CN114991878A - Turbine pressure side half-split cooling seal beard trailing edge blade and forming method - Google Patents
Turbine pressure side half-split cooling seal beard trailing edge blade and forming method Download PDFInfo
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- CN114991878A CN114991878A CN202210598550.2A CN202210598550A CN114991878A CN 114991878 A CN114991878 A CN 114991878A CN 202210598550 A CN202210598550 A CN 202210598550A CN 114991878 A CN114991878 A CN 114991878A
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
- F01D5/146—Shape, i.e. outer, aerodynamic form of blades with tandem configuration, split blades or slotted blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/70—Shape
- F05D2250/71—Shape curved
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
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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
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a turbine pressure side half-split cooling seal beard trailing edge blade and a forming method, and belongs to the technical field of passive flow control of aircraft power parts. The problem of current aeroengine's high-pressure turbine blade can't reduce the trailing edge loss and improve the regional cooling efficiency of trailing edge when guaranteeing trailing edge intensity is solved. The blade adopts the high pressure turbine movable vane profile of taking pressure side half split slit cooling, and lip and the trailing edge on the blade all adopt the seal beard structure, the pressure side and the afterbody of lip all present sinusoidal undulation along the leaf height direction, the pressure side and the suction side and the trailing edge point of trailing edge all present sinusoidal undulation along the leaf height direction. The cooling blade is mainly applied to pressure side half-split slot cooling blades in aeroengines and gas turbine turbines.
Description
Technical Field
The invention belongs to the technical field of passive flow control of aircraft power parts, and particularly relates to a turbine pressure side half-split cooling seal beard trailing edge blade and a forming method thereof.
Background
Improving the efficiency of aircraft engines is an important goal of modern aircraft engine design. Increasing efficiency can start with both increasing inlet temperature before the turbine and reducing losses. Increasing the inlet temperature before the turbine requires higher cooling performance of the turbine blades, while reducing losses requires the introduction of more advanced flow control measures. Turbine blade trailing edge shed vortices are a significant source of loss, with trailing edge losses accounting for over 1/3 in blade profile losses. Furthermore, the trailing edge region of the turbine blade is very difficult. Due to the thin construction of the turbine trailing edge, it is difficult to arrange internal cooling, typically by film cooling. The layout of turbine trailing edge cooling typically employs pressure side half-slit cooling, full-slit cooling, and trailing edge film hole cooling. The introduction of cold air will also bring a new source of losses. Moreover, the interaction of the cold air with the main flow will also seriously affect the film cooling efficiency. With the ever increasing demands placed on aircraft engine performance, there is a need to enter more advanced flow control methods for reducing turbine blade trailing edge losses and improving trailing edge cooling performance.
Several techniques have been employed to reduce the trailing edge losses of turbine blades and to improve the trailing edge cooling performance. El-Gendi applies equidistant micropores to the turbine blade trailing edge for increasing substrate pressure and reducing trailing edge losses. The results show a 0.7% increase in trailing edge substrate pressure and a 3% increase in overall loss. However, such equidistant micropores have the disadvantage of a considerable weakening of the strength of the trailing edge of the turbine blade. The elliptical trailing edge can delay the separation of the boundary layer, weaken the falling strength of the trailing edge and further reduce the loss of the trailing edge. However, an elliptical trailing edge does not completely eliminate shed vortices, so there is still room for further reduction in the loss at the trailing edge. In terms of active flow control, Bernardini et al employs pulsating air blowing to suppress trailing edge shedding vortices. This pulsating blowing pattern can alter the shedding pattern of the vortex. However, the use of pulsating blows introduces additional losses. In terms of improving the trailing edge cooling performance, the results of Effendy et al indicate that reducing the thickness of the lip portion can significantly improve the film cooling efficiency of the pressure side half-split cooling. However, reducing the thickness of the lip still reduces the strength of the trailing edge of the blade. The adoption of different lip structures can improve the air film cooling efficiency of the surface of the split seam to different degrees. However, these structures also weaken the lip to varying degrees. The current active and passive control methods have the characteristics of weakening the trailing edge strength and introducing additional losses in terms of control of the trailing edge losses and cooling performance, and therefore more effective flow control means are sought for controlling the flow in the trailing edge region.
The harbour seal beard structure was found to inhibit karman vortex street behind the cylinder. Shed vortices exist in aircraft engine components similar to behind cylinders, and so the beard structure of harps has also been applied to some components of aircraft engines. Prasad and Ricklick applied seal beard structures to the bypass flow columns of the internal cooling channels of turbine blades. Studies have shown that a streaming column with seal beard structure can reduce pressure loss in the cooling channel. Shyam et al applied a seal beard structure to the blade profile of the entire turbine blade. The seal beard leaf profile is insensitive to inlet airflow angle and can reduce the total pressure loss coefficient to a certain extent. Luo et al applied a seal beard structure to the trailing edge of an uncooled turbine blade. The blade profile with the seal beard structure at the tail edge can obviously reduce the energy loss coefficient of the turbine guide blade cascade and the turbine stage, and improve the efficiency of the turbine stage.
For high pressure turbine blades of aircraft engines, the trailing edge needs to be cooled due to the high inlet temperature before the turbine. Pressure side half-split slot cooling is a common form of trailing edge cooling. For such cooled airfoil, there is still a need for a flow control method that reduces trailing edge losses and improves cooling efficiency in the trailing edge region while maintaining trailing edge strength.
Disclosure of Invention
In view of the above, the present invention is directed to a turbine pressure side half-slit cooling seal beard trailing edge blade and a forming method thereof, so as to solve the problem that the high-pressure turbine blade of the existing aircraft engine cannot ensure the trailing edge strength, reduce the trailing edge loss, and improve the cooling efficiency of the trailing edge region.
In order to achieve the purpose, the invention adopts the following technical scheme: the utility model provides a half slit cooling seal beard trailing edge blade of turbine pressure side, the blade adopts the high-pressure turbine movable vane profile of taking the half slit cooling of pressure side, and lip and the trailing edge on the blade all adopt seal beard structure, the pressure side and the afterbody of lip all present sinusoidal undulation along the leaf height direction, the pressure side and the suction side and the trailing edge point of trailing edge all present sinusoidal undulation along the leaf height direction.
Further, the pressure side and tail undulating sinusoids of the lip are out of phase by π.
Further, the sinusoids undulating on the suction side and the pressure side of the trailing edge are in phase.
Further, the suction side and the pressure side of the trailing edge are out of phase with the undulating sinusoid at the trailing edge point by π.
Further, the cool air passage of the vane is provided with two rows of circumferential flow ribs.
The invention also provides a lip design method for cooling the seal beard trailing edge blade on the pressure side of the turbine by using the half split slit, wherein the lip is controlled by two control points, namely a control point I and a control point II, the control point II is positioned at the tail end of the lip, and the control point I is positioned at the upstream d of the tail end of the lip l A is d l The first cut-off point and the second cut-off point are respectively positioned at the upstream 1.5d of the tail end of the lip part according to the thickness of the lip part l And d l The blade profile is used for dividing the original blade profile and the beard lip, the blade profile before the cut-off point is kept unchanged, the cut-off point I is connected with the control point I through a three-order Bezier curve, the control point I, the control point II and the cut-off point II are connected through an elliptic curve, the control point I and the control point II are in sinusoidal change along the spanwise direction to form a control line I and a control line II, and the blade profiles are stacked in the spanwise direction to form a three-dimensional blade.
The control equation of the first control line is as follows:
the control equation of the second control line is as follows:
the invention also relates toThe tail edge is controlled by three control points, namely a control point three, a control line point four and a control point five, the control point four is positioned at the tail end of the tail edge, and the interception point three and the interception point four are positioned at the upstream of the tail edge point by 1.5d t D is treatment of t Representing the thickness of the trailing edge, control points three and five being located upstream of the trailing edge point d t And the cutoff point and the control points are connected by adopting a three-order Bezier curve, the control points are connected by adopting an elliptic curve, the control point III, the control line IV and the control point V are in sinusoidal change along the spanwise direction to form a control line III, a control line IV and a control line V, and the blade profiles are stacked in the spanwise direction to form the three-dimensional blade.
The control equations of the third control line and the fifth control line are as follows:
the control equation of the control line four is as follows:
the invention also provides a method for forming the turbine pressure side half-slit cooling seal beard trailing edge blade, the blade is cast by a lost wax method, a ceramic core is used as an inner core for lost wax casting, and the blade is molded by casting, and the blade passing through X-ray flaw detection is polished and refined to obtain an accurate beard trailing edge profile.
Compared with the prior art, the invention has the beneficial effects that: the invention is mainly applied to the pressure side half-slit cooling blade in the turbine of an aeroengine and a gas turbine, and the seal beard structure is applied to the trailing edge of the turbine blade to reduce the trailing edge loss by utilizing the characteristic that the seal beard structure can effectively inhibit the falling vortex, thereby improving the efficiency of the engine. In addition, the structure can inhibit the mixing of cold air and main flow, thereby improving the air film efficiency of the trailing edge of the turbine blade, prolonging the service life of the turbine blade and improving the safety of an engine.
The invention further provides a design method and a forming method of the pressure side half-slit cooling turbine blade, and the seal beard structure is applied to the lip and the tail edge of the pressure side half-slit cooling turbine blade and used for reducing the loss of the tail edge and improving the air film cooling efficiency of the tail edge. Not only the aerodynamic loss is reduced but also the air film cooling efficiency is improved, and the energy loss coefficients of the prototype and the seal beard trailing edge blade profile are shown in table 1. The energy loss of the blade at the trailing edge of the beard is reduced by 0.9 percent compared with the prototype. Shed vortices behind the prototype lip and trailing edge were completely suppressed as shown in fig. 9 and 10. The method can reduce the blade profile loss by inhibiting the trailing edge shedding vortex, reduce the impact of the wake on the downstream blades and obviously improve the aerodynamic performance of the turbine. The efficiency of the slot surface adiabatic film cooling of the prototype and the seal beard trailing edge blade is shown in fig. 11 and 12. The falling vortex of the lip is restrained, so that the air film cooling efficiency of the slit surface of the leaf profile of the tail edge of the harp seal beard close to the lip area and the tail edge part area is obviously improved.
TABLE 1 prototype and coefficients of energy loss of the trailing edge blade of beard
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention, illustrate embodiments of the invention and together with the description serve to explain the invention and do not constitute a limitation of the invention. In the drawings:
FIG. 1 is a schematic structural view of a turbine pressure side half-slit cooling seal beard trailing edge blade according to the present invention;
FIG. 2 is a schematic view of the trailing edge pressure side configuration of the present invention;
FIG. 3 is a schematic view of the trailing edge suction side configuration of the present invention;
FIG. 4 is a schematic view of the lip control point position according to the present invention;
FIG. 5 is a schematic diagram of the location of the trailing edge control point according to the present invention;
FIG. 6 is a schematic view of a pressure side control line according to the present invention;
FIG. 7 is a schematic view of a suction side control line according to the present invention;
FIG. 8 is a graph of a sinusoidal control curve for the control line of the present invention;
FIG. 9 is a diagram of a prototype blade trail according to the present invention;
FIG. 10 is a trailing image of a trailing blade of a beard according to the present invention;
FIG. 11 is a graph of the cooling efficiency of the thermal insulation film on the surface of the cleft of the prototype blade according to the present invention;
FIG. 12 is a diagram of the cooling efficiency of the thermal insulation film on the surface of the cleft of the blade at the trailing edge of the beard according to the present invention;
FIG. 13 is a schematic view of the original airfoil configuration of a turbine bucket according to the present invention;
FIG. 14 is a schematic cross-sectional view A-A of FIG. 13 according to the present invention;
fig. 15 is a schematic view of an embodiment of a half-slit cooling seal beard trailing edge blade on the pressure side of an aviation turbine in a cascade flow domain.
1-lip, 2-trailing edge.
Detailed Description
The technical solution in the embodiments of the present invention will be clearly and completely explained below with reference to the drawings in the embodiments of the present invention. It should be noted that, in the present invention, the embodiments and features of the embodiments may be combined with each other without conflict, and the described embodiments are only a part of the embodiments of the present invention, not all of the embodiments.
Referring to fig. 1-15 to illustrate the embodiment, the blade profile of an aviation turbine pressure side half-slit cooled seal beard trailing edge blade is shown in fig. 1, the base blade profile of the blade is a high-pressure turbine moving blade profile with pressure side half-slit cooling, two rows of circumferential flow ribs are arranged in a cooling air channel of the blade for enhancing heat exchange and increasing the strength of the blade, a seal beard structure is adopted for a lip 1 and a trailing edge 2 on the blade, a detailed structure of a pressure side of a trailing edge part is shown in fig. 2, and a detailed structure of a suction side of the trailing edge part is shown in fig. 3. The pressure side and the tail of the lip 1 both present sinusoidal fluctuation along the blade height direction, the sinusoidal curve of the pressure side and the tail fluctuation of the lip 1 has a phase difference of pi, and the pressure side, the suction side and the tail point of the tail edge 2 both present sinusoidal fluctuation along the blade height direction. The phases of the fluctuant sine curves of the suction side and the pressure side of the trailing edge 2 are the same, and the phase difference between the suction side and the pressure side of the trailing edge 2 and the fluctuant sine curves at the trailing edge point is pi.
For a typical pressure side half-split cooled airfoil, the lip portion is generally square or rounded, while the trailing edge portion is generally rounded. This configuration creates shedding vortices behind both the lip and trailing edge. The shedding vortex generated by the lip can roll high-temperature fuel gas to the surface of the split seam, and simultaneously roll cooling gas away from the surface of the split seam. This action can exacerbate the intermingling of the cold gas and the main flow, which can reduce the film cooling efficiency of the split surfaces and also increase the aerodynamic losses. After the trailing edge, the alternately falling karman vortices are generated. This shedding vortex is the major source of trailing edge loss. And the adoption of the seal beard structure can enable the separation boundary layer at the lip and the tail edge to generate the span-wise migration, and obviously inhibit the formation of the shedding vortex behind the lip and the tail edge. The inhibition of the shedding vortex of the lip can improve the air film cooling efficiency of the surface of the split seam and reduce the aerodynamic loss to a certain extent. The suppression of trailing edge shedding vortices can significantly reduce aerodynamic losses.
In the embodiment, the lip design method for cooling seal beard trailing edge blades on the pressure side of the turbine through half split gaps is characterized in that the lip 1 is controlled by two control points, namely a control point I and a control point II, as shown in fig. 4, the control point II is positioned at the tail end of the lip, and the control point I is positioned at the upstream d of the tail end of the lip l A is d l The first cut-off point and the second cut-off point are respectively positioned at the upstream 1.5d of the tail end of the lip part according to the thickness of the lip part l And d l The blade profile is used for dividing the original blade profile and the beard lip, the blade profile before the cut-off point is kept unchanged, the cut-off point I is connected with the control point I by adopting a three-order Bessel curve, the control point I, the control point II and the cut-off point II are connected by adopting an elliptic curve, the control point I and the control point II are in sinusoidal change along the spanwise direction to form a control line I and a control line II, and the blade profile is divided into an original blade profile and a beard lipThe forms are stacked in the spanwise direction to form a three-dimensional blade as shown in fig. 6 and 7. The sinusoidal control curve is shown in fig. 8.
The control equation of the first control line is as follows:
the control equation of the second control line is as follows:
in the embodiment, the design method of the tail edge of the turbine pressure side half-split cooling seal beard tail edge blade is characterized in that the tail edge 2 is controlled by three control points, namely a control point three, a control line point four and a control point five, as shown in figure 5, the control point four is positioned at the tail end of the tail edge, and the interception point three and the interception point four are positioned at the upstream 1.5d of the tail edge point t A is d t Representing the thickness of the trailing edge, control points three and five being located upstream of the trailing edge point d t The cutoff points and the control points are connected by a third-order Bezier curve, the control points are connected by an elliptic curve, the control point III, the control line point IV and the control point V are in sinusoidal change along the spanwise direction to form a control line III, a control line IV and a control line V, and the blade profiles are stacked in the spanwise direction to form a three-dimensional blade, as shown in FIGS. 6 and 7. The sinusoidal control curve is shown in fig. 8.
The control equations of the third control line and the fifth control line are as follows:
the control equation of the control line four is as follows:
the embodiment is a method for forming a turbine pressure side half-slit cooling seal beard trailing edge blade.
The aviation turbine pressure side half-split slit cooling seal beard trailing edge blade is designed on the basis of an original blade profile. The original blade profile is a high pressure turbine bucket blade profile as shown in fig. 13 and 14. The parameters of the ribs in the cooling channels are shown in section a-a. The values corresponding to the labeled parameters in fig. 13 and 14 are given in table 2.
TABLE 2 turbine blade raw Profile parameters
A seal beard structure was applied to the trailing edge portion of the original blade profile, which was placed in an aviation turbine cascade, with one flow path element intercepted as shown in fig. 15. The mainstream exit mach number is in the range of high subsonic velocity for the flow characteristics of the high pressure turbine. Cooling channels are arranged inside the blade. The cold air flows out from the cooling channel with the ribs, covers the surface of the split seam and is mixed with the main stream of high-temperature fuel gas.
The embodiments of the invention disclosed above are intended merely to aid in the explanation of the invention. The examples are not intended to be exhaustive or to limit the invention to the precise forms disclosed. Many modifications and variations are possible in light of the above teaching. The embodiments were chosen and described in order to best explain the principles of the invention and the practical application, to thereby enable others skilled in the art to best understand the invention for and utilize the invention.
Claims (10)
1. The utility model provides a turbine pressure side is half splits seam cooling seal beard trailing edge blade which characterized in that: the blade adopts the high pressure turbine movable vane profile of taking pressure side half split slit cooling, and lip (1) and trailing edge (2) on the blade all adopt the seal beard structure, the pressure side and the afterbody of lip (1) all present sinusoidal undulation along the leaf height direction, the pressure side and the suction side and the trailing edge point of trailing edge (2) all present sinusoidal undulation along the leaf height direction.
2. The turbine pressure side half-split cooled seal beard trailing edge blade of claim 1, wherein: the pressure side and tail part of the lip (1) undulate sinusoids differing in phase by pi.
3. The turbine pressure side half-split cooled seal beard trailing edge blade of claim 1, wherein: the phases of the sinusoids of the undulation of the suction side and the pressure side of the trailing edge (2) are the same.
4. The turbine pressure side half-split slit cooled seal beard trailing edge blade of claim 1 or 3, wherein: the suction side and the pressure side of the trailing edge (2) differ in phase from the undulating sinusoid at the trailing edge point by pi.
5. The turbine pressure side half-split cooled seal beard trailing edge blade of claim 1, wherein: the cold air channel of the blade is provided with two rows of circumferential flow ribs.
6. The method for designing the lip of the turbine pressure side half-split cooling seal beard trailing edge blade according to claim 1, characterized in that: the lip (1) is controlled by two control points, namely a control point I and a control point II, wherein the control point II is positioned at the tail end of the lip, and the control point I is positioned at the upstream d of the tail end of the lip l A is d l The first cut-off point and the second cut-off point are respectively positioned at the upstream 1.5d of the tail end of the lip part according to the thickness of the lip part l And d l The blade profile before the truncation point is kept unchanged, the first truncation point is connected with the first control point by adopting a three-order Bessel curve, the first control point, the second control point and the second truncation point are connected by adopting an elliptic curve, and the first control point and the second control point are in sinusoidal change along the spanwise direction to form a first control line and a second control lineSecond, the profiles are stacked in the span-wise direction to form a three-dimensional blade.
7. The method for designing the lip of the turbine pressure side half-split slit cooled seal beard trailing edge blade according to claim 6, wherein the method comprises the following steps:
the control equation of the first control line is as follows:
the control equation of the second control line is as follows:
8. the method for designing the trailing edge of the turbine pressure side half-split cooling seal beard trailing edge blade according to claim 1, characterized in that: the tail edge (2) is controlled by three control points, namely a control point three, a control line point four and a control point five, wherein the control point four is positioned at the tail end of the tail edge, and the interception point three and the interception point four are positioned at the upstream 1.5d of the tail edge point t D is treatment of t Representing the thickness of the trailing edge, control points three and five being located upstream of the trailing edge point by d t And the cutoff points are connected with the control points by adopting a three-order Bessel curve, the control points are connected by adopting an elliptic curve, the control point III, the control line IV and the control point V are in sinusoidal change along the spanwise direction to form a control line III, a control line IV and a control line V, and the blade profiles are stacked in the spanwise direction to form the three-dimensional blade.
9. The method for designing the trailing edge of the turbine pressure side half-split cooling seal beard trailing edge blade according to claim 8, wherein the method comprises the following steps:
the control equations of the third control line and the fifth control line are as follows:
the control equation of the control line four is as follows:
10. the method for forming the turbine pressure side half-split slit cooled seal beard trailing edge blade according to claim 1, characterized in that: the blade is cast by a lost wax method, a ceramic core is used as an inner core for lost wax casting, and the blade is molded by casting, and is polished and refined by X-ray flaw detection to obtain an accurate beard trailing edge profile.
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| CN114991878B (en) | 2023-07-07 |
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