CN115098966A - Power turbine blade of turboprop engine and design method thereof - Google Patents

Power turbine blade of turboprop engine and design method thereof Download PDF

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Publication number
CN115098966A
CN115098966A CN202210758915.3A CN202210758915A CN115098966A CN 115098966 A CN115098966 A CN 115098966A CN 202210758915 A CN202210758915 A CN 202210758915A CN 115098966 A CN115098966 A CN 115098966A
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blade
design
angle
profile
turbine
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欧阳玉清
张绍文
房兴龙
刘冬华
单熠君
谭锋
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Hunan Aviation Powerplant Research Institute AECC
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Hunan Aviation Powerplant Research Institute AECC
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    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06T—IMAGE DATA PROCESSING OR GENERATION, IN GENERAL
    • G06T17/00—Three dimensional [3D] modelling, e.g. data description of 3D objects
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/04—Ageing analysis or optimisation against ageing

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Abstract

The application provides a power turbine blade of a turboprop engine and a design method thereof, wherein the method comprises the following steps: step 1, selecting an initial attack angle as zero degree according to the design requirement of a power turbine at a design state point, and carrying out primary turbine blade modeling design; step 2, building a calculation model according to the preliminary turbine blade modeling design result, carrying out multi-state calculation analysis, and obtaining radial distribution of blade outlet airflow angles at typical state points; step 3, adjusting geometric modeling parameters of the blades to carry out blade profile optimization design according to radial distribution of blade outlet airflow angles at typical state points and optimizing the range of inlet attack angles; step 4, carrying out turbine characteristic calculation based on the optimized blades and obtaining assessment parameters of different state points; and 5, evaluating the evaluation parameters, iteratively adjusting the blade modeling parameters and calculating the turbine characteristics according to the evaluation result until the blades meet all preset conditions. The method and the device can improve the efficiency of the turbine at the non-design point and reduce the oil consumption rate of the whole turbine in the full-envelope range.

Description

Turboprop engine power turbine blade and design method thereof
Technical Field
The application relates to the field of turboprop engines, in particular to a power turbine blade of a turboprop engine and a design method thereof.
Background
The turboprop engine has the advantages of good economy, large take-off tension, simple maintenance, strong environmental applicability and the like, and is widely applied to the fields of small and medium-sized branch passenger planes and general airplanes. The propeller is usually driven by a power turbine during operation of the turboprop. Due to differences in aircraft loads during flight.
Because of the limitation of the linear speed of the blade tip of the propeller, the power turbine of the turboprop engine usually has a plurality of different working rotating speeds, for example, in a ground takeoff state, the required power of the engine is higher, and the working rotating speed of the power turbine is usually higher; in the cruise state in the air, if the rotor speed is kept high, the tip of the propeller blade is made to be ultrasonic, so that the working efficiency of the propeller is reduced. The plurality of different operating speeds presents significant challenges to the design of the power turbine. The traditional power turbine of the turboshaft engine is usually designed with a constant rotating speed, the rotating speed of the power turbine is kept constant in take-off, cruise, continuous and other states, and the load and attack angle of the turbine blade are changed slightly in different working states. Therefore, in the design of the turbine blade, the turbine blade is usually designed according to aerodynamic parameters such as blade load level and speed triangle at the design point by only considering the attack angle of the turbine at the design rotation speed. For a turboprop engine power turbine, the design of the blades needs to consider the balance among a plurality of different rotating speeds. Because the efficiency of the power turbine directly affects the oil consumption rate of the whole engine, if the performance of a non-design state point is not fully considered in the blade profile design process, the efficiency of the turbine is reduced, the circulating oil consumption rate of the whole engine is increased, and the use economy in the whole life cycle is affected.
Disclosure of Invention
The embodiment of the application provides a design method for a power turbine blade of a turboprop engine on the one hand, and aims to solve the technical problems that the loss of an attack angle at a non-design point is increased, the efficiency of a turbine is reduced, and the overall oil consumption rate of the engine in a full-envelope range is influenced because the design requirement of a certain state point is only considered in the conventional power turbine blade.
The technical scheme adopted by the application is as follows:
a design method for a power turbine blade of a turboprop engine comprises the following steps:
step 1, selecting an initial attack angle as zero degree according to the design requirement of a power turbine at a design state point, and carrying out primary turbine blade modeling design;
step 2, building a calculation model according to a preliminary turbine blade modeling design result, carrying out multi-state calculation analysis, and obtaining radial distribution of blade outlet airflow angles at typical state points, wherein the typical state points comprise high-altitude cruise state points and takeoff state points;
step 3, according to the radial distribution of the blade outlet airflow angle of the typical state point and the optimized inlet attack angle range, adjusting the geometric modeling parameters of the blade and developing the blade profile optimization design of the blade;
step 4, carrying out turbine characteristic calculation based on the optimized blades to obtain assessment parameters of different state points, wherein the assessment parameters comprise power turbine efficiency, power and physical flow of the different state points;
step 5, judging the assessment parameters, if the assessment parameters meet the design requirements, calculating the strength, vibration and service life of the rotor, and if the calculation results of the strength, vibration and service life meet the requirements, finishing the design of the turbine blade; and if any one of the calculation results of the efficiency, the power, the physical flow, the strength, the vibration and the service life of the power turbine does not meet the requirements, returning to the step 3, readjusting the modeling parameters of the blades, designing the blade profile of the blades, and performing calculation and analysis on the turbine characteristics again until the optimized blades meet all preset conditions.
Further, in step 1, the design state point is selected as a high altitude cruise state point, and the blade attack angle i is determined by the following formula:
i＝β k -β；
wherein β is the blade inlet flow angle; beta is a beta k An angle is constructed for the blade inlet.
Further, in step 1, when the preliminary turbine blade modeling design is developed, an inlet blade angle is determined according to the selected initial attack angle condition, two-dimensional basic blade profiles with different radial heights are constructed through the preferred geometric parameters including an outlet blade angle, a leading edge radius, a consistency, a chord length and a trailing edge blockage degree, and the basic blade profiles are overlapped along the radial direction according to a certain stacking rule to form a three-dimensional preliminary turbine blade modeling.
Further, in step 3, the geometric modeling parameters of the blades are adjusted by increasing the chord length of the blades, the height of the blades, the wedge angle of the inlet and the outlet or adjusting the inlet structure angle beta of the blades k In the manner of (a).
Further, the design state point vane inlet configuration angle β k Satisfies the following conditions: beta is a k ＝β des +i des The said i des The angle of attack is between 3 DEG and 5 DEG, beta des The blade inlet flow angle is the design condition point.
The turbine blade is characterized by comprising 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 plurality of two-dimensional basic blade profiles along the radial direction according to a certain rule; the three-dimensional entity of the power turbine blade of the turboprop engine is formed by stacking N two-dimensional basic blade profiles along the radius direction by a certain rule, wherein N is a natural number and is not less than 3; the two-dimensional basic blade profile comprises a front edge, a tail edge, a blade basin profile line and a blade back profile line, and is designed by selecting different inlet and outlet construction angles, a front edge wedge angle, a tail edge bending angle, a front edge radius, a tail edge radius, consistency, chord length and tail edge blockage degree;
each two-dimensional foundation blade profile comprises a blade root section foundation blade profile, a blade tip section foundation blade profile and a plurality of blade leaf section foundation blade profiles, wherein the blade root section foundation blade profile is defined as the two-dimensional foundation blade profile at a blade height which is 5% of the distance 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 position which is 95% of the blade height away from the upper surface of the lower edge plate; the blade-in-section basic blade profile is defined as a two-dimensional basic blade profile with the blade profile height between the blade root and the blade tip section;
radius value r of the leading edge 1 Between 0.55 and 0.9mm, the radius r of the trailing edge t And is between 0.25 and 0.35 mm.
Further, the value of the wedge angle of the leading edge of the two-dimensional foundation blade profile is between 32 and 48 degrees.
Furthermore, the ratio of the area of the blade root section basic blade profile surface to the area of the blade tip section basic blade profile surface of the two-dimensional basic blade profile is between 1.8 and 2.5.
Further, the ratio of the chord length of the blade tip section basic blade profile of the two-dimensional basic blade profile to the chord length of the blade root section basic blade profile is between 0.8 and 1.0.
Compared with the prior art, the method has the following beneficial effects:
the design method reduces the attack angle loss of a non-design point caused by the large-range change of the rotating speed of a turbine through a design idea different from a conventional constant-rotating-speed power turbine blade, considers the performance of the non-design point of the turbine on the premise of ensuring the performance of the design point and the structural strength of the blade, solves the problems of excessive reduction of the efficiency of the non-design point and increase of the oil consumption rate of the whole machine caused by the rotating speed change of different states of the power turbine of the turboprop engine, can improve the efficiency of the turbine of the non-design point, and reduces the oil consumption rate of the whole machine in the range of a full flight envelope; the maximum equivalent stress of the blade root is reduced, and the strength life of the blade is prolonged; the method is suitable for a plurality of different physical rotating speeds, and the attack angle adaptability is better.
In addition to the objects, features and advantages described above, other objects, features and advantages will be apparent from the present application. The present application will now be described in further detail with reference to the accompanying drawings.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the application and, together with the description, serve to explain the application and are not intended to limit the application. In the drawings:
FIG. 1 is a schematic flow chart of a method for designing a power turbine blade of a turboprop engine according to a preferred embodiment of the present application.
FIG. 2 is a schematic view of the vane airflow angle radial distribution of the preferred embodiment of the present application.
FIG. 3 is a three-dimensional solid representation of a turbine blade in accordance with a preferred embodiment of the present application.
FIG. 4 is a schematic view of a two-dimensional base profile of a preferred embodiment of the present application.
In the figure: 1. a two-dimensional basal leaf profile; 2. a leaf body; 3. a leaf shroud; 4. an upper edge plate; 5. a lower flange plate; 6. a tenon; 7. a blade tip section basic blade profile; 8. a blade root section basic blade profile; 9. a leaf basin molded line; 10. a blade back profile; 11. a trailing edge; 12. a leading edge.
Detailed Description
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present application will be described in detail below with reference to the embodiments with reference to the attached drawings.
Compared with a gas turbine, the power turbine has the advantages that the tip leakage loss of the power turbine is low due to the fact that the blade is provided with the shroud, meanwhile, due to the fact that the aspect ratio is large, the secondary flow influence area of the channel vortex and other end areas is smaller than that of the gas turbine, on the other hand, due to the fact that the inlet temperature of the power turbine is low, the power turbine rotor is usually designed without cooling, cold air mixing loss is also small, and therefore compared with the gas turbine, the key of influencing the flow loss of the power turbine is blade profile loss and attack angle loss. Unreasonable blade load distribution and large attack angle loss are important factors causing the aerodynamic performance of the power turbine to be reduced. According to the thought, the main point of the method is to provide the blade type design method capable of effectively balancing the attack angle conditions of different working states, so that the attack angle loss of non-design points is effectively reduced on the premise of ensuring the performance of the design points.
As shown in FIG. 1, the preferred embodiment of the present application provides a method for designing a power turbine blade of a turboprop engine, comprising the steps of:
s1, selecting an initial attack angle as zero degree according to the design requirement of the power turbine at the design state point, and carrying out the primary modeling design of the turbine blade;
s2, building a calculation model according to the preliminary turbine blade modeling design result, carrying out multi-state calculation analysis, and obtaining blade outlet airflow angle radial distribution (shown in figure 2) of typical state points, wherein the typical state points comprise a high-altitude cruise state point and a takeoff state point;
s3, adjusting geometric modeling parameters of the blades according to radial distribution of the outlet airflow angles of the blades at the typical state points and optimizing the range of the inlet attack angle, and developing blade profile optimization design;
s4, carrying out turbine characteristic calculation based on the optimized blade, and obtaining assessment parameters of different state points, wherein the assessment parameters comprise power turbine efficiency, power and physical flow of the different state points;
s5, judging the assessment parameters, if the assessment parameters meet the design requirements, calculating the strength, the vibration and the service life of the rotor, and if the calculation results of the strength, the vibration and the service life meet the requirements, finishing the design of the turbine blade; and if any one of the calculation results of the efficiency, the power, the physical flow, the strength, the vibration and the service life of the power turbine does not meet the requirements, returning to the step S3, readjusting the modeling parameters of the blades, designing the blade profile of the blades, and performing calculation and analysis on the turbine characteristics again until the optimized blades meet all preset conditions.
The design method is different from the design idea of the conventional constant-rotating-speed power turbine blade, the attack angle loss of a non-design point caused by the large-range change of the rotating speed of the turbine can be reduced, the performance of the non-design point of the turbine is considered on the premise of ensuring the performance of the design point and the structural strength of the blade, the problems of excessive reduction of the efficiency of the non-design point and increase of the oil consumption rate of the whole machine caused by the change of the rotating speeds in different states of the power turbine of the turboprop engine are solved, the efficiency of the turbine of the non-design point can be improved, and the oil consumption rate of the whole machine in the full flight envelope range is reduced; the maximum equivalent stress of the blade root is reduced, and the strength life of the blade is prolonged; the method is suitable for a plurality of different physical rotating speeds, and the adaptability of the attack angle is better.
Preferably, for a turboprop engine, the maximum cruise power turbine speed is generally lower than the ground takeoff speed, and therefore, in step S1, the design state point is selected as the high altitude cruise state point, and the blade angle of attack i is determined by the following equation:
i＝β k -β；
wherein β is the blade inlet flow angle; beta is a k An angle is constructed for the blade inlet.
Preferably, in step S1, when the preliminary turbine blade modeling design is developed, the inlet blade angle is determined according to the selected initial attack angle condition, two-dimensional basic blade profiles with different radial heights are constructed by using the geometric parameters preferably including the outlet blade angle, the leading and trailing edge radius, the consistency, the chord length, and the trailing edge blockage degree, and the basic blade profiles are radially stacked according to a certain stacking rule to form a three-dimensional preliminary turbine blade modeling.
Preferably, in step S3, the blade geometry parameters are adjusted by increasing the chord length, the height, the inlet-outlet wedge angle or adjusting the blade inlet configuration angle β of the blade k In the manner of (a).
Preferably, the design state point vane inlet configuration angle β k Satisfies the following conditions: beta is a k ＝β des +i des The said i des The angle of attack is between 3 DEG and 5 DEG, beta des The blade inlet flow angle is the design condition point. In the embodiment, when the geometric profile of the blade is modeled, the given blade inlet structure angle is larger than the blade inlet airflow angle, so that the attack angle i of the blade is in a positive value in a high-altitude cruising state, and the blade inlet airflow angle is gradually increased along with the increase of the working rotating speed of the turbine and the unchanged geometric structure angle of the bladeTherefore, the angle of attack i gradually changes from a positive value to a negative value. Based on the design, the blade provided by the embodiment gives consideration to two design states of takeoff and cruise, so that the attack angle of the turbine in the two design states is not too large, the loss of the attack angle is not obviously increased, and the aerodynamic efficiency of the turbine in the full flight envelope range is ensured.
As shown in fig. 3 and 4, another preferred embodiment of the present application further provides a power turbine blade of a turboprop engine, wherein the power turbine blade (working blade) of the turboprop engine comprises a blade body 2, a blade shroud 3, an upper edge plate 4, a lower edge plate 5 and a tenon 6, and the blade body 2 is formed by stacking a plurality of two-dimensional basic blade profiles 1 according to a certain rule along a radial direction; the three-dimensional entity of the power turbine blade of the turboprop engine is formed by stacking N two-dimensional basic blade profiles 1 along the radius direction by a certain rule, wherein N is a natural number and is not less than 3; the two-dimensional basic blade profile 1 consists of a front edge 12, a tail edge 11, a blade basin profile 9 and a blade back profile 10, and is designed by selecting different inlet and outlet construction angles, a front edge wedge angle, a tail edge bending angle, a front edge radius, a tail edge radius, consistency, chord length and tail edge blockage degree;
each two-dimensional basic blade profile 1 comprises a blade root section basic blade profile 8, a blade tip section basic blade profile 7 and a plurality of blade leaf section basic blade profiles, wherein the blade root section basic blade profile 8 is defined as the two-dimensional basic blade profile at a blade height of 5% away from the upper surface of the lower edge plate; the tip section basic blade profile 7 is defined as a two-dimensional basic blade profile at a blade height of 95% from the upper surface of the lower edge plate; the blade-in-section basic blade profile is defined as a two-dimensional basic blade profile with the blade profile height between the blade root and the blade tip section;
radius value r of the leading edge 1 Between 0.55 and 0.9mm, the radius r of the trailing edge t Between 0.25 mm and 0.35mm, in this embodiment, each two-dimensional basic blade profile section r is taken t All values are 0.3, r 1 The values were all 0.7.
Preferably, the two-dimensional basic blade profile adopts a design with a larger leading edge wedge angle, so that the value of the leading edge wedge angle of the two-dimensional basic blade profile is between 32 and 48 degrees, and the attack angle sensitivity of the turbine blade at different rotating speeds is reduced. In the present embodiment, the root and tip section base blade profile leading edge wedge angles are 42 ° and 34 °, respectively.
Preferably, the turbine blade is designed to have a shroud, and a large root tip area ratio is used to increase the strength margin of the blade root, so that the ratio of the profile area of the blade root section basic blade profile to the profile area of the blade tip section basic blade profile of the two-dimensional basic blade profile is 1.8 to 2.5, which is 2.33 in this embodiment.
Preferably, when the blade tip section basic blade profile is designed, the chord length b can be reduced to reduce the weight of the blade shroud, further reduce the maximum equivalent stress level of the blade root and the blade shroud, and improve the strength life of the blade, so that the ratio of the chord length of the blade tip section basic blade profile to the chord length of the blade root section basic blade profile of the two-dimensional basic blade profile is between 0.8 and 1.0, and the value in this embodiment is 0.85.
In summary, according to the power turbine blade of the turboprop engine and the design method thereof provided by the embodiments, the angle of attack and the modeling parameters of the blade at the design point are reasonably selected, so that the angle of attack loss of the non-design point due to the large-range change of the rotating speed of the turbine is reduced, the performance of the non-design point of the turbine is considered on the premise of ensuring the performance of the design point and the structural strength of the blade, and the oil consumption of the whole engine in the full flight envelope is reduced.
The above description is only for the purpose of illustrating the preferred embodiments of the present application and is not to be construed as limiting the present application, and any modifications, equivalents, improvements, etc. made within the spirit and principle of the present application should be included in the scope of the present application.

Claims (9)

1. A design method for a power turbine blade of a turboprop engine is characterized by comprising the following steps:
step 1, selecting an initial attack angle as zero degree according to the design requirements of a power turbine at a design state point, and developing a primary turbine blade modeling design;
step 2, building a calculation model according to a preliminary turbine blade modeling design result, carrying out multi-state calculation analysis, and obtaining radial distribution of blade outlet airflow angles at typical state points, wherein the typical state points comprise high-altitude cruise state points and takeoff state points;
step 3, according to the radial distribution of the blade outlet airflow angle of the typical state point and the optimized inlet attack angle range, adjusting the geometric modeling parameters of the blade and developing the blade profile optimization design of the blade;
step 4, carrying out turbine characteristic calculation based on the optimized blades to obtain assessment parameters of different state points, wherein the assessment parameters comprise power turbine efficiency, power and physical flow of the different state points;
step 5, judging the assessment parameters, if the assessment parameters meet the design requirements, calculating the strength, vibration and service life of the rotor, and if the calculation results of the strength, vibration and service life meet the requirements, finishing the design of the turbine blade; and if any one of the calculation results of the efficiency, the power, the physical flow, the strength, the vibration and the service life of the power turbine does not meet the requirements, returning to the step 3, readjusting the modeling parameters of the blades, designing the blade profile of the blades, and performing calculation and analysis on the characteristics of the turbine again until the optimized blades meet all preset conditions.
2. The method of claim 1, wherein in step 1, the design condition point is selected as a high altitude cruise condition point, and the blade angle of attack i is determined by the following equation:
i＝β k -β；
wherein β is the blade inlet flow angle; beta is a beta k An angle is constructed for the blade inlet.
3. The method for designing a power turbine blade of a turboprop engine according to claim 1, wherein in step 1, when the preliminary turbine blade model design is developed, the inlet blade angle is determined according to the selected initial attack angle condition, two-dimensional basic blade profiles at different radial heights are constructed by preferably selecting geometric parameters including the outlet blade angle, the radius of the leading edge and the trailing edge, the consistency, the chord length and the degree of blockage of the trailing edge, and the basic blade profiles are radially superposed according to a certain stacking rule to form the three-dimensional preliminary turbine blade model.
4. The method as claimed in claim 1, wherein in step 3, the geometric modeling parameters of the blade are adjusted by increasing chord length, height, inlet and outlet wedge angle, or adjusting inlet structure angle β of the blade k In the manner of (a).
5. The method of claim 1, wherein the design condition point blade inlet configuration angle β is defined as the design condition point blade inlet configuration angle β k Satisfies the following conditions: beta is a beta k ＝β des +i des The said i des The angle of attack is between 3 DEG and 5 DEG, beta des The blade inlet flow angle is the design condition point.
6. A power turbine blade of a turboprop engine is characterized by comprising 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 plurality of two-dimensional basic blade profiles along the radial direction according to a certain rule; the three-dimensional entity of the power turbine blade of the turboprop engine is formed by stacking N two-dimensional basic blade profiles along the radius direction by a certain rule, wherein N is a natural number and is not less than 3; the two-dimensional basic blade profile comprises a front edge, a tail edge, a blade basin profile line and a blade back profile line, and is designed by selecting different inlet and outlet construction angles, a front edge wedge angle, a tail edge bending angle, a front edge radius, a tail edge radius, consistency, chord length and tail edge blockage degree;
each two-dimensional foundation blade profile comprises a blade root section foundation blade profile, a blade tip section foundation blade profile and a plurality of blade leaf section foundation blade profiles, wherein the blade root section foundation blade profile is defined as the two-dimensional foundation blade profile at a blade height which is 5% of the distance 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 position 95% of the blade height away from the upper surface of the lower edge plate; the blade-in-section basic blade profile is defined as a two-dimensional basic blade profile with the blade profile height between the blade root and the blade tip section;
radius value r of the leading edge 1 Between 0.55 and 0.9mm, the radius r of the trailing edge t And is between 0.25 mm and 0.35 mm.
7. The turboprop engine power turbine blade of claim 6, wherein a leading edge wedge angle of the two-dimensional base blade profile is between 32-48 °.
8. The turboprop engine power turbine blade of claim 6, wherein the ratio of the root cross-sectional root profile area to the tip cross-sectional root profile area of the two-dimensional root profile is between 1.8 and 2.5.
9. The turboprop engine power turbine blade according to claim 6, wherein a ratio of a tip section airfoil chord length to a root section airfoil chord length of the two-dimensional airfoil is between 0.8 and 1.0.
CN202210758915.3A 2022-06-29 2022-06-29 Power turbine blade of turboprop engine and design method thereof Pending CN115098966A (en)

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115982892A (en) * 2023-03-17 2023-04-18 潍柴动力股份有限公司 Blade design method, blade and related equipment
CN116090137A (en) * 2023-03-17 2023-05-09 潍柴动力股份有限公司 Turbine blade, turbine blade design method and apparatus
CN116401767A (en) * 2023-04-18 2023-07-07 中国航发湖南动力机械研究所 Design method of blade body super-flying-off blade

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115982892A (en) * 2023-03-17 2023-04-18 潍柴动力股份有限公司 Blade design method, blade and related equipment
CN116090137A (en) * 2023-03-17 2023-05-09 潍柴动力股份有限公司 Turbine blade, turbine blade design method and apparatus
CN116401767A (en) * 2023-04-18 2023-07-07 中国航发湖南动力机械研究所 Design method of blade body super-flying-off blade
CN116401767B (en) * 2023-04-18 2024-06-04 中国航发湖南动力机械研究所 Design method of blade body super-flying-off blade

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