Interstage support preposed three-rotor mechanism, engine and blade profile optimization method

Abstract

The invention belongs to the field of aero-engines, and particularly relates to a three-rotor mechanism with front interstage supports, an engine and a blade profile optimization method. According to the three-rotor mechanism with the front interstage support, a turbine meridian inner end wall is arranged in a turbine meridian outer end wall to form an annular channel; a high-pressure turbine assembly, a low-pressure turbine assembly and a power turbine assembly are sequentially arranged in the annular channel from the inlet end to the outlet end of the annular channel. The low-pressure turbine assembly comprises interstage supporting guide vanes and working vanes, the interstage supporting guide vanes are arranged close to the high-pressure turbine assembly, and the working vanes are arranged close to the power turbine assembly. By integrally designing the low-pressure turbine guide vane and the interstage support, the preposition of the interstage support is realized, the coordination compactness of the three-rotor mechanism and the whole engine structure is improved, the axial span of a power turbine rotor fulcrum in a front output shaft form is effectively shortened, and the robustness of the power turbine rotor structure is enhanced.

Description

Technical Field

[0001] The invention belongs to the field of aeroengines, and in particular relates to a three-rotor mechanism with front interstage supports, an engine and a blade profile optimization method. Background Art

[0002] In order to achieve goals such as high engine power-to-weight ratio and good reliability, advanced high-power aviation gas turboshaft / propeller engines adopt a three-rotor configuration design concept, namely, it includes a high-pressure turbine, a low-pressure turbine and a power turbine. The low-pressure turbine, as a key component of the core engine, is located between the high-pressure turbine outlet and the power turbine inlet, and is mainly used to drive the low-pressure compressor. The quality of the low-pressure turbine design determines the overall design level of the engine, and its development level plays a vital role in the development of aviation engines.

[0003] In the existing design scheme, the low-pressure turbine adopts a conventional design, the interstage support is still placed downstream of the low-pressure turbine, the axial spacing of the low-pressure turbine is not fully utilized, and the axial span of the power turbine rotor fulcrum in the form of the front output shaft is increased, resulting in poor robustness of the power turbine; at the same time, the aerodynamic design of the low-pressure turbine is designed from a single perspective, without considering the coordinated compactness of the components and the overall structure, and the low-pressure turbine and the interstage support are not integrated in the design, resulting in a low power-to-weight ratio of the overall machine and increased manufacturing and processing costs.

[0004] In view of this, the present invention is proposed. Summary of the invention

[0005] In order to solve the technical problems existing in the prior art, the present invention provides a three-rotor mechanism, an engine and a blade optimization method with a pre-placed interstage support. The present invention realizes the pre-placement of the interstage support by integrating the low-pressure turbine guide vane and the interstage support, thereby improving the coordinated compactness of the three-rotor mechanism and the overall structure of the engine, effectively shortening the axial span of the power turbine rotor fulcrum in the form of a front output shaft, and enhancing the poor robustness of the power turbine rotor structure.

[0006] The present invention includes the following technical solutions:

[0007] A first aspect of the present invention provides a three-rotor mechanism with an interstage support in front, comprising a turbine meridian outer end wall, a turbine meridian inner end wall, a high-pressure turbine assembly, a low-pressure turbine assembly and a power turbine assembly; the turbine meridian inner end wall is arranged inside the turbine meridian outer end wall to form an annular channel; from the inlet end to the outlet end of the annular channel, the high-pressure turbine assembly, the low-pressure turbine assembly and the power turbine assembly are arranged in sequence in the annular channel;

[0008] The low-pressure turbine assembly includes an inter-stage support guide vane and a working blade. The inter-stage support guide vane is arranged close to the high-pressure turbine assembly, and the working blade is arranged close to the power turbine assembly.

[0009] Furthermore, the working blades are non-cooling shrouded blades.

[0010] Furthermore, an accessory passage is provided on the interstage support guide vane.

[0011] The second invention of the present invention provides a three-rotor engine with a front interstage support, including the three-rotor mechanism described above.

[0012] A third aspect of the present invention provides a method for optimizing the turbine blade profile of a three-rotor mechanism with a front interstage support, comprising the following steps:

[0013] Determine the structural features of the turbine meridian outer end wall and the turbine meridian inner end wall of the three-rotor mechanism as described in claims 1-3 according to the aerodynamic design constraints, and design the blade profiles of the interstage support guide vanes and the working blades according to the aerodynamic design constraints;

[0014] According to the structural features, the blade profiles of the interstage support guide vanes and the working blades, the matching relationship between the high-pressure turbine assembly and the low-pressure turbine assembly, and the cooling seal of the three-rotor mechanism, a three-dimensional simulation joint calculation is performed on the three-rotor mechanism, and the performance parameters of the preset state points of the three-rotor mechanism are obtained;

[0015] If the performance parameters have a margin, the blade profiles of the interstage support guide vanes and the working blades are obtained; if the performance parameters do not have a margin, the blade profiles of the interstage support guide vanes and the working blades are redesigned according to the aerodynamic design constraints of the three-rotor mechanism to carry out the three-dimensional simulation joint calculation until the obtained performance parameters have a margin.

[0016] Furthermore, the structural features are the expansion angle of the turbine meridian inner end wall and the expansion angle of the turbine meridian outer end wall.

[0017] Furthermore, the aerodynamic design constraints include a low-pressure turbine load coefficient, a low-pressure turbine flow coefficient, a low-pressure turbine reaction force, a low-pressure turbine expansion ratio and a low-pressure turbine efficiency.

[0018] Furthermore, the low-pressure turbine load factor is determined according to the low-pressure turbine rim power, the low-pressure turbine mid-diameter tangential speed, the inlet tangential speed of the low-pressure turbine working blades and the outlet tangential speed of the low-pressure turbine working blades.

[0019] Furthermore, the low-pressure turbine flow coefficient is determined according to the low-pressure turbine outlet axial velocity and the low-pressure turbine mid-diameter tangential velocity.

[0020] Furthermore, the low-pressure turbine reaction force is determined according to the low-pressure turbine mid-diameter tangential velocity, the inlet tangential velocity of the low-pressure turbine working blades, and the outlet tangential velocity of the low-pressure turbine working blades.

[0021] By adopting the above technical solution, the present invention has the following advantages:

[0022] 1. The present invention realizes the front placement of the interstage support by integrating the low-pressure turbine guide vane and the interstage support, thereby improving the coordinated compactness of the three-rotor mechanism and the overall structure of the engine, effectively shortening the axial span of the power turbine rotor fulcrum in the form of the front output shaft, and enhancing the poor robustness of the power turbine rotor structure.

[0023] 2. The working blades of the low-pressure turbine of the present invention are selected as non-cooled shrouded blades, which has the advantage of further ensuring the performance of the low-pressure turbine.

[0024] 3. The interstage supported front three-rotor mechanism of the present invention has the advantages of compact structure, convenient unit installation, high power-to-weight ratio, etc.

[0025] 4. The blade profile optimization method of the present invention can design and optimize the interstage support guide vanes and working blades of the low-pressure turbine of the interstage support front three-rotor mechanism, which has the advantage of further improving the performance of the low-pressure turbine.

[0026] Other features and advantages of the present invention will be described in the following description, and partly become apparent from the description, or understood by practicing the present invention. The purpose and other advantages of the present invention can be realized and obtained by the structures pointed out in the description and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0028] Figure 1 This is a schematic structural diagram of a three-rotor mechanism with a front interstage support in an embodiment of the present invention;

[0029] Figure 2 Schematic diagram of the mid-diameter velocity triangle of a low-pressure turbine in an embodiment of the present invention;

[0030] Figure 3 Schematic diagram of the structure of the thin blades of the interstage support guide vanes in an embodiment of the present invention;

[0031] Figure 4 Schematic diagram of the structure of the thick blades of the interstage support guide vanes in an embodiment of the present invention;

[0032] Figure 5 This is a schematic diagram of the structure of the working blade in an embodiment of the present invention;

[0033] In the figure: 1-turbine meridian outer end wall, 2-turbine meridian inner end wall, 3-high-pressure turbine assembly, 4-low-pressure turbine assembly, 41-interstage support guide vanes, 42-working blades, 5-power turbine assembly. DETAILED DESCRIPTION

[0034] The following description provides many different embodiments or examples for implementing different features of the present invention. The components and arrangements described in the following specific examples are only used to simplify the present invention and are only used as examples, not to limit the present invention.

[0035] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0036] In a first aspect, this embodiment provides a three-rotor mechanism with a front interstage support, such as Figure 1 As shown, it includes a turbine meridian outer end wall 1, a turbine meridian inner end wall 2, a high-pressure turbine assembly 3, a low-pressure turbine assembly 4 and a power turbine assembly 5; the turbine meridian inner end wall 2 is arranged in the turbine meridian outer end wall 1 to form an annular channel; from the inlet end to the outlet end of the annular channel, the high-pressure turbine assembly 3, the low-pressure turbine assembly 4 and the power turbine assembly 5 are arranged in sequence in the annular channel;

[0037] The low-pressure turbine assembly 4 includes an inter-stage support guide vane 41 and a working blade 42 . The inter-stage support guide vane 41 is arranged close to the high-pressure turbine assembly 3 , and the working blade 42 is arranged close to the power turbine assembly 5 .

[0038] Figure 1 In order to make it easier to clearly define the positions of the interstage support guide vanes 41 and the working blades 42 of the low-pressure turbine assembly 4, they correspond to the positions of the low-pressure turbine assembly 4 in the three-rotor mechanism. In the figure, only part of the structure of the three-rotor mechanism is shown, and the actual turbine meridian outer end wall 1 and the turbine meridian inner end wall 2 should be annular structures.

[0039] In some embodiments, the working blades 42 are non-cooled shrouded blades.

[0040] In some embodiments, Figure 1 As shown, an accessory passage is provided on the inter-stage support guide vane 41 .

[0041] In some embodiments, the inter-stage support guide vane 41 includes thick blades and thin blades, and the thick blades and the thin blades are evenly spaced.

[0042] In some embodiments, the number of the thick blades is 5 and the number of the thin blades is 15, and 3 thin blades are arranged between any two adjacent thick blades.

[0043] In some embodiments, the maximum cross-sectional circle of the thick blade is sufficient to pass through a pipe with a diameter of 20 mm.

[0044] The design of the interstage support guide vanes 41 can provide a high-performance flow channel for the mainstream gas on the one hand, and can provide structural support and oil and air pipelines on the other hand.

[0045] A second aspect of this embodiment provides a three-rotor engine with a front interstage support, including the three-rotor mechanism described above.

[0046] A third aspect of the present embodiment provides a method for optimizing a turbine blade profile of a three-rotor mechanism with a front interstage support, comprising the following steps:

[0047] Determine the structural features of the turbine meridian outer end wall 1 and the turbine meridian inner end wall 2 of the three-rotor mechanism as described in claims 1-3 according to the aerodynamic design constraints, and design the blade profiles of the interstage support guide vanes 41 and the working blades 42 according to the aerodynamic design constraints;

[0048] According to the structural features, the blade profiles of the interstage support guide vanes 41 and the working blades 42, the matching relationship between the high-pressure turbine assembly 3 and the low-pressure turbine assembly 4, and the cooling seal of the three-rotor mechanism, a three-dimensional simulation joint calculation is performed on the three-rotor mechanism, and the performance parameters of the preset state points of the three-rotor mechanism are obtained;

[0049] If the performance parameters satisfy the requirements with a margin, the blade profiles of the interstage support guide vanes 41 and the working blades 42 are obtained; if the performance parameters do not satisfy the requirements with a margin, the blade profiles of the interstage support guide vanes 41 and the working blades 42 are redesigned according to the aerodynamic design constraints of the three-rotor mechanism to carry out the three-dimensional simulation calculation until the obtained performance parameters satisfy the requirements with a margin.

[0050] Among them, the preset state points include maximum climb state, maximum cruise state, normal take-off state and maximum take-off state.

[0051] Among them, the performance parameters include maximum climb increase, maximum cruise increase, normal takeoff increase and maximum takeoff increase. The increase can include efficiency increase and power increase.

[0052] In some embodiments, the structural features are the divergence angle of the turbine meridian inner end wall 2 and the divergence angle of the turbine meridian outer end wall 1 .

[0053] In some embodiments, the aerodynamic design constraints include low-pressure turbine load coefficient, low-pressure turbine flow coefficient, low-pressure turbine reaction force, low-pressure turbine expansion ratio and low-pressure turbine efficiency. The blade profiles of the interstage support guide vanes 41 and the working blades 42 are designed according to the low-pressure turbine load coefficient and the low-pressure turbine flow coefficient, which has the advantage of improving the efficiency of the low-pressure turbine; the blade profiles of the interstage support guide vanes 41 and the working blades 42 are designed according to the low-pressure turbine reaction force, which ensures the turning angle of the blade exhaust flow and the outlet Mach number, and reduces the relative total temperature of the working blade 42 inlet, which has the advantage of improving the service life of the low-pressure turbine.

[0054] In some embodiments, the low-pressure turbine load factor is determined according to the low-pressure turbine rim power, the low-pressure turbine mid-diameter tangential speed, the inlet tangential speed of the low-pressure turbine blades 42 , and the outlet tangential speed of the low-pressure turbine blades 42 .

[0055] In some embodiments, the low-pressure turbine flow coefficient is determined according to the low-pressure turbine outlet axial velocity and the low-pressure turbine mid-diameter tangential velocity.

[0056] In some embodiments, the low-pressure turbine reaction force is determined according to the low-pressure turbine mid-diameter tangential velocity, the inlet tangential velocity of the low-pressure turbine blades 42 , and the outlet tangential velocity of the low-pressure turbine blades 42 .

[0057] The low-pressure turbine load factor μ is calculated by the formula get.

[0058] Among them, Figure 2 As shown, L u is the LP turbine wheel rim work, u is the LP turbine mid-diameter tangential velocity (reflecting the LP turbine wheel stress and life level), C 1u The tangential velocity of the low-pressure turbine blade 42, C 2u is the tangential speed of the low-pressure turbine.

[0059] The low-pressure turbine flow coefficient Φ is calculated by the formula get.

[0060] Among them, Figure 2 As shown, C 2a is the axial speed at the low-pressure turbine outlet.

[0061] Low pressure turbine motion reaction force Ω k (used to distribute the pressure drop between the low-pressure turbine interstage support guide vanes 41 and the working blades 42) by the formula get.

[0062] It should be noted that the above formulas are all dimensionless calculations.

[0063] Exemplary:

[0064] When the expansion angle of the turbine meridian inner end wall 2 is 4° and the expansion angle of the turbine meridian outer end wall 1 is 13°; the turbine load factor is 1.33, the flow coefficient is 0.56, the reaction force is 0.42, the expansion ratio is 1.85, and the turbine efficiency is 0.885;

[0065] The blade profile characteristics of the interstage support guide vane 41 are as follows: Figure 3 As shown, the leading edge diameter of the thin blade ranges from 5.0 mm to 13.0 mm, and the trailing edge diameter ranges from 1.0 mm to 2.0 mm; Figure 4 As shown, the leading edge diameter of the thick blade is 5.0 mm to 13.0 mm, and the trailing edge diameter is 1.0 mm to 2.0 mm;

[0066] like Figure 5 As shown, the blade profile characteristics of the working blade 42 are: the leading edge diameter is 2.0 mm to 3.5 mm, the trailing edge diameter is 0.8 mm to 1.2 mm, the tip-root blade area ratio is 0.49; the lift coefficient of the working blade 42 is 0.75 to 0.90;

[0067] The three-dimensional simulation results of the performance parameters of the preset state points of the low-pressure turbine mechanism are shown in the following table:

[0068] parameter Maximum climb Maximum cruise Normal takeoff Maximum takeoff Efficiency increase +1.14％ +1.14％ +0.68％ +0.80％ Power increase +1.59％ +0.46％ +1.16％ +1.24％

[0069] It can be seen from the table that the performance parameters all have margins, so the blade profiles of the interstage support guide vanes 41 and the working blades 42 mentioned above are the final blade profiles.

[0070] It should be noted that as long as the efficiency increase and the power increase are greater than 0, it means there is a margin, and if they are less than or equal to 0, it means there is no margin.

[0071] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installation", "connection", and "connection" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection, an electrical connection, or mutual communication; it can be a direct connection, or an indirect connection through an intermediate medium, it can be the internal connection of multiple components or the interaction relationship of multiple components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0072] Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent substitutions for some of the technical features therein; and these modifications or substitutions do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A three-rotor mechanism with front interstage supports, characterized in that: The invention comprises a turbine meridian outer end wall, a turbine meridian inner end wall, a high-pressure turbine assembly, a low-pressure turbine assembly and a power turbine assembly; the turbine meridian inner end wall is arranged inside the turbine meridian outer end wall to form an annular channel; from the inlet end to the outlet end of the annular channel, the high-pressure turbine assembly, the low-pressure turbine assembly and the power turbine assembly are arranged in sequence in the annular channel; The low-pressure turbine assembly includes an inter-stage support guide vane and a working blade. The inter-stage support guide vane is arranged close to the high-pressure turbine assembly, and the working blade is arranged close to the power turbine assembly.

2. The three-rotor mechanism with front interstage supports according to claim 1, characterized in that: The working blades are non-cooling shrouded blades.

3. A three-rotor mechanism with front interstage supports according to any one of claims 1-2, characterized in that: An accessory passage is provided on the interstage support guide vane.

4. A three-rotor engine with front interstage support, characterized in that: It comprises a three-rotor mechanism as claimed in any one of claims 1 to 3.

5. A method for optimizing turbine blade profile of a three-rotor mechanism with pre-positioned interstage supports, characterized in that: The steps include: Determine the structural features of the turbine meridian outer end wall and the turbine meridian inner end wall of the three-rotor mechanism as described in claims 1-3 according to the aerodynamic design constraints, and design the blade profiles of the interstage support guide vanes and the working blades according to the aerodynamic design constraints; According to the structural features, the blade profiles of the interstage support guide vanes and the working blades, the matching relationship between the high-pressure turbine assembly and the low-pressure turbine assembly, and the cooling seal of the three-rotor mechanism, a three-dimensional simulation joint calculation is performed on the three-rotor mechanism, and the performance parameters of the preset state points of the three-rotor mechanism are obtained; If the performance parameters have a margin, the blade profiles of the interstage support guide vanes and the working blades are obtained; If the performance parameters do not have a margin, the blade profiles of the interstage support guide vanes and the working blades are redesigned according to the aerodynamic design constraints of the three-rotor mechanism to carry out the three-dimensional simulation calculation until the performance parameters obtained have a margin.

6. The turbine blade profile optimization method of a three-rotor mechanism with pre-positioned interstage supports according to claim 5, characterized in that: The structural features are the expansion angle of the turbine meridian inner end wall and the expansion angle of the turbine meridian outer end wall.

7. The turbine blade profile optimization method of a three-rotor mechanism with pre-positioned interstage supports according to claim 6, characterized in that: The aerodynamic design constraints include a low-pressure turbine load coefficient, a low-pressure turbine flow coefficient, a low-pressure turbine reaction force, a low-pressure turbine expansion ratio and a low-pressure turbine efficiency.

8. The turbine blade profile optimization method of a three-rotor mechanism with pre-positioned interstage supports according to claim 7, characterized in that: The low-pressure turbine load factor is determined according to the low-pressure turbine rim power, the low-pressure turbine mid-diameter tangential speed, the inlet tangential speed of the low-pressure turbine working blades and the outlet tangential speed of the low-pressure turbine working blades.

9. A turbine blade profile optimization method for a three-rotor mechanism with pre-positioned interstage supports according to any one of claims 7-8, characterized in that: The low-pressure turbine flow coefficient is determined according to the low-pressure turbine outlet axial velocity and the low-pressure turbine mid-diameter tangential velocity.

10. A turbine blade profile optimization method for a three-rotor mechanism with pre-positioned interstage supports according to any one of claims 7-8, characterized in that: The low-pressure turbine reaction force is determined according to the low-pressure turbine mid-diameter tangential velocity, the inlet tangential velocity of the low-pressure turbine working blades, and the outlet tangential velocity of the low-pressure turbine working blades.

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