A turboprop engine power turbine blade and design method thereof

Abstract

The present application provides a turboprop engine power turbine blade and a design method thereof, the method comprising the following steps: Step 1: According to the design requirements of the power turbine at the design state point, the initial angle of attack is selected as zero, and a preliminary turbine blade shape design is carried out; Step 2: According to the preliminary turbine blade shape design results, a calculation model is built, and multi-state calculation analysis is carried out to obtain the radial distribution of the blade outlet airflow angle at the typical state point; Step 3: According to the radial distribution of the blade outlet airflow angle at the typical state point, the inlet angle of attack range is optimized, and the blade geometric modeling parameters are adjusted to carry out blade blade type optimization design; Step 4: The turbine characteristic calculation based on the optimized blade is carried out to obtain the assessment parameters at different state points; Step 5: The assessment parameters are judged, and the blade modeling parameters and the turbine characteristics are iteratively adjusted according to the judgment results until the blade meets all preset conditions. The present application can improve the turbine efficiency at non-design points and reduce the fuel consumption rate of the whole machine within the full envelope range.

Description

Technical Field

[0001] The present application relates to the field of turboprop engines, and in particular, to a turboprop engine power turbine blade and a design method thereof. Background Art

[0002] Turboprop engines have the advantages of good economy, large take-off thrust, simple maintenance, strong environmental adaptability, etc., and are widely used in the field of small and medium-sized regional passenger aircraft and general aircraft. During the operation of a turboprop engine, the propeller is usually driven by a power turbine. Due to the different loads of the aircraft during flight.

[0003] Due to the limitation of the linear speed of the propeller blade tip, the power turbine of a turboprop engine usually has multiple different operating speeds. For example, in the ground takeoff state, the engine requires higher power and the power turbine usually has a higher operating speed; in the air cruising state, if the rotor speed is still maintained at a high level, the propeller blade tip will be supersonic, resulting in reduced propeller efficiency. Therefore, in the cruising state, a lower operating speed is usually used based on the overall performance of the propulsion system. Multiple different operating speeds pose a great challenge to the design of the power turbine. Traditional turboshaft engine power turbines usually adopt a constant speed design. The power turbine speed remains constant in takeoff, cruising, and continuous states. Under different working conditions, the load and angle of attack of the turbine blades change little. Therefore, when designing turbine blades, it is usually only necessary to consider the turbine angle of attack at the design speed, and complete the turbine blade design based on aerodynamic parameters such as blade load level and speed triangle at the design point. For the power turbine of a turboprop engine, the design of the blades needs to consider the balance between multiple different speeds. Because the efficiency of the power turbine directly affects the fuel consumption rate of the entire engine, if the performance of non-design state points is not carefully considered during the blade design process, it will inevitably lead to a decrease in turbine efficiency and an increase in the overall cycle fuel consumption rate, affecting the economic efficiency of use throughout the life cycle. Summary of the invention

[0004] On the one hand, an embodiment of the present application provides a method for designing power turbine blades for a turboprop engine to solve the technical problem that the existing power turbine blades only consider the design requirements of a certain state point, resulting in increased angle of attack losses at non-design points, reduced turbine efficiency, and affecting the overall fuel consumption rate within the full envelope range of the engine.

[0005] The technical solutions adopted in this application are as follows:

[0006] A method for designing a turboprop engine power turbine blade comprises the following steps:

[0007] Step 1: According to the design requirements of the power turbine at the design state point, the initial angle of attack is selected as zero degrees, and preliminary turbine blade shape design is carried out;

[0008] Step 2: Build a calculation model based on the preliminary turbine blade shape design results, carry out multi-state calculation analysis, and obtain the radial distribution of the blade outlet airflow angle at typical state points, wherein the typical state points include a high-altitude cruising state point and a take-off state point;

[0009] Step 3: According to the radial distribution of the airflow angle at the blade outlet at the typical state point, the inlet angle of attack range is optimized, the blade geometric modeling parameters are adjusted, and the blade profile optimization design is carried out;

[0010] Step 4: Carry out turbine characteristic calculation based on optimized blades to obtain assessment parameters at different state points, wherein the assessment parameters include power turbine efficiency, power, and physical flow at different state points;

[0011] Step 5: Evaluate the assessment parameters. If all of them meet the design requirements, calculate the strength, vibration and life of the rotor. If the strength, vibration and life calculation results meet the requirements, the turbine blade design is completed. If any of the calculation results of the power turbine efficiency, power, physical flow, strength, vibration and life do not meet the requirements, return to step 3, readjust the blade shaping parameters, perform blade profile design, and recalculate and analyze the turbine characteristics until the optimized blade meets all preset conditions.

[0012] Furthermore, in step 1, the design state point is selected as the high-altitude cruising state point, and the blade attack angle i is determined by the following formula:

[0013] i=β k -β;

[0014] Where β is the blade inlet airflow angle; β k The blade inlet configuration angle.

[0015] Furthermore, in step 1, when carrying out the preliminary turbine blade shape design, the inlet blade angle is determined according to the selected initial angle of attack, and then by optimizing the geometric parameters including the outlet blade angle, the leading and trailing edge radius, the consistency, the chord length, and the trailing edge blockage, a two-dimensional basic blade profile at different radial heights is constructed, and each basic blade profile is superimposed radially according to a certain stacking rule to form a three-dimensional preliminary turbine blade shape.

[0016] Furthermore, in step 3, the blade geometric parameters are adjusted by increasing the blade chord length, blade height, inlet and outlet wedge angles or adjusting the blade inlet structural angle β k way to proceed.

[0017] Furthermore, the blade inlet structural angle β at the design state point k Satisfy: β k =β des +i des , the i desis the attack angle at the design state point, which is between 3° and 5°, β des is the blade inlet airflow angle at the design state point.

[0018] On the other hand, the present application also provides a turboprop engine power turbine blade, characterized in that the turboprop engine power turbine blade includes a blade body, a blade crown, an upper edge plate, a lower edge plate and a tenon, the blade body is formed by stacking a number of two-dimensional basic blade profiles radially according to a certain rule; the three-dimensional entity of the turboprop engine power turbine blade is formed by stacking N two-dimensional basic blade profiles along the radial direction according to a certain rule, wherein N is a natural number, N≥3; the two-dimensional basic blade profile includes a leading edge, a trailing edge, a blade basin profile line and a blade back profile line, and is designed by selecting different inlet and outlet structural angles, leading edge wedge angles, trailing edge wedge angles, trailing edge bending angles, leading edge radius, trailing edge radius, consistency, chord length, and trailing edge blockage degree.

[0019] Each two-dimensional basic blade profile includes a blade root section basic blade profile, a blade tip section basic blade profile and a plurality of blade mid-section basic blade profiles, wherein the blade root section basic blade profile is defined as a two-dimensional basic blade profile at 5% of the blade height from the upper surface of the lower edge plate; the blade tip section basic blade profile is defined as a two-dimensional basic blade profile at 95% of the blade height from the upper surface of the lower edge plate; the blade mid-section basic blade profile is defined as a two-dimensional basic blade profile with a blade profile height between the blade root and blade tip sections;

[0020] The radius r1 of the leading edge is between 0.55 and 0.9 mm, and the radius r t , between 0.25 and 0.35 mm.

[0021] Furthermore, the leading edge wedge angle of the two-dimensional basic blade profile is between 32° and 48°.

[0022] Furthermore, the ratio of the basic blade profile surface area of ​​the blade root cross section to the basic blade profile surface area of ​​the blade tip cross section of the two-dimensional basic blade profile is between 1.8 and 2.5.

[0023] Furthermore, the ratio of the basic blade profile chord length of the blade tip section to the basic blade profile chord length of the blade root section of the two-dimensional basic blade profile is between 0.8 and 1.0.

[0024] Compared with the prior art, this application has the following beneficial effects:

[0025] The present application provides a turboprop engine power turbine blade and a design method thereof. The design method reduces the angle of attack loss caused by a wide range of turbine speed changes at non-design points by using a design concept different from that of conventional constant-speed power turbine blades. While ensuring the design point performance and the blade structural strength, the non-design point performance of the turbine is taken into account, and the problem of excessive reduction in non-design point efficiency and increased overall fuel consumption rate due to speed changes in different states of the turboprop engine power turbine is solved. The non-design point turbine efficiency can be improved, and the overall fuel consumption rate within the full flight envelope can be reduced; the maximum equivalent stress at the blade root can be reduced, and the strength and life of the blade can be increased; the method can adapt to multiple different physical speeds, and the angle of attack adaptability is better.

[0026] In addition to the above-described purposes, features and advantages, the present application also has other purposes, features and advantages. The present application will be further described in detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments and descriptions of the present application are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0028] Figure 1 It is a flow chart of a turboprop engine power turbine blade design method according to a preferred embodiment of the present application.

[0029] Figure 2 It is a schematic diagram of the radial distribution of the airflow angle of the blades of the preferred embodiment of the present application.

[0030] Figure 3 It is a three-dimensional solid schematic diagram of the turbine blade of the preferred embodiment of the present application.

[0031] Figure 4 It is a schematic diagram of the two-dimensional basic blade profile of the preferred embodiment of the present application.

[0032] In the figure: 1. Two-dimensional basic blade shape; 2. Blade body; 3. Blade crown; 4. Upper edge plate; 5. Lower edge plate; 6. Tenon; 7. Basic blade shape of blade tip section; 8. Basic blade shape of blade root section; 9. Blade basin line; 10. Blade back line; 11. Trailing edge; 12. Leading edge. DETAILED DESCRIPTION

[0033] It should be noted that, in the absence of conflict, the embodiments and features in the embodiments of the present application can be combined with each other. The present application will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0034] Compared with gas turbines, power turbines have lower tip leakage losses due to the crowned blades. At the same time, due to the large aspect ratio, the end zone secondary flow influence area such as the channel vortex is also smaller than that of gas turbines. On the other hand, due to the low inlet temperature of the power turbine, the power turbine rotor usually adopts a non-cooling design, so the cold air mixing loss is also small. Therefore, compared with gas turbines, the key factors affecting the flow loss of power turbines are blade loss and angle of attack loss. Unreasonable blade load distribution and large angle of attack loss are important factors that lead to reduced aerodynamic performance of power turbines. According to this idea, the main purpose of this application is to propose a blade design method that can effectively balance the angle of attack conditions in different working conditions, so as to effectively reduce the angle of attack loss of non-design points while ensuring the performance of the design point.

[0035] like Figure 1 As shown, a preferred embodiment of the present application provides a method for designing a turboprop engine power turbine blade, comprising the following steps:

[0036] S1: According to the design requirements of the power turbine at the design state point, the initial angle of attack is selected as zero degrees, and preliminary turbine blade shape design is carried out;

[0037] S2: Build a computational model based on the preliminary turbine blade design results, conduct multi-state computational analysis, and obtain the radial distribution of the blade outlet airflow angle at typical state points (see Figure 2 ), the typical state points include a high altitude cruising state point and a take-off state point;

[0038] S3: According to the radial distribution of the airflow angle of the blade outlet at the typical state point, the inlet angle of attack range is optimized, the blade geometric modeling parameters are adjusted, and the blade profile optimization design is carried out;

[0039] S4: Carry out turbine characteristic calculation based on optimized blades to obtain assessment parameters at different state points, wherein the assessment parameters include power turbine efficiency, power, and physical flow at different state points;

[0040] S5: Evaluate the assessment parameters. If all of them meet the design requirements, calculate the strength, vibration and life of the rotor. If the strength, vibration and life calculation results meet the requirements, the turbine blade design is completed. If any of the calculation results of the power turbine efficiency, power, physical flow, strength, vibration and life do not meet the requirements, return to step S3, readjust the blade shaping parameters, perform blade profile design, and recalculate and analyze the turbine characteristics until the optimized blade meets all preset conditions.

[0041] The present embodiment provides a method for designing power turbine blades for a turboprop engine. The method is different from the conventional constant speed power turbine blade design concept and can reduce the angle of attack loss at non-design points due to a wide range of turbine speed changes. While ensuring the design point performance and the blade structural strength, the non-design point performance of the turbine is taken into account, and the problem of excessive reduction in non-design point efficiency and increased overall fuel consumption rate due to speed changes in different states of the turboprop engine power turbine is solved. The non-design point turbine efficiency can be improved and the overall fuel consumption rate within the full flight envelope can be reduced; the maximum equivalent stress at the blade root can be reduced and the blade strength and life can be increased; it can adapt to multiple different physical speeds and have better angle of attack adaptability.

[0042] Preferably, for a turboprop engine, the maximum cruise point power turbine speed is usually lower than the ground takeoff state speed. Therefore, in step S1, the design state point is selected as the high-altitude cruise state point, and the blade attack angle i is determined by the following formula:

[0043] i=β k -β;

[0044] Where β is the blade inlet airflow angle; β k The blade inlet configuration angle.

[0045] Preferably, in step S1, when carrying out preliminary turbine blade shape design, the inlet blade angle is determined according to the selected initial angle of attack, and then two-dimensional basic blade profiles at different radial heights are constructed by optimizing geometric parameters including the outlet blade angle, front and trailing edge radius, consistency, chord length, and trailing edge blockage. The basic blade profiles are superimposed radially according to a certain stacking rule to form a three-dimensional preliminary turbine blade shape.

[0046] Preferably, in step S3, the blade geometric parameters are adjusted by increasing the blade chord length, blade height, inlet and outlet wedge angles or adjusting the blade inlet structural angle β k way to proceed.

[0047] Preferably, the blade inlet configuration angle β at the design state point k Satisfy: β k =β des +i des , the i des is the attack angle at the design state point, which is between 3° and 5°, β desis the blade inlet airflow angle at the design state point. In this embodiment, when shaping the blade geometric profile, the given blade inlet structural angle is greater than the blade inlet airflow angle, so that when it is in the high-altitude cruising state, the blade angle of attack i is at a positive value, and as the turbine operating speed increases, the blade geometric structural angle remains unchanged, and the blade inlet airflow angle gradually increases, so the angle of attack i gradually changes from a positive value to a negative value. Based on this design, the blade provided in this embodiment takes into account both the take-off and cruising design states, so that the turbine angle of attack in both design states is not too large to cause a significant increase in the angle of attack loss, thereby ensuring the turbine aerodynamic efficiency within the full flight envelope.

[0048] like Figure 3 and Figure 4 As shown, another preferred embodiment of the present application further provides a turboprop engine power turbine blade, the turboprop engine power turbine blade (working blade) includes a blade body 2, a blade crown 3, an upper edge plate 4, a lower edge plate 5 and a tenon 6, the blade body 2 is formed by stacking a number of two-dimensional basic blade profiles 1 along the radial direction according to a certain rule; the three-dimensional entity of the turboprop engine power turbine blade is formed by stacking N two-dimensional basic blade profiles 1 along the radial direction according to a certain rule, wherein N is a natural number, N≥3; the two-dimensional basic blade profile 1 is composed of a leading edge 12, a trailing edge 11, a blade basin profile 9 and a blade back profile 10, and is designed by selecting different inlet and outlet structural angles, leading edge wedge angles, trailing edge wedge angles, trailing edge bending angles, leading edge radius, trailing edge radius, consistency, chord length, and trailing edge blockage degree.

[0049] Each two-dimensional basic blade profile 1 includes a blade root section basic blade profile 8, a blade tip section basic blade profile 7 and a plurality of blade mid-section basic blade profiles, wherein the blade root section basic blade profile 8 is defined as a two-dimensional basic blade profile at a distance of 5% of the blade height from the upper surface of the lower edge plate; the blade tip section basic blade profile 7 is defined as a two-dimensional basic blade profile at a distance of 95% of the blade height from the upper surface of the lower edge plate; the blade mid-section basic blade profile is defined as a two-dimensional basic blade profile with a blade profile height between the blade root and blade tip sections;

[0050] The radius r1 of the leading edge is between 0.55 and 0.9 mm, and the radius r t , between 0.25 and 0.35 mm, in this embodiment, the two-dimensional basic blade section r t The values ​​are all 0.3, and the r1 values ​​are all 0.7.

[0051] Preferably, the two-dimensional basic blade profile is designed with a larger leading edge wedge angle, so the leading edge wedge angle of the two-dimensional basic blade profile is between 32° and 48° to reduce the angle of attack sensitivity of the turbine blade at different speeds. In this embodiment, the leading edge wedge angles of the basic blade profile at the root and tip sections are 42° and 34° respectively.

[0052] Preferably, since the turbine blade adopts a crown design, a larger root-tip area ratio is adopted to increase the strength margin of the blade root portion. Therefore, the ratio of the basic blade profile surface area of ​​the blade root section to the basic blade profile surface area of ​​the blade tip section of the two-dimensional basic blade profile is between 1.8 and 2.5, and the value in this embodiment is 2.33.

[0053] Preferably, when designing the basic blade profile of the blade tip section, the chord length b can be reduced to reduce the weight of the blade crown, thereby reducing the maximum equivalent stress level of the blade root and the blade crown and improving the strength and life of the blade. Therefore, the ratio of the chord length of the basic blade profile of the blade tip section to the chord length of the basic blade profile of the blade root section of the two-dimensional basic blade profile is between 0.8 and 1.0, and the value in this embodiment is 0.85.

[0054] In summary, the above-mentioned embodiments provide a turboprop engine power turbine blade and a design method thereof, which reduces the angle of attack loss at non-design points due to a wide range of turbine speed changes by reasonably selecting the blade angle of attack and shape parameters at the design point. While ensuring the design point performance and the blade structural strength, the non-design point performance of the turbine is taken into account, thereby reducing the fuel consumption of the entire machine within the full flight envelope.

[0055] The above description is only a preferred embodiment of the present application and is not intended to limit the present application. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present application should be included in the protection scope of the present application.

Claims

1. A method for designing a turboprop engine power turbine blade, characterized in that: The following steps are involved: Step 1: According to the design requirements of the power turbine at the design state point, the initial angle of attack is selected as zero degrees, and preliminary turbine blade shape design is carried out; Step 2: Build a computational model based on the preliminary turbine blade shape design results, conduct multi-state computational analysis, and obtain the radial distribution of blade outlet airflow angles at typical state points, including high-altitude cruising state points and take-off state points; Step 3: According to the radial distribution of the airflow angle at the blade outlet at the typical state point, select the inlet angle of attack range, adjust the blade geometric modeling parameters, and carry out blade profile optimization design; Step 4: Carry out turbine characteristic calculation based on optimized blades to obtain assessment parameters at different state points, wherein the assessment parameters include power turbine efficiency, power, and physical flow at different state points; Step 5: Evaluate the assessment parameters. If all of them meet the design requirements, calculate the strength, vibration and life of the rotor. If the strength, vibration and life calculation results meet the requirements, the turbine blade design is completed. If any of the calculation results of the power turbine efficiency, power, physical flow, strength, vibration and life do not meet the requirements, return to step 3, readjust the blade shaping parameters, perform blade profile design, and recalculate and analyze the turbine characteristics until the optimized blade meets all preset conditions.

2. The turboprop engine power turbine blade design method according to claim 1, characterized in that: In step 1, the design state point is selected as the high altitude cruising state point, and the blade angle of attack is i Determined by the following formula: i= β k - β ； in, β is the blade inlet airflow angle; β k The blade inlet configuration angle.

3. The turboprop engine power turbine blade design method according to claim 1, characterized in that: In step 1, when carrying out the preliminary turbine blade shape design, the inlet blade angle is determined according to the selected initial angle of attack, and then by selecting geometric parameters including the outlet blade angle, front and trailing edge radius, consistency, chord length, and trailing edge blockage, a two-dimensional basic blade profile at different radial heights is constructed. Each basic blade profile is superimposed radially according to a certain stacking rule to form a three-dimensional preliminary turbine blade shape.

4. The turboprop engine power turbine blade design method according to claim 1, characterized in that: In step 3, the blade geometric parameters are adjusted by increasing the blade chord length, blade height, inlet and outlet wedge angles or adjusting the blade inlet structural angle β k way to proceed.

5. The turboprop engine power turbine blade design method according to claim 1, characterized in that: Blade inlet structural angle at design state point β k satisfy: β k = β des + i des , i des is the angle of attack at the design state point, which is between 3° and 5°. β des is the blade inlet airflow angle at the design state point.

6. A turboprop engine power turbine blade, characterized in that: The turboprop engine power turbine blade comprises a blade body, a blade crown, an upper edge plate, a lower edge plate and a tenon, wherein the blade body is formed by stacking a number of two-dimensional basic blade profiles radially according to a certain rule; the three-dimensional entity of the turboprop engine power turbine blade is formed by stacking N two-dimensional basic blade profiles along the radial direction according to a certain rule, wherein N is a natural number, N≥3; the two-dimensional basic blade profile comprises a leading edge, a trailing edge, a blade basin profile line and a blade back profile line, and is designed by selecting different inlet and outlet structural angles, leading edge wedge angles, trailing edge wedge angles, trailing edge bending angles, leading edge radius, trailing edge radius, consistency, chord length, and trailing edge blockage degree. Each two-dimensional basic blade profile includes a blade root section basic blade profile, a blade tip section basic blade profile and a plurality of blade mid-section basic blade profiles, wherein the blade root section basic blade profile is defined as a two-dimensional basic blade profile at a distance of 5% of the blade height from the upper surface of the lower edge plate; the blade tip section basic blade profile is defined as a two-dimensional basic blade profile at a distance of 95% of the blade height from the upper surface of the lower edge plate; the blade mid-section basic blade profile is defined as a two-dimensional basic blade profile with a blade profile height between the blade root and blade tip sections; The radius value of the leading edge r 1 is between 0.55 and 0.9 mm, the radius of the trailing edge r t , between 0.25~0.35mm.

7. The turboprop engine power turbine blade according to claim 6, characterized in that: The leading edge wedge angle of the two-dimensional basic blade profile is between 32° and 48°.

8. The turboprop engine power turbine blade according to claim 6, characterized in that: The ratio of the basic blade profile surface area of ​​the blade root section to the basic blade profile surface area of ​​the blade tip section of the two-dimensional basic blade profile is between 1.8 and 2.

5.

9. The turboprop engine power turbine blade according to claim 6, characterized in that: The ratio of the basic blade profile chord length of the blade tip section to the basic blade profile chord length of the blade root section of the two-dimensional basic blade profile is between 0.8 and 1.0.

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