CN115013089B - Method and system for designing rear turbine casing rectifying support plate with wide working condition backward shielding - Google Patents
Method and system for designing rear turbine casing rectifying support plate with wide working condition backward shielding Download PDFInfo
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- CN115013089B CN115013089B CN202210648464.8A CN202210648464A CN115013089B CN 115013089 B CN115013089 B CN 115013089B CN 202210648464 A CN202210648464 A CN 202210648464A CN 115013089 B CN115013089 B CN 115013089B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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- 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
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Abstract
The invention provides a method and a system for designing a turbine rear casing rectifying support plate with backward shielding under wide working conditions, wherein the method is used for respectively designing characteristic element sections with different blade heights so as to ensure good matching of three-dimensional flow of a low-pressure turbine outlet; for each characteristic element section, describing a blade profile by respectively utilizing an inlet geometric angle, an outlet geometric angle, a grid pitch, a chord length, a shielding width and an axial extension length; respectively obtaining molded lines of a pressure surface and a suction surface of the blade by using a secondary deflection curve method, and respectively modeling the front edge and the tail edge of the blade by using an ellipse and an arc in combination with radius parameters of the front edge and the tail edge; and realizing three-dimensional forming of the blade according to a required stacking mode. The method has the advantages of high degree of freedom, few required variables and flexible and convenient adjustment, can flexibly adjust the backward shielding rate of the final-stage low-pressure turbine while performing good pneumatic rectification, even realizes complete backward shielding, and is beneficial to quickly designing the rear casing part with low pneumatic loss and strong infrared stealth efficiency.
Description
Technical Field
The invention belongs to the field of design of aviation turbine engines, and particularly relates to a method and a system for designing a rear turbine casing rectifying support plate with backward shielding under a wide working condition, in particular to a parameterization design method for the rear turbine casing rectifying support plate with both rectification and backward shielding.
Background
In the aircraft stealth technology, radar stealth and infrared stealth are two main aspects. The gas turbine engine is used as the main power source of the fighter plane, and the air inlet/exhaust system of the gas turbine engine is the main radar scattering source and the infrared radiation source. The prerequisite for achieving stealth of a combat aircraft is to reduce the infrared and radar characteristics of the air intake/exhaust system as much as possible. The infrared detection utilizes the physical characteristics of the object, namely all objects with the temperature higher than absolute zero (-273.15 ℃) can emit infrared radiation to the outside at any time, and the intensity of the infrared radiation can be increased sharply along with the increase of the temperature. The gas turbine engine can discharge a large amount of high-temperature gas in the operation process, and the temperature is far higher than the ambient temperature, so that the backward infrared stealth difficulty of the airplane is greatly increased.
The key of infrared stealth of an engine exhaust system is to reduce the infrared radiation intensity of the exhaust system in the backward direction as much as possible. The specific method comprises the following steps: 1) By adjusting the components of the fuel gas and adding substances containing infrared shielding materials into the fuel gas, the sprayed fuel gas can form an infrared shielding layer, so that the infrared radiation of the tail flame is weakened; 2) An air injection technology, which introduces low-temperature gas of the surrounding environment into a tail nozzle and mixes the low-temperature gas with high-temperature fuel gas, thereby weakening the infrared radiation of the tail flame; 3) And shielding the high-temperature radiation source, and arranging a shielding device at the downstream of the low-pressure turbine so as to avoid or reduce the probability of detection by the infrared detection device.
The low-pressure turbine component for shielding high temperature by using the turbine rear casing rectifying support plate is a potential way for improving the infrared stealth efficiency of the aircraft engine. The turbine rear casing rectification support plate needs to give consideration to rectification, bearing and backward shielding functions, and the design and research difficulty is extremely high. On one hand, the rectifying support plate needs to guide the airflow after the final-stage low-pressure turbine in the axial direction so as to improve the thrust of the engine. On the other hand, the rectification extension board needs shelter from the low pressure turbine part of high temperature to reduce the infrared radiation intensity behind, and then promote the infrared stealthy efficiency behind the engine. At present, the design method of the turbine rear casing rectifying support plate capable of flexibly realizing backward full shielding is extremely deficient. The invention fully considers the design requirements of backward shielding and pneumatic rectification of the rectification support plate, provides a parameterization design method and a parameterization design system of the rectification support plate based on a secondary deflection curve, and can provide theoretical and technical support for researching and developing a turbine rear casing with wide working condition and backward shielding.
Disclosure of Invention
In order to solve the problems in the prior art, the invention provides a parameterization design method for a rectification support plate of a turbine rear casing of an aircraft engine, which can realize rectification and backward shielding. Compared with the existing design method, the method can conveniently and flexibly adjust the backward shielding rate of the final-stage low-pressure turbine while performing pneumatic rectification, even realize the backward complete shielding, is beneficial to quickly designing the rear casing part with low pneumatic loss and strong infrared stealth effect, reduces the temperature of the tail nozzle as far as possible under the condition of not reducing the running temperature of the engine, and enables the outlet airflow of the tail nozzle to be along the axial direction of the engine as far as possible so as to realize the maximization of the thrust.
In order to achieve the purpose, the invention adopts the technical scheme that: a method for designing a turbine rear casing rectifying support plate capable of shielding backwards under wide working conditions comprises the following steps:
s1, selecting characteristic sections with different leaf heights to respectively shape; aiming at the different characteristic cross sections, the different characteristic cross sections respectively comprise a turning section and an axial extension section; the turning section is controlled by 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid distance and shielding width, and the axial extension section is controlled by the axial extension length;
s2, respectively obtaining a turning section pressure surface molded line and a turning section suction surface molded line which meet 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, a chord length, a grid pitch and a shielding width by using a secondary deflection curve method;
s3, obtaining a pressure surface molded line and a suction surface molded line of the axial extension section through setting the axial extension length behind the turning section, and respectively combining the pressure surface molded line and the suction surface molded line of the turning section and the axial extension section to obtain a complete geometric molded line of a pressure surface and a suction surface of the blade;
s4, determining the diameters of the front edge and the tail edge of the blade according to the structural requirements, and modeling the front edge and the tail edge of the blade by utilizing an ellipse and an arc respectively so as to obtain initial primitive blade profiles with different characteristic sections;
s5, respectively fitting the complete geometric molded lines of the pressure surface and the suction surface of the blade obtained in the S3; meanwhile, adding a plurality of control points at different axial positions, adjusting and iterating the molded lines of the blade, and modeling the front edge and the tail edge of the blade by combining the obtained molded lines to obtain an elementary blade profile meeting the pneumatic performance requirement;
and S6, determining a blade stacking form, setting circumferential and axial stacking parameters, and combining the blade stacking form and the element blade profile obtained in the S5 to obtain the three-dimensional shape of the blade.
And S1, selecting at least 3 different characteristic element sections for modeling respectively.
The inlet geometric angle is determined according to the inlet airflow angle at the local leaf height so as to reduce the flow separation of the inlet; the outlet geometry angle is set to 90 degrees to achieve axial venting; the chord length is adjusted according to the structural size limitation and the flow turning angle; the grid pitch is determined according to the number of the blades of the rectifying support plate 2 and the pitch diameter of the local blade height; by using the width of the shadedThe backward shielding rate of the final-stage low-pressure turbine part is accurately controlled, and the full shielding design is realized; by axial extensionlTo determine the length of the axial extension.
Width of shieldingdSpecifically, the method comprises the following steps: the projection coincidence distance of the front edge of the current rectifying support plate 2 and the tail edge of the adjacent rectifying support plate 2 on the cross section of the outlet of the engine is determined if the front edge of the current rectifying support plate and the tail edge of the adjacent rectifying support plate 2 coincide with each otherdIs positive; if the two are not coincident, thendIs negative whendWhen the number is 0, the rectifying support plate 2 just blocks the backward observation sight from the tail of the engine, and the high-temperature low-pressure turbine part is shielded forward and backward;dand the observation sight angle shielded by the rectification support plate 2 is increased along with the increase of the angle.
And S5, respectively fitting the molded lines of the pressure surface and the suction surface of the blade obtained in the S3 by using a non-uniform B-spline curve.
In S2, the molded line is integrally divided into a front edge section, a middle section and a tail edge section by a secondary deflection curve method, wherein the front edge section is a 1-section Bezier spline curve, and deflection from the near-axial direction to the circumferential direction is realized; the tail edge section is a 1-section Bezier spline curve, and deflection from the circumferential direction to the axial direction is realized; the middle section is a Bezier spline curve with 1 section or 2 sections and connects the front edge section and the tail edge section.
The blade root adopts a C-shaped flow passage outline, and the blade top adopts an S-shaped flow passage outline.
In addition, the invention provides a design system of a turbine rear casing rectification support plate with wide-working-condition backward shielding, which comprises a design element determining module, a molded line design module, an initial primitive blade profile obtaining module, a primitive blade profile updating module and a blade three-dimensional modeling design module.
The design element determining module is used for selecting characteristic sections with different blade heights to respectively shape; aiming at the different characteristic cross sections, the different characteristic cross sections respectively comprise a turning section and an axial extension section; the turning section is controlled by 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid distance and shielding width, and the axial extension section is controlled by the axial extension length;
the molded line design module firstly utilizes a secondary deflection curve method to respectively obtain a turning section pressure surface molded line and a turning section suction surface molded line which meet 5 geometric parameters of an inlet geometric angle, an outlet geometric angle, chord length, grid spacing and shielding width; then obtaining a pressure surface molded line and a suction surface molded line of an axial extension section by giving an axial extension length behind the turning section, and respectively combining the pressure surface molded line and the suction surface molded line of the turning section and the axial extension section to obtain a complete geometric molded line of a pressure surface and a suction surface of the blade;
the initial primitive blade profile obtaining module is used for determining the diameters of the front edge and the tail edge of the blade according to the structural requirements, and modeling the front edge and the tail edge of the blade by utilizing an ellipse and an arc respectively so as to obtain initial primitive blade profiles with different characteristic sections;
the element blade profile updating module is used for respectively fitting complete geometric molded lines of the pressure surface and the suction surface of the obtained blade; meanwhile, adding a plurality of control points at different axial positions, adjusting and iterating the molded lines of the blade, and modeling the front edge and the tail edge of the blade by combining the obtained molded lines to obtain an elementary blade profile meeting the requirement of pneumatic performance;
and the blade three-dimensional modeling design module is used for determining a blade stacking form, setting circumferential and axial stacking parameters, and combining the blade stacking form and the element blade profile obtained in the S5 to obtain the three-dimensional modeling of the blade.
The invention also provides computer equipment which comprises a processor and a memory, wherein the memory is used for storing the computer executable program, the processor reads the computer executable program from the memory and executes the computer executable program, and the processor can realize the design method of the turbine rear casing rectifying support plate for the rear shielding under the wide working condition when executing the computer executable program.
The invention can also provide a computer readable storage medium, wherein a computer program is stored in the computer readable storage medium, and when the computer program is executed by a processor, the method for designing the turbine rear casing rectifying support plate with wide-working-condition backward shielding can be realized.
Compared with the prior art, the invention has at least the following beneficial effects:
the invention designs the characteristic element sections with different blade heights respectively, ensures good matching of three-dimensional flow at the outlet of the low-pressure turbine, and describes the geometric profile of the blade by respectively utilizing the inlet geometric angle, the outlet geometric angle, the grid pitch, the chord length, the shielding width and the length of the axial extension section for each characteristic element section. On one hand, the molded lines of the pressure surface and the suction surface of the blade are controlled by utilizing a secondary deflection curve method and combining typical characteristic geometric parameters, and the front edge and the tail edge of the blade are molded by utilizing an ellipse and an arc; the method can meet the design requirement of directly adjusting the backward shielding rate (including full shielding) through the shielding width, and has the advantages of clear physical significance of geometric parameters, convenience and flexibility in adjustment and high modeling precision; on the other hand, the blade profile line fine adjustment based on the control points and the three-dimensional stacking forming of the blades are fully considered, and the geometric profile of the rectifying support plate with high pneumatic performance and backward full shielding can be better designed.
The design method of the rectification support plate can conveniently and flexibly adjust the backward shielding rate of the final-stage low-pressure turbine while performing pneumatic rectification, even realize the backward complete shielding, is beneficial to quickly designing the rear casing part with low pneumatic loss and strong infrared stealth effect, reduces the temperature of the tail nozzle as far as possible under the condition of not reducing the running temperature of the engine, and enables the outlet airflow of the tail nozzle to be along the axial direction of the engine as far as possible so as to realize the maximization of the thrust; the obtained rectifying support plate of the rear turbine casing is positioned behind the final-stage low-pressure turbine and in front of the afterburning flame stabilizer, and on one hand, the rectifying support plate guides airflow at the outlet of the low-pressure turbine in the axial direction, and on the other hand, backward shielding of low-pressure turbine parts is realized through molded line bending.
Drawings
Fig. 1 is a schematic layout of a turbine rear casing fairing support plate, wherein (a) is a schematic partial sectional view in a three-dimensional manner, and (b) is a schematic structural end face view.
Fig. 2 is a three-dimensional schematic view of a rectifying support plate and schematic views of profiles with different blade heights.
FIG. 3 is a diagram of different blade height primitive profiles and their design parameters, wherein (a) is the root design profile and design parameters, (b) is the in-blade design profile and design parameters, and (c) is the tip design profile and design parameters.
In the drawings, 1-hub; 2, a rectifying support plate; 3, a case; 4-a tail cone; 21-blade root profile; 22-profile in blade airfoil; 23-blade tip profile; 211-the blade root turning section pressure surface; 212-suction surface of blade root turning section; 213-pressure surface of axial extension section of blade root; 214-suction surface of axial extension section of blade root; 215-root flow channel center line; 221-pressure side of turning section in leaf; 222-suction surface of turning segment in leaf; 223-pressure surface of axial extension section in the lobe; 224-the suction surface of the axial extension section of the lobe; 225-center line of flow channel in lobe; 231-pressure surface of turning section of blade top; 232-suction surface of turning section of leaf top; 233-pressure surface of axial extension section of blade tip; 234-suction surface of axial extension section of blade tip; 235-blade tip flowpath centerline.
Detailed Description
In order to make the technical solutions and advantages related to the present application clearer, the technical solutions related to the present application will be explained in more detail below with reference to the drawings in the description of the drawings of the present application. The embodiments described herein are merely some embodiments and not all embodiments of the present application, and the explanation with reference to the drawings is for the purpose of explanation and should not be construed as a limitation of the present application. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the present application without inventive step, are within the scope of the present application. Embodiments of the present application will be described in detail below with reference to the drawings.
The invention provides a method for designing a turbine rear casing rectifying support plate with wide working condition backward shielding, which specifically comprises the following steps:
s1, considering that a low-pressure turbine part of an aeroengine belongs to a rotating part, the internal flow of the low-pressure turbine part of the aeroengine presents a high three-dimensional flow characteristic, and the flow states of different blade height sections are different. The design method selects 3 or more than 3 different characteristic sections to respectively shape so as to adapt to the airflow angle distribution flowing at the outlet of the low-pressure turbine. Aiming at the different characteristic cross sections, the different characteristic cross sections respectively comprise a turning section and an axial extension section; the turning section is controlled by 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid distance and shielding width, and the axial extension section is controlled by the axial extension length;
and S2, respectively obtaining the turning section pressure surface molded line and the turning section suction surface molded line which meet 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid spacing and shielding width by using a secondary deflection curve method. The molded line is integrally divided into a front edge section, a middle section and a tail edge section by a secondary deflection curve method, wherein the front edge section is a 1-section Bezier spline curve, so that deflection from the near-axial direction to the circumferential direction is realized; the tail edge section is a 1-section Bezier spline curve, and deflection from the circumferential direction to the axial direction is realized; the middle section is a Bezier spline curve with 1 section or 2 sections and connects the front edge section and the tail edge section. The inlet geometry angle is determined based on the inlet airflow angle at the current leaf height to reduce inlet flow separation. The outlet geometry angle is typically set at 90 degrees to achieve axial venting. The chord length is adjusted according to the structural size limitation and the flow turning angle. The grid distance is determined according to the number of blades of the rectifying support plate 2 and the pitch diameter of the local blade height. The shielding width parameter can accurately control the backward shielding rate of the final-stage low-pressure turbine part, and the full shielding design is realized.
S3, obtaining a pressure surface molded line and a suction surface molded line of the axial extension section through setting the axial extension length behind the turning section, and respectively combining the pressure surface molded line and the suction surface molded line of the turning section and the axial extension section to obtain a complete geometric molded line of a pressure surface and a suction surface of the blade;
s4, determining the diameters of the front edge and the tail edge of the blade according to the structural requirements, and modeling the front edge and the tail edge of the blade by utilizing an ellipse and an arc respectively so as to obtain initial primitive blade profiles with different characteristic sections;
and S5, considering that the rectifying support plate 2 and the meridian flow path are diffusion flow, the flow in the rectifying support plate is sensitive to geometric molded lines, and fine design is needed. And respectively fitting the complete geometric molded lines of the pressure surface and the suction surface of the blade by using a non-uniform B-spline curve. On the basis, a plurality of control points are added at different axial positions, and the molded lines of the blade, particularly the molded lines of the suction surface, are subjected to fine tuning iteration to obtain the primitive blade profile meeting the requirement of pneumatic performance.
S6, three-dimensional blade modeling is carried out, and firstly, a blade stacking form is determined, wherein the stacking form comprises stacking of a front edge and a tail edge. And further setting circumferential and axial stacking parameters to realize the curved and swept three-dimensional modeling of the blade, and combining the primitive blade profiles with different blade heights and the three-dimensional stacking rule to obtain the three-dimensional modeling of the blade.
Specifically, as shown in fig. 1, the rectifying support plate 2 is located after the last stage low-pressure turbine of the aircraft engine and before the afterburner flame stabilizer; the rectifying support plate 2 is arranged on the hub 1 and the casing 3, the hub 1 and the casing 3 are coaxially arranged, the tail cone 4 is smoothly connected with the hub 1, and the tail cone 4 is arranged at an air outlet. The rectifying support plates 2 are uniformly distributed on the hub 1, so that the air flow at the outlet of the final-stage low-pressure turbine can be rectified. More importantly, the bent line of the rectifying support plate 2 can also shield the final-stage low-pressure turbine from the rear direction, so that the infrared radiation signal from the rear direction of the engine is reduced.
Considering that the low-pressure turbine part of an aircraft engine belongs to a rotating part, the outlet airflow of the low-pressure turbine part of the aircraft engine presents a high three-dimensional characteristic and has different outlet airflow angles in different blade height sections. In order to adapt to the distribution of the air flow angle flowing at the outlet of the low-pressure turbine, the section of 3 or more different characteristic elements is selected for modeling, and as can be seen from fig. 2, three sections, namely a blade root profile 21, a blade leaf profile 22 and a blade top profile 23, have different geometric characteristics respectively, so that the flow and shielding requirements of different blade height areas can be well met.
FIG. 3 shows a parameterized design method and modeling example of the primitive leaf profile. The different characteristic sections are composed of turning sections and axial extension sections. The turning sections of which pass through the inlet geometric angle (alpha) respectively 1 、α 2 、α 3 ) Outlet geometry angle (beta) 1 、β 2 、β 3 ) Chord length (C 1 、C 2 、C 3 ) Grid pitch (a)P 1 、P 2 、P 3 ) And occlusion width (d 1 、d 2 、d 3 ) A total of 5 geometric parameters were controlled separately. Inlet geometry angle
The inlet attack angle can be effectively reduced according to the determination of the inlet airflow angle at the high position of the ground vane, so that the flow separation phenomenon at the inlet of the rectifying support plate 2 is avoided. Outlet geometry angle (beta) 1 、β 2 、β 3 ) Typically set at about 90 degrees to maximize axial exhaust and thrust. Chord length (C 1 、C 2 、C 3 ) Adjustments are made based on axial dimensional limitations and flow turning angles. Grid pitch (P 1 、P 2 、P 3 ) And determining according to the number of blades of the rectifying support plate 2 and the pitch diameter of the local blade height. The axial extension is mainly defined by an axial extension length (l 1 、l 2 、l 3 ) And (5) controlling.
In particular, it is possible to use, for example, the design method utilizes the shielding width (d 1 、d 2 、d 3 ) The backward shielding rate of the final-stage low-pressure turbine part is accurately controlled. The parameters are defined as: the projection coincidence distance of the front edge of the current rectifying support plate 2 and the tail edge of the adjacent rectifying support plate 2 on the cross section of the outlet of the engine. If the two are superposed, thendIs positive; if the two are not coincident, thendIs negative. When in usedAnd when the number is 0, the representative rectifying support plate 2 just can block the backward observation sight from the tail part of the engine, and the high-temperature low-pressure turbine part is shielded forward and backward. With followingdThe larger the angle of the observation sight line which can be shielded by the rectification support plate 2 is, the larger the shielding effect of the side and the rear can be enhanced. In the example shown in FIG. 3, the occlusion widthdA small positive value is used.
Then, for the turning section, according to the above 5 geometric parameters, the blade root pressure surface profile 211, the blade leaf pressure surface profile 221, the blade tip pressure surface profile 231, the blade root suction surface profile 212, the blade leaf suction surface profile 222, and the blade tip suction surface profile 232 are respectively described by using a secondary deflection curve; for the axial extension segment, according to the axial extension length, a blade root pressure surface molded line 213, a blade leaf pressure surface molded line 223, a blade top pressure surface molded line 233, a blade root suction surface molded line 214, a blade leaf suction surface molded line 224 and a blade top suction surface molded line 234 are respectively described by straight lines; and (4) combining the molded lines of the turning section and the axial extension section to obtain the complete geometric molded line of the pressure surface and the suction surface of the blade.
Further, the diameters of the front edge and the tail edge of the blade are determined according to the structural requirements, and the front edge and the tail edge of the blade are modeled by utilizing an ellipse and an arc respectively, so that initial primitive blade profiles with different characteristic sections are obtained; through the steps, the primitive leaf profiles with different characteristic sections can be efficiently and quickly obtained.
Further, as can be seen from the modeling example of the full-shielding rectification support plate 2, the geometric characteristics of the blade profiles of the rectification support plates 2 with different blade heights are completely different. At the blade root, the airflow angle of the incoming flow is small, the tangential velocity component is large, and the flow turning angle generated for realizing axial flow guiding is large. To achieve full aft shielding, the resulting flow path profile is "C" shaped, as shown by the root flow path centerline 215. At the top of the blade, the airflow angle of the incoming flow is large, the tangential velocity component is small, and the flow turning angle generated for realizing axial flow guiding is small. To achieve full aft shielding, the resulting flow path profile is "S" shaped, as shown by the tip flow path centerline 235. The reasons for the different flow channel profiles at different blade heights are: when the turning angle is larger, the circumferential deflection angle of the blade profile is larger, and axial flow guiding and backward shielding can be realized simultaneously with smaller chord length; when the turning angle is smaller, the circumferential deflection angle of the blade profile is smaller, and if the C-shaped profile is adopted to realize backward full shielding, a longer chord length is required, which is not beneficial to controlling the axial size; and the adoption of the S-shaped profile can realize axial flow guiding and backward full shielding with relatively shorter chord length.
Further, in order to carry out fine design on the blade profile of the rectifying support plate 2, a profile line fine tuning method based on a non-uniform B-spline curve is also developed. Firstly, the molded lines of the pressure surface and the suction surface of the blade are respectively fitted by using a non-uniform B spline curve. On the basis, a plurality of control points are generated at different axial positions, and the molded line of the blade, particularly the molded line of the suction surface, is subjected to fine tuning and iteration to obtain the elementary blade profile meeting the pneumatic performance requirement.
Finally, as shown in fig. 2, a certain three-dimensional stacking manner is selected to perform three-dimensional forming on the blade. First, a blade stacking pattern is determined, including leading edge and trailing edge stacking, etc. And further setting circumferential and axial stacking parameters to realize the curved and swept three-dimensional modeling of the blade. The three-dimensional modeling of the blade can be obtained by combining the primitive blade profiles with different blade heights and the three-dimensional stacking rule.
In addition, the invention can also provide computer equipment which comprises a processor and a memory, wherein the memory is used for storing computer executable programs, the processor reads part or all of the computer executable programs from the memory and executes the computer executable programs, and when the processor executes the part or all of the computer executable programs, the design method of the turbine rear casing rectifying support plate with the rear shielding under the wide working condition can be realized.
The invention also provides a design system of the turbine rear casing rectifying support plate with wide working condition backward shielding, which comprises a design element determining module, a molded line design module, an initial element blade profile obtaining module, an element blade profile updating module and a blade three-dimensional modeling design module.
The design factor determining module is used for selecting characteristic sections with different blade heights to respectively shape; aiming at the different characteristic cross sections, the different characteristic cross sections respectively comprise a turning section and an axial extension section; the turning section is controlled by 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid distance and shielding width, and the axial extension section is controlled by the axial extension length;
the molded line design module firstly utilizes a secondary deflection curve method to respectively obtain a turning section pressure surface molded line and a turning section suction surface molded line which meet 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid spacing and shielding width; then obtaining a geometric molded line of an extension section by giving an axial extension length behind the turning section, and combining the molded line of the pressure surface of the turning section and the molded line of the suction surface of the turning section to obtain a complete geometric molded line of the pressure surface and the suction surface of the blade;
the initial element blade profile acquisition module is used for determining the diameters of the front edge and the tail edge of the blade according to the structural requirements, and modeling the front edge and the tail edge of the blade by utilizing an ellipse and an arc respectively so as to obtain initial element blade profiles with different characteristic sections;
the element blade profile updating module is used for respectively fitting complete geometric molded lines of the pressure surface and the suction surface of the obtained blade; meanwhile, adding a plurality of control points at different axial positions, adjusting and iterating the molded lines of the blade, and modeling the front edge and the tail edge of the blade by combining the obtained molded lines to obtain an elementary blade profile meeting the requirement of pneumatic performance;
and the blade three-dimensional modeling design module is used for determining the stacking form of the blades, setting circumferential and axial stacking parameters, and combining the stacking form of the blades and the element blade model obtained in the S5 to obtain the three-dimensional modeling of the blades.
In another aspect, the present invention provides a computer-readable storage medium, in which a computer program is stored, and when the computer program is executed by a processor, the method for designing a turbine rear casing rectifying support plate with a wide-operating-condition backward shielding according to the present invention can be implemented.
The computer equipment can adopt a notebook computer, a desktop computer or a workstation.
The processor may be a Central Processing Unit (CPU), graphics Processing Unit (GPU), digital Signal Processor (DSP), application Specific Integrated Circuit (ASIC), or an off-the-shelf programmable gate array (FPGA).
The memory of the invention can be an internal storage unit of a notebook computer, a desktop computer or a workstation, such as a memory and a hard disk; external memory units such as removable hard disks, flash memory cards may also be used.
Computer-readable storage media may include computer storage media and communication media. Computer storage media includes volatile and nonvolatile, removable and non-removable media implemented in any method or technology for storage of information such as computer readable instructions, data structures, program modules or other data. The computer-readable storage medium may include: a Read Only Memory (ROM), a Random Access Memory (RAM), a Solid State Drive (SSD), or an optical disc. The Random Access Memory may include a resistive Random Access Memory (ReRAM) and a Dynamic Random Access Memory (DRAM).
The above description is only for the purpose of creating a preferred embodiment of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can substitute or change the technical solution and the inventive concept of the present invention within the technical scope of the present invention.
Claims (5)
1. A method for designing a turbine rear casing rectifying support plate with wide working condition backward shielding is characterized by comprising the following steps:
s1, selecting characteristic sections with different leaf heights to respectively shape; aiming at the different characteristic cross sections, the different characteristic cross sections respectively comprise a turning section and an axial extension section; wherein, the first and the second end of the pipe are connected with each other,the turning section is controlled by 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid distance and shielding width, and the axial extension section is controlled by the axial extension length; wherein the inlet geometry angle is determined based on the inlet airflow angle at the local blade height to reduce the flow separation at the inlet; the outlet geometry angle is set to 90 degrees to achieve axial venting; the chord length is adjusted according to the size limitation of the axial structure and the flow turning angle; the grid distance is determined according to the number of blades of the rectifying support plate (2) and the pitch diameter of the local blade height; by using the width of the shadedThe backward shielding rate of the final-stage low-pressure turbine part is accurately controlled, and the full shielding design is realized; by axial extensionlTo determine the length of the axial extension; width of shieldingdSpecifically, the method comprises the following steps: the axial projection coincidence distance of the front edge of the current rectifying support plate (2) and the tail edge of the adjacent rectifying support plate (2) on the cross section of the outlet of the engine is equal to the axial projection coincidence distance of the front edge of the current rectifying support plate and the tail edge of the adjacent rectifying support plate (2) on the cross section of the outlet of the engine, and if the front edge of the current rectifying support plate and the tail edge of the adjacent rectifying support plate are coincident with each other, the front edge of the current rectifying support plate and the tail edge of the adjacent rectifying support plate are coincident with each otherdIs positive; if the two are not coincident, thendIs negative whendWhen the number is 0, the rectification support plate (2) just blocks the backward observation sight from the tail of the engine, and the high-temperature low-pressure turbine part is shielded from the front and the back;dthe observation sight angle shielded by the rectification support plate (2) is increased along with the increase of the angle;
s2, respectively obtaining a turning section pressure surface molded line and a turning section suction surface molded line which meet 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, a chord length, a grid pitch and a shielding width by using a secondary deflection curve method; the molded line is integrally divided into a front edge section, a middle section and a tail edge section by a secondary deflection curve method, wherein the front edge section is a 1-section Bezier spline curve, and deflection from the near-axial direction to the circumferential direction is realized; the tail edge section is a 1-section Bezier spline curve, and deflection from the circumferential direction to the axial direction is realized; the middle section is a Bezier spline curve towards 1 section or 2 sections and is connected with the front edge section and the tail edge section;
s3, obtaining a pressure surface molded line and a suction surface molded line of the axial extension section through setting the axial extension length behind the turning section, and respectively combining the pressure surface molded line and the suction surface molded line of the turning section and the axial extension section to obtain a complete geometric molded line of a pressure surface and a suction surface of the blade;
s4, determining the diameters of the front edge and the tail edge of the blade according to the structural requirements, and modeling the front edge and the tail edge of the blade by utilizing an ellipse and an arc respectively so as to obtain initial primitive blade profiles with different characteristic sections; the blade root adopts a C-shaped flow channel outline, and the blade top adopts an S-shaped flow channel outline;
s5, fitting complete geometric molded lines of the pressure surface and the suction surface of the blade obtained in the S3 respectively by using a non-uniform B-spline curve; meanwhile, adding a plurality of control points at different axial positions, adjusting and iterating the molded lines of the blade, and modeling the front edge and the tail edge of the blade by combining the obtained molded lines to obtain an elementary blade profile meeting the pneumatic performance requirement;
and S6, determining a blade stacking form, setting circumferential and axial stacking parameters, and combining the blade stacking form and the element blade profile obtained in the S5 to obtain the three-dimensional shape of the blade.
2. The method for designing the turbine rear casing rectifying support plate with the wide-working-condition backward shielding function according to claim 1, wherein in S1, at least 3 characteristic element sections with different blade heights are selected to be respectively shaped.
3. A design system of a turbine rear casing rectification support plate with wide working condition backward shielding is characterized by being used for realizing the design method of claim 1 or 2, and comprising a design element determining module, a molded line design module, an initial element blade profile obtaining module, an element blade profile updating module and a blade three-dimensional modeling design module;
the design element determining module is used for selecting characteristic sections with different blade heights to respectively shape; aiming at the different characteristic cross sections, the different characteristic cross sections respectively comprise a turning section and an axial extension section; the turning section is controlled by 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid distance and shielding width, and the axial extension section is controlled by the axial extension length;
the molded line design module firstly utilizes a secondary deflection curve method to respectively obtain a turning section pressure surface molded line and a turning section suction surface molded line which meet 5 geometric parameters including an inlet geometric angle, an outlet geometric angle, chord length, grid spacing and shielding width; then obtaining a pressure surface molded line and a suction surface molded line of an axial extension section by giving an axial extension length behind the turning section, and respectively combining the pressure surface molded line and the suction surface molded line of the turning section and the axial extension section to obtain a complete geometric molded line of a pressure surface and a suction surface of the blade;
the initial element blade profile acquisition module is used for determining the diameters of the front edge and the tail edge of the blade according to the structural requirements, and modeling the front edge and the tail edge of the blade by utilizing an ellipse and an arc respectively so as to obtain initial element blade profiles with different characteristic sections;
the element blade profile updating module is used for respectively fitting complete geometric molded lines of the pressure surface and the suction surface of the obtained blade; meanwhile, adding a plurality of control points at different axial positions, adjusting and iterating the molded lines of the blade, and modeling the front edge and the tail edge of the blade by combining the obtained molded lines to obtain an elementary blade profile meeting the pneumatic performance requirement;
the blade three-dimensional modeling design module is used for determining a blade stacking form, setting circumferential and axial stacking parameters and combining the blade stacking form and the obtained element blade profile to obtain the three-dimensional modeling of the blade.
4. A computer device, comprising a processor and a memory, wherein the memory is used for storing a computer executable program, the processor reads the computer executable program from the memory and executes the computer executable program, and the processor can realize the design method of the turbine rear casing fairing plate of the wide-working-condition rear shielding in claim 1 or 2 when executing the computer executable program.
5. A computer-readable storage medium, wherein the computer-readable storage medium has a computer program stored therein, and the computer program, when executed by a processor, is capable of implementing the wide-regime aft-shielding aft-case fairing design method as claimed in claim 1 or 2.
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Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH07253001A (en) * | 1994-03-16 | 1995-10-03 | Mitsubishi Heavy Ind Ltd | Integral shroud moving blade |
| EP1259711A1 (en) * | 2000-02-17 | 2002-11-27 | ALSTOM Power N.V. | Aerofoil for an axial flow turbomachine |
| JP2012072706A (en) * | 2010-09-29 | 2012-04-12 | Hitachi Ltd | Method for modifying gas turbine device |
| CN102587997A (en) * | 2011-01-13 | 2012-07-18 | 阿尔斯通技术有限公司 | Vane for an axial flow turbomachine and corresponding turbomachine |
| JP2015068264A (en) * | 2013-09-30 | 2015-04-13 | 株式会社Ihi | Reflected electromagnetic wave attenuation structure for aircraft-engine |
| CN113217461A (en) * | 2021-05-12 | 2021-08-06 | 中南大学 | Blade, molding method and manufacturing method thereof and air compressor |
| CN113775436A (en) * | 2021-08-16 | 2021-12-10 | 中国航发贵阳发动机设计研究所 | Stealthy whirl mixing arrangement |
| CN113868793A (en) * | 2021-09-18 | 2021-12-31 | 西安热工研究院有限公司 | Novel movable blade modeling method for energy-saving reconstruction of movable blade adjustable axial flow fan of power station |
| CN114542207A (en) * | 2022-02-22 | 2022-05-27 | 中国航发沈阳发动机研究所 | Design method for outer surface modeling of turbine rear casing support plate |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2993020B1 (en) * | 2012-07-06 | 2016-03-18 | Snecma | TURBOMACHINE RECTIFIER WITH AUBES WITH IMPROVED PROFILE |
| GB2544554B (en) * | 2015-11-23 | 2018-07-04 | Rolls Royce Plc | Gas turbine engine |
-
2022
- 2022-06-09 CN CN202210648464.8A patent/CN115013089B/en active Active
Patent Citations (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH07253001A (en) * | 1994-03-16 | 1995-10-03 | Mitsubishi Heavy Ind Ltd | Integral shroud moving blade |
| EP1259711A1 (en) * | 2000-02-17 | 2002-11-27 | ALSTOM Power N.V. | Aerofoil for an axial flow turbomachine |
| US6709233B2 (en) * | 2000-02-17 | 2004-03-23 | Alstom Power N.V. | Aerofoil for an axial flow turbomachine |
| JP2012072706A (en) * | 2010-09-29 | 2012-04-12 | Hitachi Ltd | Method for modifying gas turbine device |
| CN102587997A (en) * | 2011-01-13 | 2012-07-18 | 阿尔斯通技术有限公司 | Vane for an axial flow turbomachine and corresponding turbomachine |
| JP2015068264A (en) * | 2013-09-30 | 2015-04-13 | 株式会社Ihi | Reflected electromagnetic wave attenuation structure for aircraft-engine |
| CN113217461A (en) * | 2021-05-12 | 2021-08-06 | 中南大学 | Blade, molding method and manufacturing method thereof and air compressor |
| CN113775436A (en) * | 2021-08-16 | 2021-12-10 | 中国航发贵阳发动机设计研究所 | Stealthy whirl mixing arrangement |
| CN113868793A (en) * | 2021-09-18 | 2021-12-31 | 西安热工研究院有限公司 | Novel movable blade modeling method for energy-saving reconstruction of movable blade adjustable axial flow fan of power station |
| CN114542207A (en) * | 2022-02-22 | 2022-05-27 | 中国航发沈阳发动机研究所 | Design method for outer surface modeling of turbine rear casing support plate |
Non-Patent Citations (6)
| Title |
|---|
| 世界航空动力技术的现状及发展动向;刘大响等;《北京航空航天大学学报》;20021030(第05期);全文 * |
| 低温透平膨胀机制动叶轮的优化设计;梁艳伟等;《低温与超导》;20210430(第4期);全文 * |
| 大扩张通道超声高载荷对转涡轮动叶三维设计方法研究;方祥军等;《航空学报》;20070125(第01期);全文 * |
| 超音尾喷流红外抑制方案的研究;杨旭等;《航空动力学报》;20020430(第02期);全文 * |
| 轴流式叶型的Bezier曲线设计方法;王毅等;《湘潭大学自然科学学报》;20161215(第04期);全文 * |
| 非轴对称端壁设计的高负荷涡轮气热性能研究进展;李军等;《航空发动机》;20180630(第03期);全文 * |
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