CN117077292A - Design method and system for inner and outer flow coupling structure of full-shielding turbine rear support plate - Google Patents
Design method and system for inner and outer flow coupling structure of full-shielding turbine rear support plate Download PDFInfo
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Abstract
The invention provides a design method and a system for an internal and external flow coupling structure of a full-shielding turbine rear support plate, which are used for respectively modeling cooling chambers of characteristic sections with different blade heights and ensuring uniform wall thickness and stress distribution; defining a two-dimensional shape of each characteristic section according to the relative axial chord position, offset distance and chamber length of each cooling chamber; obtaining a three-dimensional model of each cooling cavity through a skin method and Boolean operation; the positioning of the air film holes and the impact holes on each characteristic section is determined through the relative arc length positions of the inner cavity wall surfaces, the spatial positioning of each cooling hole is determined through the distance from the hub and the casing and the number of the cooling holes, the geometric configuration design of each cooling hole is completed by combining the aperture and the inclination angle, the three-dimensional cooling structure modeling is obtained by rounding on the edge line of each cavity, the required design variables are few, the adaptability to the blade shape is strong, the degree of automation is high, and the full-shielding turbine rear support plate with the impact-air film composite cooling structure is beneficial to the efficient and quick design.
Description
Technical Field
The invention belongs to the field of aeroengine research and development design, and particularly relates to a method and a system for designing an internal and external flow coupling structure of a rear support plate of a full-shielding turbine.
Background
Under the condition that the engine speed and the primary parameters of gas at the inlet of the turbine are unchanged, the afterburner adopted by the military aircraft engine can further improve the gas temperature of the tail jet flow of the engine by injecting fuel into the gas flow after the turbine for re-combustion, thereby providing stronger thrust for the maneuver of the fighter plane. In order to realize low flow resistance, high reliability, light weight and strong stealth design of parts such as a turbine rear casing and an afterburner, the multifunctional integrated design of the aircraft engine turbine rear casing and the afterburner is a leading-edge research direction in the field of aircraft engines.
The general company in the United states first proposes an afterburner and turbine rear casing integrated design concept, and is successfully applied to an F119 aeroengine, so that the technical targets of higher combustion efficiency and lower flow loss are achieved. The research in this aspect is started later in China, and the following technical bottlenecks mainly exist in the aspect of the integrated design of the rear casing and afterburner of the turbine at present: firstly, with the rise of the temperature before the turbine, the temperature of afterburning is also increased year by year, so that the thermal load of the rear part of the turbine is increased year by year, and the cooling protection of the rear part of the turbine, particularly the integrated support plate, needs to be finely designed. Secondly, the demand of new generation fighter plane for the backward infrared stealth of the engine is higher and higher, and the high-temperature radiation source of the exhaust system of the engine, namely the final low-pressure turbine, needs to be effectively shielded. Under the background, the design of the full shielding integrated turbine rear support plate is proposed, and the design thought is to modify the rectifying support plate at the downstream of the final low-pressure turbine so as to geometrically realize full shielding of the final low-pressure turbine blades. Meanwhile, cold air is introduced from the outer duct to cool the surface of the full-shielding support plate, so that the infrared radiation intensity of the full-shielding support plate is reduced. Compared with the cooling of the rotary part of the final low-pressure turbine rotor, the cooling of the stationary rectifying support plate is higher in feasibility and lower in risk. Therefore, the efficient cooling layout of the full-shielding turbine rear support plate becomes one of key technologies.
The flow in the turbine aft support is typically a diffusion flow, especially for a fully shielded support, where the flow inside would undergo a forward pressure followed by diffusion flow process, which is quite different from the flow conditions inside the turbine. Therefore, the existing cooling structure design method for the air-cooled turbine is not suitable for the full-shielding turbine rear support plate, and development of a novel cooling structure design method matched with the flow characteristic of the full-shielding support plate is needed.
Disclosure of Invention
In order to solve the problems in the prior art, the invention provides a design method of an internal and external flow coupling structure of a fully-shielded turbine rear support plate, and provides a parameterized design method of a composite cooling structure aiming at the secondary deflection geometric characteristics and diffusion flow characteristics of the fully-shielded turbine rear support plate, wherein the parameterized design method is mainly used for cooling the rear visible surface of the fully-shielded turbine rear support plate, and is hopeful to provide a method foundation and technical support for research and development of the turbine rear support plate with high reliability and strong infrared stealth efficiency.
In order to achieve the above purpose, the technical scheme adopted by the invention is as follows: a design method of an internal and external flow coupling structure of a full-shielding turbine rear support plate comprises the following steps:
s1, selecting characteristic sections with different blade heights to respectively carry out cooling chamber configurations, and determining two-dimensional molded lines of the first cooling chamber to the fifth cooling chamber in each characteristic section according to local blade profile molded lines, offset distances and relative axial positions of each cooling chamber aiming at the characteristic sections;
s2, setting circumferential and axial stacking parameters according to the appearance of the blade, and obtaining the three-dimensional shape of each cooling cavity by a masking method based on the two-dimensional molded line obtained in the S1;
s3, obtaining two-dimensional molded lines of tail edge split joints on each characteristic section according to the positions of the fifth cooling chambers and the split joint widths at the characteristic sections with different blade heights;
s4, obtaining a preliminary three-dimensional shape of the split joint along the blade height direction through a covering method based on two-dimensional molded lines of the split joint of the tail edge at each characteristic section, and obtaining a split joint structure of the tail edge in discrete distribution through Boolean operation according to the distance between the split joint and the hub and the casing and the number of the split joints;
s5, arranging the impact holes and the air film holes on each impact cavity and each air film cavity, specifically, determining punching positions of the cooling holes according to relative positions along the main flow direction and the blade height direction, and punching according to the diameter of each cooling hole and the inclination angle of the opposite wall surface;
s6, chamfering is carried out on the edges of the first cooling chamber, the fifth cooling chamber and the tail edge split joint structure, and the full-shielding turbine rear support plate three-dimensional manufacturing model with the complete impact-air film composite cooling structure is obtained.
The values of the specific geometric parameters are given as follows:
relative position S of the first cooling chamber from the leading edge 1 From which the distance t is offset LE Determining the cavity offset distance t of the first cooling chamber LE The value is 2 times of the diameter d of the air film hole FC The method comprises the steps of carrying out a first treatment on the surface of the Second to 5 Cooling Chamber center to front edge position S 2 Respectively taking positions of 20%,40%,55% and 87% relative to axial chord lengths;
film cavity offset distance t of second to 5 cooling chambers FC The value is 2 times of the diameter d of the air film hole FC Width w FC The value is 2 times of the diameter d of the air film hole FC Length L FC The value is 13% of the axial chord length; offset distance t of impact cavity IMP The value is 5-6 times of the diameter d of the air film hole FC Width w IMP From blade profile and offset distance t IMP Determine length L IMP The axial chord length is 9.6%, and the diameter of the impact hole is 2 times of the diameter of the air film holeDiameter d FC The method comprises the steps of carrying out a first treatment on the surface of the Slit width w SL The fifth cooling chamber width was taken as 50%.
The two-dimensional modeling of the first cooling chamber is specifically as follows: offset inward distance t from the suction side and pressure side lines of the blade LE For the offset molded line, the distance is 4 times t from the front edge LE Cutting off at a position, wherein the axial distance S between the center of the first cooling chamber and the front edge 1 Is 2.5 times t LE ；
And connecting the cut pressure surface side and suction surface side deflection curves by using a straight line to obtain two-dimensional molded lines of the first cooling cavity in different characteristic sections.
When the second to fifth cooling chambers are molded in two dimensions,
the suction surface and the pressure surface side molded lines of the characteristic section are offset inwards three times, and the pitch arc of the blade and the distance S from the front edge of the second cooling chamber to the fifth cooling chamber are used for 2 、S 3 、S 4 S and S 5 Determining the positions of the second cooling chamber to the fifth cooling chamber impact chamber center point;
determining the central position of the cooling chamber molded line on each offset curve according to the positions of the impact chamber central points of the second cooling chamber and the fifth cooling chamber and the normal vector of the local camber line;
determining the molded lines on two sides of the second cooling chamber according to the lengths of the air film chambers and the impact chamber of the second cooling chamber to the fifth cooling chamber;
and connecting the corresponding molded lines to obtain the two-dimensional modeling of the impact cavity and the air film cavity from the second cooling cavity to the fifth cooling cavity on each characteristic section.
Taking the midpoint of the cooling chamber wall surface molded line of the impact chamber of the current cooling chamber, which is close to one side of the blade trailing edge, and connecting the midpoint with the tail edge point of the support plate blade to form a straight line, and then respectively biasing the straight line to the pressure surface and the suction surface by 0.5 times of the tail edge slit width w SL And obtaining the two-dimensional molded line of the tail edge split joint on the corresponding characteristic section of the current cooling cavity.
In S5, only the air film cavity is designed for the first cooling cavity, and the air film hole arrangement specifically comprises the following steps:
taking a cavity front edge point of a first cooling cavity as a punching center of a first row of air film holes, taking midpoints of pressure side and suction side cavity molded lines as punching centers of a second row of air film holes and a third row of air film holes, and connecting the air film hole punching centers on 3 characteristic sections to obtain a positioning line of each row of air film holes;
according to the distance h between the air film hole array and the hub s Distance h between the gas film hole row and the casing i Air film hole spacing h H Determining positions and the number of the air film holes along the height direction of the blade;
according to the diameter d of the air film hole FC And the air film hole is arranged by forming an included angle between the center line of the air film hole and the local wall surface.
h s And h i The value is 5% of the local leaf height, and the air film hole distance h is determined according to the number of each row of air film holes H 。
In S5, the impingement hole arrangement for the second to fifth cooling chambers specifically comprises the steps of:
in the impact cavities of the second cooling cavity to the fifth cooling cavity, taking the positions of 25% and 75% of arc length of the inner wall surface on one side of the rear-oriented visible surface as the punching centers of the impact holes, and connecting the punching centers of the impact holes on 3 characteristic sections to obtain the positioning lines of each row of the impact holes;
in the blade height direction, the distance h between the impact hole row and the hub s Impact hole row and case h i The air film hole spacing is used for determining the space positioning of all impact holes in the second cooling chamber to the fifth cooling chamber; the impact hole spacing is 2 times of the air film hole spacing;
based on the spatial positioning of all impingement holes in the second to fifth cooling chambers and the impingement hole diameter d IMP The design of the impingement holes is done in a direction perpendicular to the local wall.
In S5, the air film hole arrangement for the second cooling chamber to the fifth cooling chamber specifically includes the following steps:
in the air film cavities of the second to fifth cooling chambers, 10%, 50% and 90% arc length positions of the inner wall surface on one side of the rear-oriented visible surface are taken as punching centers of air film holes, and the air film hole punching centers on 3 characteristic sections are connected to obtain positioning lines of each row of air film holes;
in the blade height direction, according to the distance h between the air film hole array and the hub s Distance h between the gas film hole row and the casing i Air film hole spacing h H Determining the space positioning of all air film holes in a fifth cooling chamber of the second cooling chamber;
according to the space positioning of all the air film holes in the fifth cooling chamber of the second cooling chamber and the diameter d of the air film holes FC And the air film hole is designed by the air film hole inclination angle.
The invention simultaneously provides an internal and external flow coupling structure design system of the full-shielding turbine rear support plate, which comprises a characteristic acquisition module, a three-dimensional modeling acquisition module, a two-dimensional molded line acquisition module of a characteristic section, a trailing edge split joint structure acquisition module, a hole structure acquisition module and a manufacturing model acquisition module;
the characteristic acquisition module is used for selecting characteristic sections with different blade heights to respectively carry out cooling chamber configurations, and determining two-dimensional molded lines of the first cooling chamber to the fifth cooling chamber in each characteristic section according to local blade profile molded lines, offset distances and relative axial positions of each cooling chamber aiming at the characteristic sections;
the three-dimensional modeling acquisition module is used for setting circumferential and axial stacking parameters according to the appearance of the blade, and obtaining the three-dimensional modeling of each cooling cavity through a masking method based on the obtained two-dimensional molded line;
the two-dimensional molded line acquisition module of the characteristic section is used for acquiring the two-dimensional molded line of the tail edge split at each characteristic section according to the position of the fifth cooling cavity and the split width at the characteristic sections of different blade heights;
the tail edge split joint structure acquisition module acquires a preliminary three-dimensional shape of the split joint along the blade height direction through a covering method based on two-dimensional molded lines of the split joint of the tail edge at each characteristic section, and then acquires the split joint structure of the tail edge in discrete distribution through Boolean operation according to the distance between the split joint and the hub and the casing and the number of the split joints;
the hole structure acquisition module is used for arranging the impact holes and the air film holes on each impact cavity and each air film cavity, specifically, the punching positions of the cooling holes are determined according to the relative positions along the main flow direction and the blade height direction, and then punching is carried out according to the diameter of each cooling hole and the inclination angle of the opposite wall surface;
the manufacturing model acquisition module is used for chamfering the edges of the first cooling chamber, the fifth cooling chamber and the trailing edge split joint structure to obtain the full-shielding turbine rear support plate three-dimensional manufacturing model with the complete impact-air film composite cooling structure.
Compared with the prior art, the invention has at least the following beneficial effects: according to the invention, cooling structure design is carried out on different characteristic sections, and the profile of the cooling cavity is obtained by offsetting the blade profile, so that relatively uniform wall thickness and stress distribution can be ensured; compared with the traditional air film cooling mode, the cooling mode combining impact and air film can provide higher comprehensive cooling efficiency; the back visible surface of the fully-shielded back support plate is subjected to key cooling, so that the back infrared radiation intensity can be greatly reduced under the same cold air quantity. The geometrical outline of the secondary deflection type of the full-shielding turbine rear support plate can be well adapted, and meanwhile, the rear visible surface can be subjected to key cooling so as to reduce the rear infrared radiation. In addition, the method carries out parameterization definition on the cooling structure, and can conveniently carry out parameter modification and parameter optimization.
Drawings
Fig. 1 is a schematic view of a fully shielded turbine aft support plate with an impingement-film composite cooling structure, where (a) is a side aft view and (b) is an aft view.
FIG. 2 shows a cooling arrangement of different blade height sections.
Fig. 3 shows the cooling layout parameters of the root section.
FIG. 4 is a definition of cooling structure parameters at different locations of the root section. Wherein (a) is a first cooling chamber structure; (b) a third cooling chamber structure; (c) a fifth cooling chamber structure;
fig. 5 is a side view of a fully shielded posterior buttress cooling structure.
In the drawing, 1-a hub; 2-a tail cone; 3-rear support plate blades with cooling structures; 4-a case; 31-blade root section; 32-blade middle section; 33 blade top sectionA noodle; 311-a first cooling chamber structure; 312-a second cooling chamber structure; 313-a third cooling chamber structure; 314-fourth cooling chamber structure; 315-fifth cooling chamber structure; cax (Cax) 1 -axial chord length of the root section; cax (Cax) 2 -axial chord length of the middle section; cax (Cax) 3 -axial chord of the top section; s is S 1 -the relative position of the first cooling chamber from the leading edge; s is S 2 -the relative position of the second cooling chamber from the leading edge; s is S 3 -the relative position of the third cooling chamber from the leading edge; s is S 4 -the relative position of the fourth cooling chamber from the leading edge; s is S 5 -the relative position of the fifth cooling chamber from the leading edge; t is t LE -the cavity offset distance of the first cooling chamber; d, d FC -diameter of the gas film holes; d, d IMP -the diameter of the impingement holes; t is t FC -film cavity offset distance of the second to 5 cooling chambers; t is t IMP -an impingement cavity offset distance of the second to 5 cooling chambers; w (w) FC -film cavity width of the second to 5 cooling chambers; w (w) IMP -impingement cavity width of the second to 5 cooling chambers; l (L) FC -film cavity length of the second to 5 cooling chambers; l (L) IMP -the impingement cavity length of the second to 5 cooling chambers; w (w) SL -a split width of the fifth cooling chamber; h is a H -distance of root air film hole from hub; h is a S -distance of top gas film hole from casing; h is a i -relative distance of the intermediate gas film holes.
Detailed Description
The method for designing the impact-air film composite cooling structure in the invention is described in detail by taking a full shielding turbine rear support plate as an example:
fig. 1 shows a schematic view of a fully shielded turbine aft support plate structure with an impingement-film composite cooling structure, (a) being a side view excluding the casing 4, and (b) being a rear view of the complete structure. As shown in the figure, a rectifying support plate 3 with a cooling structure is arranged between a hub 1 and a casing 4, a tail cone 2 is arranged behind the hub, high-temperature fuel gas at a low-pressure turbine outlet flows through the rectifying support plate 3 and forms a high-strength central vortex in a tail cone rear region through a divergent structure of the tail cone 2, and the fuel gas is decelerated to be combusted stably.
Fig. 2 shows the cooling structure modeling at selected characteristic cross sections of the impingement-film composite cooling structure design, a root section 31, a middle section 32, and a tip section 33, respectively.
FIGS. 3 and 4 illustrate specific parameter definitions of the geometry of the cooling cavity and cooling holes, taking the root section as an example;
fig. 5 shows the distribution of cooling holes on the backward visible surface of the fully-shielded support plate and a specific parameter definition.
A design method of an internal and external flow coupling structure of a full-shielding turbine rear support plate comprises the following steps:
s1, selecting 3 characteristic sections (a root section 31, a middle section 32 and a top section 33) with different blade heights to perform two-dimensional modeling of a cooling cavity, wherein modeling ideas of each section are consistent, and referring to FIG. 2;
s2, taking a root section as an example, the first cooling chamber (311) is firstly subjected to two-dimensional modeling. Offset inward a distance t from the suction side and pressure side profiles of the blade LE As an example, the offset distance t LE The value is the diameter d of the air film hole FC Is 2 times as large as the above. For the offset profile, at 4 times t from the leading edge LE At a position where it is truncated, so that the first cooling chamber centre is at an axial distance S from the leading edge 1 About 2.5 times t LE . And connecting the two curves by using one straight line to obtain the two-dimensional molded lines of the first cooling cavity in different characteristic sections. It should be noted that, considering that the leading edge of the support plate is thin and cannot be observed from the rear, the first cooling chamber is only provided with film cooling, and does not consider impingement cooling.
S3, taking the root section as an example, two-dimensional modeling is carried out on the second to fifth cooling chambers (312 to 315). Firstly, the side molded lines of the suction surface and the pressure surface of the characteristic section are inwardly biased for three times, and the distances are t respectively FC 、t FC +w FC T IMP . Wherein t is FC And w is equal to FC The value of (2) is 2 times of the diameter d of the air film hole FC ，t IMP The value of (2) is 5.5 times of the diameter d of the air film hole FC . Next, the pitch arc of the blade and the distances S from the leading edge of the second to fifth cooling chambers 2 、S 3 、S 4 S and S 5 And determining the positions of the central points of the impact cavities of the second cooling chamber and the fifth cooling chamber, and determining the central positions of the molded lines of the cooling chambers on each offset curve according to the positions and the normal vector of the local camber line. In the present example S 2 、S 3 、S 4 And S is equal to 5 The values are respectively 20%,40%,55% and 87% of axial chord lengths. Then, according to the lengths L of the film cavity and the impingement cavity of the second to fifth cooling chambers FC 、L IMP Two side molded lines of the air film cavity and the impact cavity are respectively determined, L FC And L is equal to IMP The values of (2) are 13% and 9.6% of the axial chord length respectively. Finally, connecting the corresponding molded lines to obtain two-dimensional shapes of the impact cavity and the air film cavity of the second to fifth cooling cavities on each characteristic section, referring to fig. 3 and 4;
s4, setting circumferential and axial stacking parameters according to the appearance of the blade, and performing three-dimensional stacking on two-dimensional molded lines of the cooling chambers corresponding to different characteristic sections, so that the three-dimensional modeling of each cooling chamber can be obtained through a masking method.
S5, taking the midpoint of the cooling chamber wall surface molded line of the impact chamber of the fifth cooling chamber 315, which is close to the blade trailing edge side, to be connected with the blade trailing edge point of the support plate to form a straight line, and then biasing the straight line to the pressure surface and the suction surface by 0.5 times w respectively SL And obtaining the two-dimensional molded line of the tail edge split joint on the characteristic section. Tail edge slit width w in the present example SL Take the value of the width W of the fifth cooling chamber IMP Half of (a) is provided. The two-dimensional molded lines on the corresponding characteristic sections of the rest cooling chambers are obtained by the same method.
S6, obtaining a preliminary three-dimensional modeling of the split joint along the leaf height direction through a surface covering method based on the split joint two-dimensional molded lines of the characteristic sections; then according to the distance between the split joint and the hub and the casing and the number n of the split joints SL The tail edge split joint structure with discrete distribution can be obtained through Boolean operation; there are a total of 10 split structures in the examples given by the present invention.
Referring to fig. 5, S7, since the leading edge of the blade of the stay is thin and is not observable from the rear, the first cooling chamber 311 is arranged with only film cooling, in other words, only the cooling chamberAnd the air film cavity is not provided with an impact cavity. The front edge point of the first cooling chamber 311 is taken as the punching center of the first row of air film holes, and the midpoint of the chamber molded lines of the pressure side and the suction side is taken as the punching centers of the second row and the third row of air film holes. And connecting the perforation centers of the air film holes on the 3 characteristic sections to obtain the positioning line of each row of air film holes. Further, according to the distance h between the air film hole array and the hub and the casing s And h i And air film hole pitch h H And determining the distribution of the air film holes along the blade height direction. Because the connection part of the support plate blade and the upper end wall and the lower end wall has a rounding structure, h is taken as an example s And h i The value is 5% of the local leaf height, and the air film hole distance h H Is determined by the number of each row of air film holes. Furthermore, the spatial positioning of all the film holes in the first cooling chamber can be determined. Finally, according to the diameter d of the air film hole FC And the air film hole arrangement is completed by the included angle (inclined angle) between the center line of the air film hole and the local wall surface. In this embodiment, the tilt angle of the first row of air film holes is given as 90 ° (corresponding to being perpendicular to the local line), and the tilt angles of the second and third rows of air film holes are given as 30 °.
S8, taking the positions of 25% and 75% of arc length of the inner wall surface on the side of the rear visible surface as the punching center of the impact holes in the impact cavities of the second to fifth cooling chambers (312 to 315), in other words, arranging 2 impact holes in one impact cavity. And connecting the punching centers of the impact holes on the 3 characteristic sections to obtain the positioning line of each row of the impact holes. In the blade height direction, the distance between the impact hole row and the hub and the casing is h respectively s And h i In agreement with the film holes, the impact hole pitch is 2 times (i.e. 2 h) H ). The spatial positioning of all impingement holes in the second through fifth cooling chambers may be determined. Finally, according to the diameter d of the impact hole IMP And punching the impact hole along the direction perpendicular to the local wall surface.
S9, taking the positions of 10%, 50% and 90% of arc length of the inner wall surface on the visible surface side as the punching center of the air film holes in the air film cavities of the second cooling chamber to the fifth cooling chamber (312-315), in other words, arranging 3 air film holes in one impact cavity. Connecting the air film hole punching centers on 3 characteristic sectionsAnd obtaining the positioning lines of each row of air film holes. In the blade height direction, the distance between the air film hole array and the hub and the casing is h respectively s And h i The distance between the air film holes is h H . The spatial positioning of all the film holes in the second cooling chamber to the fifth cooling chamber can be determined. Finally, according to the diameter d of the air film hole FC And the modeling of the air film hole is completed by the inclination angle of the air film hole. In the embodiment of the invention, the inclination angles of all the air film holes of the second cooling chamber to the fifth cooling chamber are 30 degrees, so that good air film adherence can be ensured.
S10, generating a rounding on each ridge line according to the structural strength and the processing and manufacturing requirements, wherein the rounding scheme adopted by the invention is as follows: forming a radius on a ridge line with a radius of 3mm in the impact cavities of the first cooling chamber and the second to fifth cooling chambers, and forming a radius on the ridge line with a radius of 1mm in the air film cavities of the second to fifth cooling chambers; in the blade trailing edge split, a radius of 2mm is used as a radius to form a rounding on the edge line. Thus, the complete three-dimensional geometric model of the rear support plate with the impact-air film composite cooling structure for the full shielding turbine can be obtained.
The invention simultaneously provides an internal and external flow coupling structure design system of the full-shielding turbine rear support plate, which comprises a characteristic acquisition module, a three-dimensional modeling acquisition module, a two-dimensional molded line acquisition module of a characteristic section, a trailing edge split joint structure acquisition module, a hole structure acquisition module and a manufacturing model acquisition module;
the characteristic acquisition module is used for selecting characteristic sections with different blade heights to respectively carry out cooling chamber configurations, and determining two-dimensional molded lines of the first cooling chamber to the fifth cooling chamber in each characteristic section according to local blade profile molded lines, offset distances and relative axial positions of each cooling chamber aiming at the characteristic sections;
the three-dimensional modeling acquisition module is used for setting circumferential and axial stacking parameters according to the appearance of the blade, and obtaining the three-dimensional modeling of each cooling cavity through a masking method based on the obtained two-dimensional molded line;
the two-dimensional molded line acquisition module of the characteristic section is used for acquiring the two-dimensional molded line of the tail edge split at each characteristic section according to the position of the fifth cooling cavity and the split width at the characteristic sections of different blade heights;
the tail edge split joint structure acquisition module acquires a preliminary three-dimensional shape of the split joint along the blade height direction through a covering method based on two-dimensional molded lines of the split joint of the tail edge at each characteristic section, and then acquires the split joint structure of the tail edge in discrete distribution through Boolean operation according to the distance between the split joint and the hub and the casing and the number of the split joints;
the hole structure acquisition module is used for arranging the impact holes and the air film holes on each impact cavity and each air film cavity, specifically, the punching positions of the cooling holes are determined according to the relative positions along the main flow direction and the blade height direction, and then punching is carried out according to the diameter of each cooling hole and the inclination angle of the opposite wall surface;
the manufacturing model acquisition module is used for chamfering the edges of the first cooling chamber, the fifth cooling chamber and the trailing edge split joint structure to obtain the full-shielding turbine rear support plate three-dimensional manufacturing model with the complete impact-air film composite cooling structure.
The above is only for illustrating the technical idea of the present invention, and the protection scope of the present invention is not limited by this, and any modification made on the basis of the technical scheme according to the technical idea of the present invention falls within the protection scope of the claims of the present invention.
Claims (10)
1. The design method of the inner and outer flow coupling structure of the full-shielding turbine rear support plate is characterized by comprising the following steps of:
s1, selecting characteristic sections with different blade heights to respectively carry out cooling chamber configurations, and determining two-dimensional molded lines of the first cooling chamber to the fifth cooling chamber in each characteristic section according to local blade profile molded lines, offset distances and relative axial positions of each cooling chamber aiming at the characteristic sections;
s2, setting circumferential and axial stacking parameters according to the appearance of the blade, and obtaining the three-dimensional shape of each cooling cavity by a masking method based on the two-dimensional molded line obtained in the S1;
s3, obtaining two-dimensional molded lines of tail edge split joints on each characteristic section according to the positions of the fifth cooling chambers and the split joint widths at the characteristic sections with different blade heights;
s4, obtaining a preliminary three-dimensional shape of the split joint along the blade height direction through a covering method based on two-dimensional molded lines of the split joint of the tail edge at each characteristic section, and obtaining a split joint structure of the tail edge in discrete distribution through Boolean operation according to the distance between the split joint and the hub and the casing and the number of the split joints;
s5, arranging the impact holes and the air film holes on each impact cavity and each air film cavity, specifically, determining punching positions of the cooling holes according to relative positions along the main flow direction and the blade height direction, and punching according to the diameter of each cooling hole and the inclination angle of the opposite wall surface;
s6, chamfering is carried out on the edges of the first cooling chamber, the fifth cooling chamber and the tail edge split joint structure, and the full-shielding turbine rear support plate three-dimensional manufacturing model with the complete impact-air film composite cooling structure is obtained.
2. The method for designing an internal and external flow coupling structure of a rear support plate of a full-shielding turbine according to claim 1, wherein the specific geometric parameters are given by the following values:
relative position S of the first cooling chamber from the leading edge 1 From which the distance t is offset LE Determining the cavity offset distance t of the first cooling chamber LE The value is 2 times of the diameter d of the air film hole FC The method comprises the steps of carrying out a first treatment on the surface of the Second to 5 Cooling Chamber center to front edge position S 2 Respectively taking positions of 20%,40%,55% and 87% relative to axial chord lengths;
film cavity offset distance t of second to 5 cooling chambers FC The value is 2 times of the diameter d of the air film hole FC Width w FC The value is 2 times of the diameter d of the air film hole FC Length L FC The value is 13% of the axial chord length; offset distance t of impact cavity IMP The value is 5-6 times of the diameter d of the air film hole FC Width w IMP From blade profile and offset distance t IMP Determine length L IMP The axial chord length is 9.6 percent, and the diameter of the impact hole is 2 times of the diameter d of the air film hole FC The method comprises the steps of carrying out a first treatment on the surface of the Slit width w SL The fifth cooling chamber width was taken as 50%.
3. Root of Chinese characterThe method for designing an internal and external flow coupling structure of a full-shielding turbine rear support plate according to claim 1, wherein the two-dimensional modeling of the first cooling chamber is specifically as follows: offset inward distance t from the suction side and pressure side lines of the blade LE For the offset molded line, the distance is 4 times t from the front edge LE Cutting off at a position, wherein the axial distance S between the center of the first cooling chamber and the front edge 1 Is 2.5 times t LE ；
And connecting the cut pressure surface side and suction surface side deflection curves by using a straight line to obtain two-dimensional molded lines of the first cooling cavity in different characteristic sections.
4. The method for designing an internal and external flow coupling structure of a full-shielding turbine rear support plate according to claim 1, wherein when the second to fifth cooling chambers are two-dimensionally molded,
the suction surface and the pressure surface side molded lines of the characteristic section are offset inwards three times, and the pitch arc of the blade and the distance S from the front edge of the second cooling chamber to the fifth cooling chamber are used for 2 、S 3 、S 4 S and S 5 Determining the positions of the second cooling chamber to the fifth cooling chamber impact chamber center point;
determining the central position of the cooling chamber molded line on each offset curve according to the positions of the impact chamber central points of the second cooling chamber and the fifth cooling chamber and the normal vector of the local camber line;
determining the molded lines on two sides of the second cooling chamber according to the lengths of the air film chambers and the impact chamber of the second cooling chamber to the fifth cooling chamber;
and connecting the corresponding molded lines to obtain the two-dimensional modeling of the impact cavity and the air film cavity from the second cooling cavity to the fifth cooling cavity on each characteristic section.
5. The method for designing an internal and external flow coupling structure of a rear support plate of a fully shielded turbine according to claim 1, wherein a straight line is formed by connecting a midpoint of a cooling chamber wall surface molded line of an impingement chamber of a current cooling chamber near a trailing edge of a blade with a trailing edge point of a blade of the support plate, and then the straight line is respectively biased towards a pressure surface and a suction surfaceSetting 0.5 times of the width w of the split joint of the tail edge SL And obtaining the two-dimensional molded line of the tail edge split joint on the corresponding characteristic section of the current cooling cavity.
6. The method for designing an internal and external flow coupling structure of a full shielding turbine rear support plate according to claim 1, wherein in S5, only a gas film cavity is designed for the first cooling chamber, and the gas film hole arrangement specifically comprises the following steps:
taking a cavity front edge point of a first cooling cavity as a punching center of a first row of air film holes, taking midpoints of pressure side and suction side cavity molded lines as punching centers of a second row of air film holes and a third row of air film holes, and connecting the air film hole punching centers on 3 characteristic sections to obtain a positioning line of each row of air film holes;
according to the distance h between the air film hole array and the hub s Distance h between the gas film hole row and the casing i Air film hole spacing h H Determining positions and the number of the air film holes along the height direction of the blade;
according to the diameter d of the air film hole FC And the air film hole is arranged by forming an included angle between the center line of the air film hole and the local wall surface.
7. The method for designing an internal and external flow coupling structure of a rear support plate of a fully shielded turbine as set forth in claim 6, wherein h s And h i The value is 5% of the local leaf height, and the air film hole distance h is determined according to the number of each row of air film holes H 。
8. The method for designing an internal and external flow coupling structure of a full-shielding turbine rear support plate according to claim 1, wherein in S5, the arrangement of impingement holes for the second to fifth cooling chambers specifically includes the steps of:
in the impact cavities of the second cooling cavity to the fifth cooling cavity, taking the positions of 25% and 75% of arc length of the inner wall surface on one side of the rear-oriented visible surface as the punching centers of the impact holes, and connecting the punching centers of the impact holes on 3 characteristic sections to obtain the positioning lines of each row of the impact holes;
in the blade height direction, the impact hole array and the hubDistance h of (2) s Impact hole row and case h i The air film hole spacing is used for determining the space positioning of all impact holes in the second cooling chamber to the fifth cooling chamber; the impact hole spacing is 2 times of the air film hole spacing;
based on the spatial positioning of all impingement holes in the second to fifth cooling chambers and the impingement hole diameter d IMP The design of the impingement holes is done in a direction perpendicular to the local wall.
9. The method for designing an internal and external flow coupling structure of a full shielding turbine rear support plate according to claim 1, wherein in S5, the gas film hole arrangement for the second to fifth cooling chambers specifically includes the following steps:
in the air film cavities of the second to fifth cooling chambers, 10%, 50% and 90% arc length positions of the inner wall surface on one side of the rear-oriented visible surface are taken as punching centers of air film holes, and the air film hole punching centers on 3 characteristic sections are connected to obtain positioning lines of each row of air film holes;
in the blade height direction, according to the distance h between the air film hole array and the hub s Distance h between the gas film hole row and the casing i Air film hole spacing h H Determining the space positioning of all air film holes in a fifth cooling chamber of the second cooling chamber;
according to the space positioning of all the air film holes in the fifth cooling chamber of the second cooling chamber and the diameter d of the air film holes FC And the air film hole is designed by the air film hole inclination angle.
10. The design system is characterized by comprising a characteristic acquisition module, a three-dimensional modeling acquisition module, a two-dimensional molded line acquisition module of a characteristic section, a trailing edge split joint structure acquisition module, a hole structure acquisition module and a manufacturing model acquisition module;
the characteristic acquisition module is used for selecting characteristic sections with different blade heights to respectively carry out cooling chamber configurations, and determining two-dimensional molded lines of the first cooling chamber to the fifth cooling chamber in each characteristic section according to local blade profile molded lines, offset distances and relative axial positions of each cooling chamber aiming at the characteristic sections;
the three-dimensional modeling acquisition module is used for setting circumferential and axial stacking parameters according to the appearance of the blade, and obtaining the three-dimensional modeling of each cooling cavity through a masking method based on the obtained two-dimensional molded line;
the two-dimensional molded line acquisition module of the characteristic section is used for acquiring the two-dimensional molded line of the tail edge split at each characteristic section according to the position of the fifth cooling cavity and the split width at the characteristic sections of different blade heights;
the tail edge split joint structure acquisition module acquires a preliminary three-dimensional shape of the split joint along the blade height direction through a covering method based on two-dimensional molded lines of the split joint of the tail edge at each characteristic section, and then acquires the split joint structure of the tail edge in discrete distribution through Boolean operation according to the distance between the split joint and the hub and the casing and the number of the split joints;
the hole structure acquisition module is used for arranging the impact holes and the air film holes on each impact cavity and each air film cavity, specifically, the punching positions of the cooling holes are determined according to the relative positions along the main flow direction and the blade height direction, and then punching is carried out according to the diameter of each cooling hole and the inclination angle of the opposite wall surface;
the manufacturing model acquisition module is used for chamfering the edges of the first cooling chamber, the fifth cooling chamber and the trailing edge split joint structure to obtain the full-shielding turbine rear support plate three-dimensional manufacturing model with the complete impact-air film composite cooling structure.
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN117521563A (en) * | 2024-01-08 | 2024-02-06 | 中国空气动力研究与发展中心计算空气动力研究所 | Pneumatic data processing method based on impeller mechanical turbulence wall distance calculation |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN117521563A (en) * | 2024-01-08 | 2024-02-06 | 中国空气动力研究与发展中心计算空气动力研究所 | Pneumatic data processing method based on impeller mechanical turbulence wall distance calculation |
| CN117521563B (en) * | 2024-01-08 | 2024-03-15 | 中国空气动力研究与发展中心计算空气动力研究所 | Pneumatic data processing method based on impeller mechanical turbulence wall distance calculation |
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