991043A1 (en) | 2016-10-27 | 2019-09-30 | Руаг Швайц Аг | PRODUCTION OF FIBER REINFORCED POLYMER |
| JP7039401B2 (en) | 2018-06-28 | 2022-03-22 | 三菱重工業株式会社 | Composite material and method of curing composite material |
| JP7114367B2 (en) | 2018-06-28 | 2022-08-08 | 三菱重工業株式会社 | Forming method for composite structure |
| CN111186147B (en) * | 2020-02-24 | 2021-10-08 | 江苏亨睿碳纤维科技有限公司 | Forming method for preparing lightweight automobile parts by using continuous fiber mixed chopped fibers |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1391558A (en) * | 1972-03-23 | 1975-04-23 | Boeing Co | Honeycomb structural panels |
| US4954382A (en) * | 1988-11-01 | 1990-09-04 | American Cyanamid Company | Interleaf layer in fiber reinforced resin laminate composites |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2198835B1 (en) * | 1972-09-11 | 1975-03-07 | Rhone Poulenc Ind | |
| CA1239572A (en) * | 1983-09-21 | 1988-07-26 | Roger A. Stonier | Method for cocuring a composite skin directly to honeycomb core |
| US4598007A (en) * | 1985-02-28 | 1986-07-01 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Light weight fire resistant graphite composites |
| GB8817669D0 (en) * | 1988-07-25 | 1988-09-01 | Short Brothers Ltd | Means for attenuating sound energy |
| FR2726500B1 (en) * | 1994-11-09 | 1997-01-10 | Eurocopter France | PROCESS FOR PRODUCING COMPOSITE SANDWICH PANELS AND PANELS OBTAINED THEREBY |
-
1997
- 1997-01-06 AU AU18229/97A patent/AU1822997A/en not_active Abandoned
- 1997-01-06 WO PCT/US1997/000075 patent/WO1997025198A1/en active Application Filing
- 1997-01-06 JP JP52529497A patent/JP3913275B2/en not_active Expired - Lifetime
- 1997-01-06 CN CN97192447A patent/CN1101751C/en not_active Expired - Lifetime
- 1997-01-06 EP EP97903737A patent/EP0883484A1/en not_active Withdrawn
-
2006
- 2006-07-24 JP JP2006200715A patent/JP4407964B2/en not_active Expired - Lifetime
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1391558A (en) * | 1972-03-23 | 1975-04-23 | Boeing Co | Honeycomb structural panels |
| US4954382A (en) * | 1988-11-01 | 1990-09-04 | American Cyanamid Company | Interleaf layer in fiber reinforced resin laminate composites |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104703781A (en) * | 2012-10-04 | 2015-06-10 | 波音公司 | Composite structure having a stabilizing element |
| CN104973232A (en) * | 2014-02-04 | 2015-10-14 | 波音公司 | Bonded tab and machining device |
| US10377122B2 (en) | 2014-02-04 | 2019-08-13 | The Boeing Company | Bonded tab and tooling device |
| US11007766B2 (en) | 2014-02-04 | 2021-05-18 | The Boeing Company | Bonded tab and tooling device |
| CN109159518A (en) * | 2018-08-28 | 2019-01-08 | 丹阳丹金航空材料科技有限公司 | A kind of aviation aircraft composite plate |
Also Published As
| Publication number | Publication date |
|---|---|
| JP2000502968A (en) | 2000-03-14 |
| JP3913275B2 (en) | 2007-05-09 |
| WO1997025198A1 (en) | 1997-07-17 |
| JP2007015385A (en) | 2007-01-25 |
| JP4407964B2 (en) | 2010-02-03 |
| EP0883484A1 (en) | 1998-12-16 |
| AU1822997A (en) | 1997-08-01 |
| CN1211947A (en) | 1999-03-24 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN1101751C (en) | Composite honeycomb sandwich structure | |
| US6698484B1 (en) | Method for reducing core crush | |
| US6180206B1 (en) | Composite honeycomb sandwich panel for fixed leading edges | |
| US5604010A (en) | Composite honeycomb sandwich structure | |
| US4680216A (en) | Method for stabilizing thick honeycomb core composite articles | |
| US5518796A (en) | Near-surface enhancement of honeycomb sandwich structures to improve durability using a foaming fiber-filled adhesive | |
| JP5638654B2 (en) | Hybrid laminate and aircraft fuselage part | |
| WO1997025198A9 (en) | Composite honeycomb sandwich structure | |
| US4249976A (en) | Manufacture of honeycomb sandwich | |
| US6743742B1 (en) | Method of partially embedding non-woven fiber mat to reinforcing fibers impregnated with a thermoplastic resin and composition therefrom | |
| US5685940A (en) | Adhering tiedown plies in composite construction | |
| US20040200571A1 (en) | Methods of forming honeycomb sandwich composite panels | |
| US8926786B1 (en) | Composite panel and method of manufacturing the same | |
| RU2007115401A (en) | THIN LAYERED LAMINATES | |
| AU2005232257A1 (en) | Moulding materials | |
| JPH08207185A (en) | Multilayered laminate and production of multilayered laminate panel | |
| US20100043955A1 (en) | Flat-Cured Composite Structure | |
| US7138031B2 (en) | Mandrel and method for manufacturing composite structures | |
| WO2022041358A1 (en) | Novel composite panel and production method therefor | |
| JPS6092831A (en) | Method of co-curing skin member made of composite material and directly joining it to honeycomb core | |
| US6780488B2 (en) | Liquid molded hollow cell core composite articles | |
| CA2473350C (en) | Lightweight structure especially for an aircraft and method for making such a structure | |
| CN108973089B (en) | Preparation method of composite material special-shaped skin structure | |
| CA2242050C (en) | Composite honeycomb sandwich structure | |
| JPH05185542A (en) | Structure with multilayer honeycomb core and manufacture thereof |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| CX01 | Expiry of patent term |
Granted publication date: 20030219 |
|
| EXPY | Termination of patent right or utility model |