CN101567025A - Finite element modeling method used for damage process of thermal barrier coating of turbine blade - Google Patents
Finite element modeling method used for damage process of thermal barrier coating of turbine blade Download PDFInfo
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Abstract
The invention discloses a finite element modeling method used for damage processes of the thermal barrier coating of a turbine blade, belonging to the field of heat-insulation protection coating system technology of high-performance aero-engine. The method adopts a CATIA software and a finite element ABAQUS software. The method wholly models all coatings and the substrate in the CATIA software, carries out a secondary processing to geometrical models in the ABAQUS software, carries out geometrical cleaning to the model in the CATIA software, and carries out simulation, calculation and result analysis in the ABAQUS software. The method can effectively overcome the incompatibility of the CAD software such as CATIA and the like and the finite element software, considers both various properties of the heat barrier coating material and the geometrical shape of the substrate, can exactly simulate the relevant key physical quantity of the turbine blade heat barrier coating system in practical working environment, has clear concept and is convenient to be mastered and used by the technicians.
Description
Technical field
The invention belongs to the heat insulation protective coating systems technology of high-performance aeromotor field, particularly be used for the finite element modeling method of thermal barrier coating of turbine blade system destruction process.
Background technology
Thermal barrier coating (thermal barrier coatings, abbreviation TBCs) being based on stupalith has the fusing point height, pyroconductivity is low, steam forces down, characteristics such as the low and reflectivity height of radiance, ceramic powders is sprayed or is deposited on high temperature alloy hot-end component (especially turbo blade) surface, to reduce the working temperature of high-temperature component, make it avoid high temperature corrosion and high-temperature oxydation, prolong the serviceable life of high-temperature component greatly, the interior high-temperature alloy part of modern aero gas turbine engine is worked in being higher than the Service Environment of its melting temperature become possibility, and then improved the aeromotor fuel gas temperature and the thermal efficiency, thereby be widely used in Aero-Space, chemical industry, fields such as the metallurgy and the energy.
Yet in actual applications, because factors such as the layers of material parameter do not match, high-temerature creep, high-temperature interface oxidation and stupalith high-temperature phase-change cause thermal barrier coating to be subjected to the alternating action of thermal stress and compressive residual stress jointly, and along with the increase of using the time, be subjected to increasing compressive residual stress effect in the ceramic layer, the continuous nucleation, expansion and the crackle that also are accompanied by coating interface hole or Interface Crack simultaneously are connected.Along with the increase of times of thermal cycle, compressive residual stress of Zeng Daing and Interface Crack have caused ceramic coat with flexing with peel off form and metallic matrix and be separated and destroy jointly gradually.In case coating is peeled off, the hot side metal parts will be directly exposed under the high-temperature severe environment, and its consequence is very serious.Therefore domestic and international many researchists adopt multiple means such as theoretical analysis, experimental study and analogy method to study destructive process and the failure mechanism between the ceramic coat and metallic substrates under specific work environments, thereby the mission life or the active time of prediction thermal barrier coating system improve its reliability of applying.But for the thermal barrier coating system (for example turbo blade, guide vane) of labyrinth, the general theoretical analytic solution of employing that are difficult to are carried out correlative study, must rely on experiment test and limited analogy method to realize.Wherein experiment test repeatedly need spend great amount of manpower and material resources and financial resources.Therefore the finite element analogy method will become a kind of research method of main flow, not only can reduce experimentation cost, and can reduce design and lead time.
In the thermal barrier coating field, carried out Study of finite element simulation more widely at present, temperature field, displacement field, stress field and the destructive process of prediction thermal barrier coating system under certain working environment.But most of research work also mainly concentrates on fundamental research, and its research object mainly is the thermal barrier coating system (parts) of simple geometric configuration (flat plate model, semicircle model etc.).And it is very limited for self Geometric Modeling ability of common finite element software program, be difficult to set up the object model of complex geometry, therefore and its workload is very big, and complex operation, presses for that the relevant finite element of development is handled means special early stage and skill realizes.
Through new, there be not the report of discovery to the Study of finite element simulation of the thermal barrier coating of turbine blade system destruction process of complex-curved true form to looking into of prior art.
Summary of the invention
The purpose of this invention is to provide the finite element modeling method that is used for thermal barrier coating of turbine blade system destruction process, it is characterized in that, may further comprise the steps:
(1) foundation of geometric model in earlier stage
In the model, thermal barrier coating represents that with TBC thickness is h
cOxide layer represents that with TGO thickness is h
tTack coat represents that with BC thickness is h
bThe turbo blade substrate of turbo blade substrate or band base represents that with SUB thickness is h
sThe turbo blade substrate base thickness of band base is h
d
1. set up the TBC-TGO-BC-SUB-Cuboid model in CATIA software, be designated as model A, model A represents to apply the Turbine Blade Model of tack coat, oxide layer and thermal barrier coating, saves as the TBC-TGO-BC-SUB-Cuboid-A.stp file;
2. set up the TGO-BC-SUB-Cuboid model in CATIA software, be designated as Model B, Model B represents to apply the Turbine Blade Model of tack coat and oxide layer, saves as the TGO-BC-SUB-Cuboid-B.stp file;
3. set up the BC-SUB-Cuboid model in CATIA software, be designated as MODEL C, MODEL C represents to apply the Turbine Blade Model of tack coat, saves as the BC-SUB-Cuboid-C.stp file;
4. set up the SUB-Cuboid model in CATIA software, be designated as model D, model D represents turbo blade substrate model, does not promptly apply the turbo blade of any coating, saves as the SUB-Cuboid-D.stp file;
5. in CATIA software, to the turbo blade of band base, chamfering is carried out in blade and base connecting portion and each layer coating junction operate, each chamfer radius is identical, and size is r;
(2) secondary treating of geometric model
The data file of middle A, the B that sets up of step (1), C, four models of D is imported in the ASEEMBLY module of large commercial finite element software ABAQUS software:
1. model A and Model B is tangent, generate the TBC layer, the model file of TBC layer is saved as TBC.stp;
2. Model B and MODEL C is tangent, generate the TGO layer, the model file of TGO layer is saved as TGO.stp;
3. MODEL C and model D is tangent, generate the BC layer, the model file of BC layer is saved as BC.stp;
(3) how much cleanings of each layer model
With TBC layer, TGO layer and the BC layer three layer model data importing CATIA software that generates in the step (2), carry out how much cleaning works of model, leave out unnecessary part, obtain being used for the model of finite element analysis, will clear up good model then and import ABAQUS software respectively;
(4) the final generation of thermal barrier coating of turbine blade system model
Do not consider destruction, the slippage at interface, set the intact combination in interface, in the ASEEMBLY module in ABAQUS software four layer models of setting up are previously merged, promptly merge TBC.stp, TGO.stp, BC.stp and four pairing models of file of SUB-Cuboid.stp, obtain whole thermal barrier coating system model, be designated as TBCs;
(5) respectively to TBC layer, TGO layer, BC layer, blade base or blade base and base definition material parameter;
(6) mechanical boundary condition and calorifics boundary condition are set;
(7) divide grid;
(8) preserve model data, adopt latent trial and error procedure, calculate;
(9) carry out sign as a result, find out dangerous point and hazardous location.
Described h
cBe 0.15～1.60mm, described h
tBe 0.01～0.10mm, described h
bBe 0.05～0.20mm, described h
sBe 1.00～2.00mm, described h
dBe 10.00～100.00mm, described r is 0.50～2.00mm.
Beneficial effect of the present invention is: this method can effectively overcome the incompatibility on geometric model is handled between CAD software such as CATIA and the finite element software, can consider the various attributes of heat barrier coat material, can consider the geometric configuration of substrate again, can simulate the relevant key physical amount (as temperature field, stress field, dangerous point and hazardous location or the like) of thermal barrier coating of turbine blade system in actual working environment exactly, clear thinking makes things convenient for the technician to grasp and uses.This method equally also is applicable to the finite element modeling and the emulation of the laminated coating/membrane system of other type structure complexity, helps the research and development design and processes optimization of films/coatings product.
Description of drawings
Fig. 1 is the process flow diagram that is used for the finite element modeling method of thermal barrier coating of turbine blade system destruction process;
Fig. 2 is the block mold of single thermal barrier coating of turbine blade system;
Fig. 3 is the block mold that contains the thermal barrier coating of turbine blade system of base;
Fig. 4 be thermal barrier coating system bear the thermal cycle experiment mode;
Fig. 5 is the distribution of dangerous point and hazardous location in each layer of single thermal barrier coating of turbine blade system;
Fig. 6 is the distribution that contains dangerous point and hazardous location in each layer of thermal barrier coating of turbine blade system of base;
Number in the figure: 1-thermal barrier coating; The 2-oxide layer; The 3-tack coat; The 4-substrate; The 5-base.
Embodiment
The invention will be further described below in conjunction with accompanying drawing:
Embodiment 1
Choose single thermal barrier coating of turbine blade system (no base) as objective for implementation, shown in Fig. 2 (a), set up its finite element analysis model, the finite element simulation of Xingqi under the thermal cycle effect of going forward side by side.
For convenience, for simplicity, the present invention makes the following assumptions: 1) the even and equal isotropy of each layer coating, substrate and submount material; 2) each layer coating thickness is even; 3) adopt ideal elastoplastic model, and only consider the high-temerature creep of TBC layer; 4) this method has only been described the finite element simulation of single thermal barrier coating of turbine blade system.
The finite element modeling method that is used for thermal barrier coating of turbine blade system destruction process, this method flow diagram may further comprise the steps as shown in Figure 1:
(1) foundation of geometric model in earlier stage
In the model, thermal barrier coating represents that with TBC thickness is h
cOxide layer represents that with TGO thickness is h
tTack coat represents that with BC thickness is h
bThe turbo blade substrate of turbo blade substrate or band base represents that with SUB thickness is h
sWherein, h
c=0.40mm, h
t=0.10mm, h
b=0.20mm, h
s=1.80mm; Shown in Fig. 2 (b), turbo blade is followed successively by thermal barrier coating 1, oxide layer 2, tack coat 3, substrate 4 from outside to inside.
1. set up the TBC-TGO-BC-SUB-Cuboid model in CATIA software, be designated as model A, model A represents to apply the Turbine Blade Model of tack coat, oxide layer and thermal barrier coating, saves as the TBC-TGO-BC-SUB-Cuboid-A.stp file;
2. set up the TGO-BC-SUB-Cuboid model in CATIA software, be designated as Model B, Model B represents to apply the Turbine Blade Model of tack coat and oxide layer, saves as the TGO-BC-SUB-Cuboid-B.stp file;
3. set up the BC-SUB-Cuboid model in CATIA software, be designated as MODEL C, MODEL C represents to apply the Turbine Blade Model of tack coat, saves as the BC-SUB-Cuboid-C.stp file;
4. set up the SUB-Cuboid model in CATIA software, be designated as model D, model D represents turbo blade substrate model, does not promptly apply the turbo blade of any coating, saves as the SUB-Cuboid-D.stp file;
(2) secondary treating of geometric model
The data file of middle A, the B that sets up of step (1), C, four models of D is imported in the ASEEMBLY module of large commercial finite element software ABAQUS software:
1. model A and Model B is tangent, generate the TBC layer, the model file of TBC layer is saved as TBC.stp;
2. Model B and MODEL C is tangent, generate the TGO layer, the model file of TGO layer is saved as TGO.stp;
3. MODEL C and model D is tangent, generate the BC layer, the model file of BC layer is saved as BC.stp;
(3) how much cleanings of each layer model
With TBC layer, TGO layer and the BC layer three layer model data importing CATIA software that generates in the step (2), carry out how much cleaning works of model, leave out unnecessary part, obtain being used for the model of finite element analysis, will clear up good model then and import ABAQUS software respectively;
(4) the final generation of thermal barrier coating of turbine blade system model
Do not consider the destruction, slippage at interface etc., set the intact combination in interface, in the ASEEMBLY module in ABAQUS software four layer models of setting up are previously merged, promptly merge TBC.stp, TGO.stp, BC.stp and four pairing models of file of SUB-Cuboid.stp, obtain whole thermal barrier coating system model, be designated as TBCs, as shown in Figure 2;
(5) respectively to TBC layer, TGO layer, BC layer, blade base definition material parameter, the critical material parameter of thermal barrier coating system layers of material is all considered to vary with temperature and is changed, and concrete numerical value is shown in table 1～7;
The elastic modulus of table 1 layers of material
The Poisson ratio of table 2 layers of material
The thermal expansivity of table 3 layers of material
The yield strength of table 4 layers of material
The heat-conduction coefficient of table 5 layers of material
The specific heat of table 6 layers of material
The creep parameters of table 7 ceramic layer
(6) mechanical boundary condition and calorifics boundary condition are set, the mechanical boundary condition is the point of anchoring base lower surface, shown in Fig. 2 (a), the calorifics boundary condition is the temperature that TBC layer outside surface and cooling duct are set according to thermal cycle, the thermal cycle that thermal barrier coating system stood mainly comprises 2 processes as shown in Figure 4:
1. after TBCs systems produce operation was finished, total system was cooled to 20 ℃ from 400 ℃;
2. the thermal cycle mode of She Dinging is: the ceramic coat surface is warming up to 1121 ℃ in 10 minutes, and turbo blade (substrate) inner cooling channel is warming up to 700 ℃; Keeping this thermograde duration then is 40 minutes; Total system all is cooled to 20 ℃ in 10 minutes again;
(7) divide grid;
(8) preserve model data, adopt latent trial and error procedure, calculate;
(9) carry out sign as a result, find out dangerous point and hazardous location;
1. for TBC layer and TGO layer, they belong to hard brittle material, we adopt first strength theory (being the maximum tension stress failure criteria), TBC layer and TGO layer are extracted respectively, observe the distribution of major principal stress (Max.Principal stress), it is exactly the place of losing efficacy at first that major principal stress is concentrated part, i.e. dangerous point, and the zone that stress ratio is bigger is the hazardous location;
2. for BC layer, blade base, they belong to toughness material, we adopt the 3rd intensity theory (being the Mises yield criteria), BC layer, blade base are extracted respectively, observe the distribution of Mises equivalent stress and plastic strain, the place of Mises equivalent stress maximum or plastic strain maximum is exactly the dangerous point of respective layer, and the Mises stress ratio is the hazardous location than big or the bigger zone of plastic strain.
By observation and the analysis to result of calculation, the stress level of thermal barrier coating system is the highest at the terminal point of cooling stage constantly, and we should pay close attention to the dangerous point of thermal barrier coating of turbine blade system in cooling procedure and the distribution of hazardous location.When now extracting the 10th thermal cycle end, the distribution of dangerous point and hazardous location as shown in Figure 5 in each layer of thermal barrier coating system.We can find: in TBC layer (shown in Fig. 5 (a)) and TGO layer (shown in Fig. 5 (b)), leaf basin central region is the hazardous location that stress is concentrated, and the stress value in the TBC layer hazardous location is about 170.70～184.30MPa; Stress value in the TGO layer hazardous location is about 1351.00MPa～1472MPa; In BC layer (shown in Fig. 5 (c)), the leading edge of blade and the stress value of trailing edge are all than higher, and the scope of its plastic strain value is 0.76%～0.81%, but dangerous point appears at the position, middle and upper part of trailing edge, and the plastic strain at this place is 0.81%; In substrate (shown in Fig. 5 (d)), the leaf basin zone of leading edge and close trailing edge is a region of stress concentration, and the hazardous location stress value is at 537.20～581.80MPa.Destroy and at first can occur in these hazardous locations, the stress level in the zone outside the hazardous location is all lower, is the safety zone relatively, generally can not destroy.
Choose contain base the thermal barrier coating of turbine blade system as objective for implementation, shown in Fig. 3 (a), set up its finite element analysis model, the finite element simulation of Xingqi under the thermal cycle effect of going forward side by side.
For convenience, for simplicity, the present invention makes the following assumptions: 1) the even and equal isotropy of each layer coating, substrate and submount material; 2) each layer coating thickness is even; 3) base is reduced to rectangular parallelepiped; 4) adopt ideal elastoplastic model, and only consider the high-temerature creep of TBC layer; 5) this method has only been described the finite element simulation of single thermal barrier coating of turbine blade system.
Be used for the finite element modeling method of thermal barrier coating of turbine blade system destruction process, may further comprise the steps:
(1) foundation of geometric model in earlier stage
In the model, thermal barrier coating represents that with TBC thickness is h
cOxide layer represents that with TGO thickness is h
tTack coat represents that with BC thickness is h
bThe turbo blade substrate of turbo blade substrate or band base represents that with SUB thickness is h
sThe turbo blade substrate base thickness of band base is h
d, wherein, h
c=0.40mm, h
t=0.10mm, h
b=0.20mm, h
s=1.80mm, h
d=5.00mm; Shown in Fig. 3 (b), turbo blade is followed successively by thermal barrier coating 1, oxide layer 2, tack coat 3, substrate 4 from outside to inside, and turbo blade contains base 5 (shown in Fig. 3 (c)).
1. set up the TBC-TGO-BC-SUB-Cuboid model in CATIA software, be designated as model A, model A represents to apply the Turbine Blade Model of tack coat, oxide layer and thermal barrier coating, saves as the TBC-TGO-BC-SUB-Cuboid-A.stp file;
2. set up the TGO-BC-SUB-Cuboid model in CATIA software, be designated as Model B, Model B represents to apply the Turbine Blade Model of tack coat and oxide layer, saves as the TGO-BC-SUB-Cuboid-B.stp file;
3. set up the BC-SUB-Cuboid model in CATIA software, be designated as MODEL C, MODEL C represents to apply the Turbine Blade Model of tack coat, saves as the BC-SUB-Cuboid-C.stp file;
4. set up the SUB-Cuboid model in CATIA software, be designated as model D, model D represents turbo blade substrate model, does not promptly apply the turbo blade of any coating, saves as the SUB-Cuboid-D.stp file;
5. in CATIA software, blade and base connecting portion and each layer coating junction are carried out the chamfering operation, each chamfer radius is identical, and size is r, r=1.00mm;
(2) secondary treating of geometric model
The data file of middle A, the B that sets up of step (1), C, four models of D is imported in the ASEEMBLY module of large commercial finite element software ABAQUS software:
1. model A and Model B is tangent, generate the TBC layer, the model file of TBC layer is saved as TBC.stp;
2. Model B and MODEL C is tangent, generate the TGO layer, the model file of TGO layer is saved as TGO.stp;
3. MODEL C and model D is tangent, generate the BC layer, the model file of BC layer is saved as BC.stp;
(3) how much cleanings of each layer model
With TBC layer, TGO layer and the BC layer three layer model data importing CATIA software that generates in the step (2), carry out how much cleaning works of model, leave out unnecessary part, obtain being used for the model of finite element analysis, will clear up good model then and import ABAQUS software respectively;
(4) the final generation of thermal barrier coating of turbine blade system model
Do not consider the destruction, slippage at interface etc., set the intact combination in interface, in the ASEEMBLY module in ABAQUS software four layer models of setting up are previously merged, promptly merge TBC.stp, TGO.stp, BC.stp and four pairing models of file of SUB-Cuboid.stp, obtain whole thermal barrier coating system model, be designated as TBCs, as shown in Figure 3;
(5) respectively to TBC layer, TGO layer, BC layer, blade base and base definition material parameter, thermal barrier coating system layers of material parameter is provided with identical with embodiment 1, and referring to table 1-7, the substrate submount material is identical with base material, and is identical with base property;
(6) mechanical boundary condition and calorifics boundary condition are set, the mechanical boundary condition is a bit of firm banking lower surface, shown in Fig. 3 (a), the calorifics boundary condition is the temperature that TBC layer outside surface and cooling duct are set according to thermal cycle, the thermal cycle mode of thermal barrier coating system experience is identical with embodiment 1, as shown in Figure 4;
(7) divide grid;
(8) preserve model data, adopt latent trial and error procedure, calculate;
(9) carry out sign as a result, find out dangerous point and hazardous location;
1. for TBC layer and TGO layer, they belong to hard brittle material, we adopt first strength theory (being the maximum tension stress failure criteria), TBC layer and TGO layer are extracted respectively, observe the distribution of major principal stress (Max.Principal stress), it is exactly the place of losing efficacy at first that major principal stress is concentrated part, i.e. dangerous point, and the zone that stress ratio is bigger is the hazardous location;
2. for BC layer, blade base and base, they belong to toughness material, we adopt the 3rd intensity theory (being the Mises yield criteria), BC layer, blade base and base are extracted respectively, observe the distribution of Mises equivalent stress and plastic strain, the place of Mises equivalent stress maximum or plastic strain maximum is exactly the dangerous point of respective layer, and the Mises stress ratio is the hazardous location than big or the bigger zone of plastic strain.
The Stress Field Distribution of thermal barrier coating system after the 10th loop ends that contains base as shown in Figure 6.By analyzing, we can find: considered after the base of blade, very big variation has taken place in the stress distribution situation, in TBC layer (shown in Fig. 6 (a)) and TGO layer (shown in Fig. 6 (b)), dangerous point and hazardous location appear at the junction of leaf basin and base, TBC layer explosive area stress value is about 187.00～225.00MPa, and TGO layer explosive area stress value is about 2173.00～2494.00MPa; In BC layer (shown in Fig. 6 (c)), the leaf basin top of close trailing edge is dangerous point (dangerous point place stress value is 426.00MPa), and the top of leading edge also is the hazardous location, and the plastic strain in explosive area is about 3.39%～3.69%; In substrate (shown in Fig. 6 (d)), the junction of leading edge and trailing edge and base is the hazardous location, and substrate explosive area stress value is about 735.3～800.00MPa, and wherein, dangerous point appears at the junction of trailing edge and base, and stress value is 800MPa.Destroy and at first can occur in these hazardous locations, the stress level in the zone outside the hazardous location is all lower, is the safety zone relatively, generally can not destroy.
Claims (2)
1, be used for the finite element modeling method of thermal barrier coating of turbine blade system destruction process, it is characterized in that, may further comprise the steps:
(1) foundation of geometric model in earlier stage
In the model, thermal barrier coating represents that with TBC thickness is h
cOxide layer represents that with TGO thickness is h
tTack coat represents that with BC thickness is h
bThe turbo blade substrate of turbo blade substrate or band base represents that with SUB thickness is h
sThe turbo blade substrate base thickness of band base is h
d
1. set up the TBC-TGO-BC-SUB-Cuboid model in CATIA software, be designated as model A, model A represents to apply the Turbine Blade Model of tack coat, oxide layer and thermal barrier coating, saves as the TBC-TGO-BC-SUB-Cuboid-A.stp file;
2. set up the TGO-BC-SUB-Cuboid model in CATIA software, be designated as Model B, Model B represents to apply the Turbine Blade Model of tack coat and oxide layer, saves as the TGO-BC-SUB-Cuboid-B.stp file;
3. set up the BC-SUB-Cuboid model in CATIA software, be designated as MODEL C, MODEL C represents to apply the Turbine Blade Model of tack coat, saves as the BC-SUB-Cuboid-C.stp file;
4. set up the SUB-Cuboid model in CATIA software, be designated as model D, model D represents turbo blade substrate model, does not promptly apply the turbo blade of any coating, saves as the SUB-Cuboid-D.stp file;
5. in CATIA software, to the turbo blade of band base, chamfering is carried out in blade and base connecting portion and each layer coating junction operate, each chamfer radius is identical, and size is r;
(2) secondary treating of geometric model
The data file of middle A, the B that sets up of step (1), C, four models of D is imported in the ASEEMBLY module of large commercial finite element software ABAQUS software:
1. model A and Model B is tangent, generate the TBC layer, the model file of TBC layer is saved as TBC.stp;
2. Model B and MODEL C is tangent, generate the TGO layer, the model file of TGO layer is saved as TGO.stp;
3. MODEL C and model D is tangent, generate the BC layer, the model file of BC layer is saved as BC.stp;
(3) how much cleanings of each layer model
With TBC layer, TGO layer and the BC layer three layer model data importing CATIA software that generates in the step (2), carry out how much cleaning works of model, leave out unnecessary part, obtain being used for the model of finite element analysis, will clear up good model then and import ABAQUS software respectively;
(4) the final generation of thermal barrier coating of turbine blade system model
Do not consider destruction, the slippage at interface, set the intact combination in interface, in the ASEEMBLY module in ABAQUS software four layer models of setting up are previously merged, promptly merge TBC.stp, TGO.stp, BC.stp and four pairing models of file of SUB-Cuboid.stp, obtain whole thermal barrier coating system model, be designated as TBCs;
(5) respectively to TBC layer, TGO layer, BC layer, blade base or blade base and base definition material parameter;
(6) mechanical boundary condition and calorifics boundary condition are set;
(7) divide grid;
(8) preserve model data, adopt latent trial and error procedure, calculate;
(9) carry out sign as a result, find out dangerous point and hazardous location.
2, the finite element modeling method that is used for thermal barrier coating of turbine blade system destruction process according to claim 1 is characterized in that, described h
cBe 0.15～1.60mm, described h
tBe 0.01～0.10mm, described h
bBe 0.05～0.20mm, described h
sBe 1.00～2.00mm, described h
dBe 10.00～100.00mm, described r is 0.50～2.00mm.
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