CN111594277B - Nozzle vane for supersonic turbine and design method thereof - Google Patents
Nozzle vane for supersonic turbine and design method thereof Download PDFInfo
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- CN111594277B CN111594277B CN202010477534.9A CN202010477534A CN111594277B CN 111594277 B CN111594277 B CN 111594277B CN 202010477534 A CN202010477534 A CN 202010477534A CN 111594277 B CN111594277 B CN 111594277B
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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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Abstract
The invention discloses a nozzle vane for a supersonic speed turbine and a design method thereof.A blade profile of the nozzle vane is composed of two groups of curves by taking a front edge and a tail edge as dividing points, wherein one group of the curves is a pressure surface curve, the other group of the curves is a suction surface curve, a salient point is arranged on the pressure surface curve and is a first salient point, a salient point is arranged on the suction surface curve and is a second salient point, the first salient point on one nozzle vane is aligned with the second salient point on the adjacent nozzle vane and is positioned at the same position vertical to the flow line to form a throat part of a nozzle flow channel, so that the two adjacent nozzle vanes form a convergent channel at the upstream of the throat part, the two adjacent nozzle vanes form a symmetrical divergent channel at the downstream of the throat part, and correspondingly, airflow is accelerated to sonic speed and then to supersonic speed from subsonic flow at the airflow inlet side. The supersonic flow expansion wave system can be distributed more smoothly, the supersonic flow loss is reduced, and the expansion capacity of the nozzle is increased.
Description
Technical Field
The invention relates to the technical field of low-temperature liquid rocket engine turbopumps, in particular to a nozzle blade for a supersonic turbine and a design method thereof.
Background
Both domestic and foreign high-thrust liquid rocket engines adopt pumping pressure type propellant supply systems to optimize rocket design schemes and improve carrying capacity.
The turbine is an important component of the engine system, and its efficiency affects the overall engine performance. The turbine for the liquid rocket engine can be divided into an independent turbine and a front turbine according to the difference of the engine circulating system, wherein the independent turbine is used for the gas generator circulating engine, and the turbine gas of the independent turbine is directly discharged into the atmosphere after working.
In order to improve engine performance, turbine gas flow must be reduced and specific power increased, and independent turbines often employ high pressure ratio supersonic turbine designs.
Supersonic turbines refer to nozzle outlet gas flow velocities in excess of sonic velocity, typically at mach numbers greater than 2. The gas flow inside the nozzle expands from a low speed to a supersonic speed through the convergent-divergent channel, which easily causes the flow loss related to the supersonic speed flow, and the design difficulty is large. The flow characteristics of the nozzle outlet combustion gases largely determine the output power and efficiency of the turbine blades. The design of the supersonic nozzle therefore has a decisive influence on the performance of the entire turbine.
Therefore, it is an urgent technical problem to be solved by those skilled in the art to provide a nozzle vane for a supersonic turbine, so as to make the supersonic flow expansion wave system distribution smoother, reduce the supersonic flow loss, and increase the nozzle expansion capacity.
Disclosure of Invention
In view of the above, an object of the present invention is to provide a nozzle vane for a supersonic turbine and a design method thereof, which can make the supersonic flow expansion wave system distribution smoother, reduce the supersonic flow loss, and increase the nozzle expansion capacity.
In order to achieve the purpose, the invention provides the following technical scheme:
a nozzle vane for supersonic turbine has a vane profile comprising two sets of curves, one set being a pressure surface curve and the other set being a suction surface curve, with the leading edge and the trailing edge as dividing points,
the pressure surface curve is provided with a salient point which is a first salient point,
the curve of the suction surface is provided with a convex point which is a second convex point,
the first salient point on one nozzle blade is aligned with the second salient point on the adjacent nozzle blade and is positioned at the same position vertical to the streamline to form the throat part of the nozzle runner,
two adjacent nozzle vanes form a convergent channel at the upstream of the throat part, namely the air flow inlet side, and two adjacent nozzle vanes form a symmetrical divergent channel at the downstream of the throat part, namely the air flow outlet side, so that the air flow is accelerated to the sonic speed and then to the supersonic speed again from the subsonic flow at the air flow inlet side.
Preferably, the leading edge is a circular arc.
Preferably, on the curve of the suction surface,
a first section of curve of the suction surface is formed from the front edge to the second salient point, the first section of curve of the suction surface is a single-section spline curve, the curvature of the curve is monotonously increased,
and a second section of curve of the suction surface is formed from the second salient point to the tail edge, the second section of curve of the suction surface is a single section of spline curve, and the curvature of the curve is monotonically reduced.
Preferably, the first section of the curve of the suction surface is connected with the second section of the curve of the suction surface by an arc.
Preferably, on the pressure surface curve,
a first section of curve of the pressure surface is formed from the front edge to the first salient point, the first section of curve of the pressure surface is a single-section spline curve, the curvature of the curve is monotonously increased,
and a second section of curve of the pressure surface is formed from the first salient point to the trailing edge, the second section of curve of the pressure surface is a single-section spline curve, and the curvature of the curve is monotonically reduced.
Preferably, the first section of the pressure surface curve is connected with the second section of the pressure surface curve in an arc manner.
The invention also provides a design method of the nozzle vane for the supersonic turbine, which comprises the following steps:
step 1) determining the pitch diameter of nozzle blades, the height of the nozzle blades, the number of the nozzle blades and the turbine air admission degree according to turbine thermal one-dimensional calculation parameters;
step 2), the thickness of the front edge of the nozzle blade is selected according to structural limitation, and the front edge of the nozzle blade is designed into a section of circular arc;
step 3), designing a first section of curve of the suction surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically increased from the front edge to a second salient point;
step 4), designing a second section of curve of the suction surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically reduced from the second salient point to the trailing edge of the nozzle blade;
step 5), arc connection is carried out between the first section of curve of the suction surface and the second section of curve of the suction surface;
step 6), designing a first section of curve of the pressure surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically increased and the curve extends from the front edge to a first salient point;
step 7), designing a second section of curve of the pressure surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically reduced from the first salient point to the trailing edge;
step 8) the first section of curve of the pressure surface is connected with the second section of curve of the pressure surface through an arc;
step 9), taking an arc from the tail edge;
step 10) determining a blade outlet installation angle according to a turbine one-dimensional thermodynamic calculation result, and then adjusting the positions of a first salient point on one nozzle blade and a second salient point on an adjacent nozzle blade to ensure that the two are aligned, wherein only one throat part is arranged in a nozzle flow passage along the gas flow direction, and the area of the throat part can ensure the gas flow requirement;
and 11) generating a three-dimensional blade model according to the two-dimensional blade profile, calculating a three-dimensional flow field, and finely adjusting the blade profile according to a calculation result to meet the design requirement.
Preferably, the method for designing the nozzle vane for the supersonic turbine further comprises the step 12) if one process cannot meet the design requirement, the step 3 to the step 11 form an iterative process until the design requirement is met.
Preferably, the radius of the arc at the trailing edge is the smallest value according to the strength of the blade.
Preferably, step 91) of tangent transition between the curves is also included after step 9).
The invention provides a nozzle vane for a supersonic turbine, which is characterized in that the vane profile of the nozzle vane takes a front edge and a tail edge as dividing points and consists of two groups of curves, wherein one group is a pressure surface curve, the other group is a suction surface curve,
a salient point is arranged on the curve of the pressure surface and is a first salient point,
the curve of the suction surface is provided with a convex point which is a second convex point,
the first salient point on one nozzle blade is aligned with the second salient point on the adjacent nozzle blade and is positioned at the same position vertical to the streamline to form the throat part of the nozzle runner,
thus, two adjacent nozzle vanes form a convergent channel upstream of the throat, i.e. the inlet side of the gas flow, and two adjacent nozzle vanes form a symmetrical divergent channel downstream of the throat, i.e. the outlet side of the gas flow, correspondingly so that the gas flow is accelerated from subsonic flow to sonic and then to supersonic flow again at the inlet side of the gas flow.
Compared with the traditional design method, the nozzle vane for the supersonic speed turbine provided by the invention has the following different points:
the expansion channel at the downstream of the throat part is formed by simultaneously expanding the curves of the pressure surface and the suction surface of the blade profile outwards, so that a symmetrical and uniform double-cluster expansion wave system can be formed in the flow direction, the flow loss related to expansion in the air flow acceleration process is reduced, and the flow efficiency of the nozzle is further improved; and the expansion capacity of the nozzle outlet can be increased, the Mach number of the outlet is increased, and the work capacity of the turbine is improved.
The more even nozzle outlet flow also makes the moving vane internal flow more high-efficient, further improves turbine efficiency.
Therefore, the distribution of the supersonic flow expansion wave system is more gentle, the supersonic flow loss is reduced, and the expansion capacity of the nozzle is increased.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
FIG. 1 is a schematic structural view of a supersonic turbine nozzle vane according to an embodiment of the present invention;
FIG. 2 is a schematic structural diagram of a curve of a suction surface provided in an embodiment of the present invention;
FIG. 3 is a schematic diagram of a pressure surface curve according to an embodiment of the present invention;
FIG. 4 is a schematic structural view of two supersonic turbine nozzle vanes in cooperation according to an embodiment of the present invention;
fig. 5 is a schematic structural view of a flow passage and a throat of a nozzle provided in an embodiment of the present invention.
In the above fig. 1-5:
the device comprises a first salient point 1, a second salient point 2, a first section of curve 3 of a suction surface, a second section of curve 4 of the suction surface, a first section of curve 5 of a pressure surface, a second section of curve 6 of the pressure surface, a front edge 7, a tail edge 8, a throat part 9, a nozzle flow channel 10, a curve 11 of the suction surface and a curve 12 of the pressure surface.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to fig. 1 to 5, fig. 1 is a schematic structural diagram of a nozzle vane for a supersonic turbine according to an embodiment of the present invention; FIG. 2 is a schematic structural diagram of a curve of a suction surface provided in an embodiment of the present invention; FIG. 3 is a schematic diagram of a pressure surface curve according to an embodiment of the present invention; FIG. 4 is a schematic structural view of two supersonic turbine nozzle vanes in cooperation according to an embodiment of the present invention; fig. 5 is a schematic structural view of a flow passage and a throat of a nozzle provided in an embodiment of the present invention.
The nozzle vane for supersonic turbine according to the embodiment of the present invention, as shown in fig. 1, has a blade profile that is formed by two sets of curves with a leading edge 7 and a trailing edge 8 as dividing points, wherein one set is a pressure surface curve 12, as shown in fig. 3, and the other set is a suction surface curve 11, as shown in fig. 2,
the pressure surface curve 12 is provided with a convex point which is a first convex point 1,
the suction surface curve 11 has a convex point, which is a second convex point 2,
the first salient point 1 on one nozzle vane is aligned with the second salient point 2 on the adjacent nozzle vane and is positioned at the same position perpendicular to the streamline, namely, the projections of the first salient point 1 and the second salient point 2 perpendicular to the streamline are superposed to form a throat part 9 of a nozzle flow channel 10, as shown in figures 4 and 5,
thus, two adjacent nozzle vanes form a convergent channel upstream of throat 9, i.e. the inlet side of the gas flow, and two adjacent nozzle vanes form a symmetrical divergent channel downstream of throat 9, i.e. the outlet side of the gas flow, correspondingly so that the gas flow is accelerated from subsonic flow to sonic flow and then to supersonic flow again at the inlet side of the gas flow.
Compared with the traditional design method, the nozzle vane for the supersonic speed turbine provided by the invention has the following different points:
the expansion channel at the downstream of the throat part 9 is formed by simultaneously expanding a pressure surface curve 121 and a suction surface curve 11 of the blade profile outwards, so that a symmetrical and uniform double-cluster expansion wave system can be formed in the flow direction, the flow loss related to expansion in the air flow acceleration process is reduced, and the flow efficiency of the nozzle is improved; and the expansion capacity of the nozzle outlet can be increased, the Mach number of the outlet is increased, and the work capacity of the turbine is improved.
The more even nozzle outlet flow also makes the moving vane internal flow more high-efficient, further improves turbine efficiency.
Therefore, the distribution of the supersonic flow expansion wave system is more smooth, the supersonic flow loss is reduced, and the expansion capacity of the nozzle is increased.
Specifically, the front edge 7 is a section of circular arc, and the radius of the circular arc is R1-R3.
Specifically, on the suction surface curve 11,
a first section of curve 3 of the suction surface is formed from the front edge 7 to the second salient point 2, the first section of curve 3 of the suction surface is a single-section spline curve, the curvature of the curve is monotonously increased,
the second section of the curve 4 of the suction surface is from the second salient point 2 to the tail edge 8, the second section of the curve 4 of the suction surface is a single-section spline curve, and the curvature of the curve is monotonously reduced.
Specifically, a first section curve 3 of the suction surface is connected with a second section curve 4 of the suction surface through an arc, and the radius of the arc is R1-R3.
Specifically, on the pressure surface curve 12,
a first section curve 5 of the pressure surface is formed from the front edge 7 to the first salient point 1, the first section curve 5 of the pressure surface is a single-section spline curve, the curvature of the curve is monotonously increased,
the first salient point 1 to the trailing edge 8 is a pressure surface second section curve 6, the pressure surface second section curve 6 is a single section spline curve, and the curvature of the curve is monotonously reduced.
Specifically, a first section of curve 5 of the pressure surface is connected with a second section of curve 6 of the pressure surface through an arc, and the radius of the arc is R1-R3.
The embodiment of the invention also provides a design method of the nozzle blade for the supersonic turbine, which comprises the following steps:
step 1) determining the pitch diameter of the nozzle vanes, the height of the nozzle vanes, the number of the nozzle vanes and the turbine air inlet degree according to one-dimensional calculation parameters of the turbine heat;
step 2), selecting the thickness of the front edge 7 of the nozzle blade according to structural limitation, designing the front edge 7 of the nozzle blade into a section of circular arc, and taking R1-R3 as the radius of the circular arc;
step 3), designing a first section of curve 3 of the suction surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonously increased from the front edge 7 to the second salient point 2;
step 4), designing a second section of curve 4 of the suction surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically reduced from the second salient point 2 to the tail edge 8 of the nozzle blade;
step 5), arc connection is carried out between the first section of curve 3 of the suction surface and the second section of curve 4 of the suction surface, and the radius of the arc is R1-R3;
step 6), designing a first section of curve 5 of the pressure surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonously increased from the front edge 7 to the first salient point 1;
step 7), designing a second section of curve 6 of the pressure surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically reduced from the first salient point 1 to the trailing edge 8;
step 8), arc connection is carried out between the first section curve 5 of the pressure surface and the second section curve 6 of the pressure surface, and the radius of the arc is R1-R3;
step 9), taking an arc from the tail edge 8;
step 10) determining a blade outlet installation angle according to a turbine one-dimensional thermodynamic calculation result, and then adjusting the positions of a first salient point 1 on one nozzle blade and a second salient point 2 on an adjacent nozzle blade to align the two, wherein only one throat 9 is arranged in a nozzle flow passage 10 along the gas flow direction, and the area of the throat 9 can meet the gas flow requirement;
and 11) generating a three-dimensional blade model according to the two-dimensional blade profile, calculating a three-dimensional flow field, and finely adjusting the blade profile according to a calculation result to meet design requirements.
Compared with the traditional design method, the design method of the nozzle vane for the supersonic speed turbine provided by the embodiment of the invention has the following different points:
the expansion channel at the downstream of the throat part is formed by simultaneously expanding a pressure surface curve 12 and a suction surface curve 11 of the blade profile outwards, so that a symmetrical and uniform double-cluster expansion wave system can be formed in the flow direction, the flow loss related to expansion in the air flow acceleration process is reduced, and the flow efficiency of the nozzle is improved; and the expansion capacity of the nozzle outlet can be increased, the Mach number of the outlet is increased, and the work capacity of the turbine is improved.
The more even nozzle outlet flow also makes the moving vane internal flow more high-efficient, further improves turbine efficiency.
Therefore, the distribution of the supersonic flow expansion wave system is more smooth, the supersonic flow loss is reduced, and the expansion capacity of the nozzle is increased.
In order to further optimize the above solution, the method for designing the nozzle vane for the supersonic turbine further includes step 12) if one process cannot meet the design requirement, the above steps 3 to 11 form an iterative process until the design requirement is met.
To further optimize the above solution, the radius of the arc of the trailing edge 8 is taken to be the minimum value according to the blade strength.
In order to further optimize the above solution, step 9) is followed by step 91) of making a tangential transition between the curves.
The technical scheme adopted by the embodiment of the invention for solving the technical problem is as follows:
according to the thermodynamic calculation result of the turbine, the three-dimensional flow field simulation is combined, and the nozzle blade profile with uniform and symmetrical expansion section profile change is obtained by increasing the design freedom of the nozzle blade profile, so that the supersonic flow expansion wave system is distributed more smoothly, the supersonic flow loss is reduced, and the expansion capacity of the nozzle is increased.
The nozzle vane for the supersonic turbine and the design method thereof provided by the embodiment of the invention can be applied to the vane design of the supersonic vane nozzle of the independent turbine of the liquid rocket engine, and have the beneficial effects that:
(1) The freedom degree of the design of the nozzle blade profile is increased, and the internal flow efficiency of the nozzle can be effectively improved;
(2) The Mach number of the airflow at the outlet of the nozzle is improved, and the working capacity of the unit flow of the gas of the movable blades is enhanced;
(3) The uniformity of the incoming flow at the inlet of the turbine movable blade is improved, and the efficiency of the movable blade is improved;
(4) By combining a design optimization method from thermodynamic calculation to three-dimensional simulation, the accuracy, reasonability and reliability of a design result are improved;
(5) The idea and the method for designing the nozzle blade profile are clear and simplified, and the design efficiency is improved.
The two-dimensional blade profile of the nozzle blade for the supersonic turbine according to the embodiment of the present invention is shown in fig. 1.
The blade profile is composed of two groups of curves with the leading edge 7 and the trailing edge 8 as division points in fig. 1, one group is called a pressure surface curve 12, and the other group is called a suction surface curve 11.
The pressure surface curve 12 and the suction surface curve 11 of the nozzle vane for the supersonic turbine provided by the embodiment of the invention are respectively provided with a convex point, namely the first convex point 1 and the second convex point 2, the positions of the two convex points are adjusted to enable the convex points of two adjacent vanes to be positioned at the same position vertical to a streamline, so that the throat part 9 of the nozzle runner 10 is formed, the two adjacent vanes form a convergent channel at the upstream (namely the airflow inlet side) of the throat part 9, the vane profile forms a symmetrical divergent channel at the downstream (namely the airflow outlet side) of the throat part 9, and correspondingly, the airflow is accelerated to the sonic speed from the subsonic flow of the inlet and then is accelerated to the supersonic flow.
The conventional supersonic nozzle blade profile generally has only one convex point on the pressure surface curve, and the suction surface curve is a straight line or a curve with monotonous change of curvature. The flow expansion acceleration of the downstream of the throat part only depends on the unidirectional area expansion of the pressure surface to form a single cluster expansion wave system.
Compared with the traditional design method, the nozzle vane for the supersonic turbine and the design method thereof provided by the embodiment of the invention have the different points that the expanding channel at the downstream of the throat part 9 is formed by simultaneously expanding the pressure surface curve 12 and the suction surface curve 11 of the vane profile outwards, so that a symmetrical and uniform double-cluster expansion wave system can be formed in the flow direction, the flow loss related to expansion in the gas flow acceleration process is reduced, and the flow efficiency of the nozzle is further improved; and the expansion capacity of the outlet of the nozzle can be increased, the Mach number of the outlet is increased, and the working capacity of the turbine is improved. The more even nozzle outlet flow also makes the inside flow of movable vane more high-efficient, further improves turbine efficiency.
The embodiment of the invention provides a design method of a nozzle blade for a supersonic speed turbine, which comprises the following concrete steps:
(1) Determining the pitch diameter of the turbine blades, the height of the nozzle blades, the number of the blades and the air inlet degree of the turbine according to the thermal one-dimensional calculation parameters of the turbine;
(2) The thickness of the front edge 7 of the blade can be selected according to structural limitation, the front edge 7 of the blade is designed into a section of circular arc under the usually appropriate condition, and the radius of the circular arc is between R1 and R3;
(3) The first section of curve 3 of the suction surface of the blade is designed into a single section of spline curve, and the curvature of the curve is monotonously increased;
(4) The second section of curve 4 of the suction surface of the blade is designed into a single section of spline curve, and the curvature of the curve is monotonously reduced;
(5) The first section of curve 3 of the suction surface of the blade is connected with the second section of curve 4 of the suction surface by an arc, and the radius of the arc is R1-R3;
(6) The first section of curve 5 of the pressure surface of the blade is designed into a single section of spline curve, and the curvature of the curve is monotonously increased;
(7) The second section of curve 6 of the pressure surface of the blade is designed into a single section of spline curve, and the curvature of the curve is monotonously reduced;
(8) The first section of curve 5 of the pressure surface of the blade is connected with the second section of curve 6 of the pressure surface through an arc, and the radius of the arc is R1-R3;
(9) The tail edge 8 of the blade is in a circular arc shape, and the radius of the circular arc-shaped tail edge 8 is as small as possible according to the strength of the blade;
(10) The curves of the first section of the suction surface 3, the second section of the suction surface 4, the first section of the pressure surface 5 and the second section of the pressure surface 6 are in tangent transition;
(11) Determining a blade outlet installation angle according to a turbine one-dimensional thermodynamic calculation result, and then adjusting the positions of a pressure surface salient point and a suction surface salient point to ensure that a channel between blades has only one throat 9 along the gas flow direction, and the area of the throat 9 can ensure the gas flow requirement;
(12) Generating a three-dimensional blade model according to the two-dimensional blade profile, calculating a three-dimensional flow field, and finely adjusting the blade profile according to a calculation result to meet design requirements of turbine flow, outlet Mach number, airflow angle and the like;
(13) The steps 3 to 12 form an iterative process until the design requirements are met.
The previous description of the disclosed embodiments is provided to enable any person skilled in the art to make or use the present invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the invention. Thus, the present invention is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.
Claims (10)
1. A nozzle vane for supersonic turbine is characterized in that the vane profile of the nozzle vane is composed of two groups of curves with the leading edge and the trailing edge as dividing points, wherein one group is a pressure surface curve, the other group is a suction surface curve,
the pressure surface curve is provided with a salient point which is a first salient point,
the curve of the suction surface is provided with a convex point which is a second convex point,
the first salient point on one nozzle blade is aligned with the second salient point on the adjacent nozzle blade to form the throat part of the nozzle flow passage,
two adjacent nozzle vanes form a convergent channel upstream of the throat, i.e. the inlet side of the gas flow, and two adjacent nozzle vanes form a symmetrical divergent channel downstream of the throat, i.e. the outlet side of the gas flow, correspondingly such that the gas flow is accelerated from a subsonic flow to a sonic flow and then to a supersonic flow again at the inlet side of the gas flow.
2. The nozzle vane of claim 1, wherein the leading edge is a segment of a circular arc.
3. The supersonic turbine nozzle vane according to claim 1, wherein the suction surface curve is on a line,
a first section of curve of the suction surface is formed from the front edge to the second salient point, the first section of curve of the suction surface is a single-section spline curve, the curvature of the curve is monotonously increased,
and a second section of curve of the suction surface is formed from the second salient point to the tail edge, the second section of curve of the suction surface is a single section of spline curve, and the curvature of the curve is monotonically reduced.
4. The nozzle vane of claim 3, wherein the suction surface first segment curve and the suction surface second segment curve are connected by an arc of a circle.
5. The nozzle vane for a supersonic turbine according to claim 1, wherein said pressure surface curve is a curve of a pressure surface,
a first section of curve of the pressure surface is formed from the front edge to the first salient point, the first section of curve of the pressure surface is a single-section spline curve, the curvature of the curve is monotonously increased,
and a second section of curve of the pressure surface is formed from the first salient point to the trailing edge, the second section of curve of the pressure surface is a single-section spline curve, and the curvature of the curve is monotonically reduced.
6. The supersonic turbine nozzle vane according to claim 5, wherein said pressure surface first segment curve and said pressure surface second segment curve are connected by an arc of a circle.
7. A method of designing a nozzle vane for a supersonic turbine, comprising:
step 1) determining the pitch diameter of the nozzle vanes, the height of the nozzle vanes, the number of the nozzle vanes and the turbine air inlet degree according to one-dimensional calculation parameters of the turbine heat;
step 2) selecting the thickness of the front edge of the nozzle blade according to structural limitation, wherein the front edge of the nozzle blade is designed into a section of circular arc;
step 3), designing a first section of curve of the suction surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically increased from the front edge to a second salient point;
step 4), designing a second section of curve of the suction surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically reduced from the second salient point to the tail edge of the nozzle blade;
step 5), arc connection is carried out between the first section of curve of the suction surface and the second section of curve of the suction surface;
step 6), designing a first section of curve of the pressure surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically increased and the curve extends from the front edge to a first salient point;
step 7), designing a second section of curve of the pressure surface of the nozzle blade into a single section of spline curve, wherein the curvature of the curve is monotonically reduced from the first salient point to the trailing edge;
step 8) the first section of curve of the pressure surface is connected with the second section of curve of the pressure surface through an arc;
step 9), taking an arc from the tail edge;
step 10) determining a blade outlet installation angle according to a turbine one-dimensional thermodynamic calculation result, and then adjusting the positions of a first salient point on one nozzle blade and a second salient point on an adjacent nozzle blade to align the two, wherein only one throat part is arranged in a nozzle flow passage along the gas flow direction, and the throat part area can ensure the gas flow requirement;
and 11) generating a three-dimensional blade model according to the two-dimensional blade profile, calculating a three-dimensional flow field, and finely adjusting the blade profile according to a calculation result to meet design requirements.
8. The method of claim 7, further comprising the step 12) if one of the processes fails to satisfy the design requirement, the steps 3) to 11) forming an iterative process until the design requirement is satisfied.
9. The method of claim 7, wherein the radius of the trailing edge is rounded to a minimum based on blade strength.
10. The method of designing a nozzle vane for a supersonic turbine according to claim 7, further comprising, after said step 9), the steps of: the curves are in tangent transition.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010477534.9A CN111594277B (en) | 2020-05-29 | 2020-05-29 | Nozzle vane for supersonic turbine and design method thereof |
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| CN202010477534.9A CN111594277B (en) | 2020-05-29 | 2020-05-29 | Nozzle vane for supersonic turbine and design method thereof |
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