CN113836633B - Gas turbine inlet guide vane profile design method, gas turbine and aircraft engine - Google Patents
Gas turbine inlet guide vane profile design method, gas turbine and aircraft engine Download PDFInfo
- Publication number
- CN113836633B CN113836633B CN202110936959.6A CN202110936959A CN113836633B CN 113836633 B CN113836633 B CN 113836633B CN 202110936959 A CN202110936959 A CN 202110936959A CN 113836633 B CN113836633 B CN 113836633B
- Authority
- CN
- China
- Prior art keywords
- inlet guide
- guide vane
- profile
- inlet
- angle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Classifications
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/10—Geometric CAD
- G06F30/15—Vehicle, aircraft or watercraft 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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
-
- 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
-
- 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
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/20—Design reuse, reusability analysis or reusability optimisation
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Geometry (AREA)
- Theoretical Computer Science (AREA)
- General Physics & Mathematics (AREA)
- General Engineering & Computer Science (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- Mathematical Optimization (AREA)
- Computational Mathematics (AREA)
- Mathematical Analysis (AREA)
- Pure & Applied Mathematics (AREA)
- Aviation & Aerospace Engineering (AREA)
- Automation & Control Theory (AREA)
- Mechanical Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a design method of a gas turbine inlet guide vane profile, which comprises the steps of acquiring the distribution condition and characteristic parameters of isentropic Mach numbers of the surface of the inlet guide vane profile through simulation, and acquiring a low-load region and an inlet attack angle of the surface of the inlet guide vane profile; the method comprises the steps of cutting off a low-load area at the front edge of an inlet guide vane profile by adopting a front edge cutting-off mode to construct a high-load inlet guide vane profile, simultaneously increasing an inlet attack angle of the inlet guide vane profile to increase the front load of the inlet guide vane profile, amplifying and reconstructing parameters of the inlet guide vane profile in order to keep the axial installation space of a vane in an aeroengine unchanged, and accordingly, the pneumatic parameters on the surface of the inlet guide vane are reasonably distributed, the pneumatic loss is less, after the inlet guide vane is amplified, the blade grid distance of a turbine is amplified according to a corresponding proportion, and the number of the blades of the turbine is correspondingly reduced according to a proportion, so that the cold air quantity used for the inlet guide vane is reduced, and meanwhile, the manufacturing cost is reduced.
Description
Technical Field
The invention relates to the technical field of design and processing of parts of aero-engines, in particular to a design method of a blade profile of an inlet guide vane of a gas turbine. In addition, the invention also relates to a gas turbine which is designed and manufactured by the design method of the gas turbine inlet guide vane profile. Furthermore, the invention relates to an aircraft engine comprising a gas turbine as described above.
Background
In the existing middle and small-sized aviation gas turbine engine, a turbine is one of core components of a power device of the gas turbine engine, is an impeller machine which converts the energy of high-temperature and high-pressure gas into kinetic energy and mechanical energy, and has the main function of turning and expanding the high-temperature and high-pressure gas flowing through a guide blade row from a turbine inlet so as to convert the gas pressure potential energy into the gas kinetic energy, so that the high-speed gas performs mechanical work through a working blade row and is provided for a gas compressor by a gas turbine shaft for air compression. Therefore, the quality of the turbine design determines the overall design level of the engine, and the development level of the turbine design plays a crucial role in the development of the aircraft engine.
With the improvement of the performance and the power-to-weight ratio of a turbine engine, the cycle parameters are continuously improved, the temperature in front of the turbine is continuously increased, and for the gas turbine, the temperature of inlet gas reaches or even exceeds 2000K, so that the inlet guide vane needs a large amount of cold air for cooling, the cold air amount of the inlet guide vane approximately accounts for 10% of the inlet flow of the gas compressor, but if the inlet guide vane carries out bleed air cooling through a large amount of gas films, on one hand, the pneumatic loss of the inlet guide vane is increased, and the transmission efficiency of the turbine is reduced; on the other hand, the amount of available air for cooling and blending the combustion chamber is correspondingly reduced. In addition, manufacturing a large number of film holes in the body of the inlet guide vane increases process difficulty and cost.
Disclosure of Invention
The invention provides a design method of a gas turbine inlet guide vane profile, a gas turbine and an aero-engine, which aim to solve the problems that the existing aero-engine has excessive air entraining quantity of a gas turbine inlet guide vane, increases the pneumatic loss of the inlet guide vane, reduces the transmission efficiency of the turbine, and correspondingly reduces the cooling and mixing available air quantity of a combustion chamber; the technical problem of high manufacturing cost is caused by the fact that a large number of air film holes are manufactured in the blade body of the inlet guide blade to increase process difficulty.
According to one aspect of the invention, a method for designing a gas turbine inlet guide vane profile is provided, comprising the steps of: a. acquiring characteristic parameters of the blade profile of the inlet guide blade and the isentropic Mach number distribution condition of the surface of the blade profile of the inlet guide blade; b. acquiring an inlet attack angle of the inlet guide vane profile and a low-load region of the inlet guide vane profile surface according to characteristic parameters of the inlet guide vane profile and the isentropic Mach number distribution condition of the inlet guide vane profile surface, wherein the low-load region of the inlet guide vane profile surface is positioned at the front edge of the inlet guide vane profile; c. setting a preset proportion and a preset angle according to design requirements, cutting off the axial width of the preset proportion in the front edge of the inlet guide vane profile, and simultaneously increasing the inlet attack angle of the inlet guide vane profile by the preset angle; d. obtaining the amplification factor of the blade profile of the inlet guide vane according to a preset proportion; e. and amplifying the blade profile of the inlet guide blade according to the amplification factor of the blade profile of the inlet guide blade, and carrying out parametric reconstruction to keep the inlet attack angle and the tail edge thickness of the amplified blade profile of the inlet guide blade unchanged so as to obtain the blade profile of the inlet guide blade of the gas turbine required by the design requirement.
Further, the step a specifically comprises: carrying out motion simulation analysis on the assembly and operation conditions of the inlet guide vane; acquiring the axial width, inlet structure angle, outlet airflow angle and trailing edge thickness of an inlet guide vane profile; obtaining the distribution condition of the isentropic Mach number of the blade profile surface of the inlet guide vane;
further, the step b specifically comprises: acquiring an inlet attack angle according to an inlet structure angle and an outlet airflow angle of the inlet guide vane profile, wherein the inlet attack angle is the angle difference between the inlet structure angle and the outlet airflow angle; and acquiring a low-load region of the surface of the inlet guide vane profile according to the entropy Mach number distribution condition of the surface of the inlet guide vane profile, wherein the low-load region of the surface of the inlet guide vane profile is positioned at the front edge of the inlet guide vane profile.
Further, the inlet attack angle of the inlet guide vane profile is increased by a preset angle, specifically: and keeping the outlet airflow angle of the inlet guide vane profile unchanged, so that the inlet structure angle of the inlet guide vane profile is increased by a preset angle.
Further, step e specifically comprises: amplifying the blade profile of the inlet guide blade according to the amplification factor of the blade profile of the inlet guide blade; and carrying out parametric reconstruction on the blade profile of the inlet guide vane to ensure that the inlet structure angle, the outlet airflow angle and the tail edge thickness of the blade profile of the inlet guide vane after amplification are consistent with those of the blade profile of the inlet guide vane before amplification.
Further, step d specifically includes: obtaining a calculation formula according to a preset proportion and a magnification factor; the calculation formula is as follows: 1/(1-x) = y (wherein x is a preset ratio and y is a magnification); and substituting the preset proportion into a calculation formula, and obtaining the magnification factor through calculation.
Further, the preset ratio is 17%, and the preset angle is 40-60 °.
Further, the preset ratio is 41.5%, and the preset angle is 40-60 °.
According to another aspect of the invention, a gas turbine is further provided, and the gas turbine is designed and manufactured by adopting the gas turbine inlet guide vane profile design method.
According to another aspect of the invention, there is also provided an aircraft engine comprising a gas turbine as described above.
The invention has the following beneficial effects:
according to the design method of the gas turbine inlet guide vane profile, the isentropic Mach number distribution condition and the characteristic parameters of the surface of the inlet guide vane profile are obtained through simulation, and the low-load area and the inlet attack angle of the surface of the inlet guide vane profile are obtained; the method comprises the steps of cutting off a low-load area at the front edge of an inlet guide vane profile by adopting a front edge cutting-off mode to construct a high-load inlet guide vane profile, simultaneously increasing an inlet attack angle of the inlet guide vane profile to increase the front load of the inlet guide vane profile, amplifying and reconstructing parameters of the inlet guide vane profile in order to keep the axial installation space of a vane in an aeroengine unchanged, and accordingly, the pneumatic parameters on the surface of the inlet guide vane are reasonably distributed, the pneumatic loss is less, after the inlet guide vane is amplified, the blade grid distance of a turbine is amplified according to a corresponding proportion, and the number of the blades of the turbine is correspondingly reduced according to a proportion, so that the cold air quantity used for the inlet guide vane is reduced, and meanwhile, the manufacturing cost is reduced. The invention can design, process and manufacture the gas turbine inlet guide vane in various aeroengines according to the design method of the gas turbine inlet guide vane profile, under the condition that the total cold air quantity is not changed, because the number of the vanes is reduced, the cold air quantity used for a single inlet guide vane is increased, the metal wall temperature of the inlet guide vane is reduced, the research shows that the metal wall temperature is reduced by 25K, and the service life of the inlet guide vane can be increased by 2 times; or, under the condition of maintaining the temperature of the metal wall of the inlet guide vane unchanged, the total gas consumption of the inlet guide vane is reduced due to the reduction of the number of the vanes, so that the transmission efficiency of the gas turbine and the comprehensive performance of the aero-engine are improved.
In addition to the objects, features and advantages described above, other objects, features and advantages of the present invention are also provided. The present invention will be described in further detail below with reference to the drawings.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the invention and, together with the description, serve to explain the invention and not to limit the invention. In the drawings:
FIG. 1 is a block flow diagram of the steps of a gas turbine inlet guide vane airfoil design method of a preferred embodiment of the present invention;
FIG. 2 is an isentropic Mach number distribution plot for a fundamental airfoil surface of the present invention;
FIG. 3 is a schematic illustration of step c of the gas turbine inlet guide vane profile design method of the preferred embodiment of the present invention;
FIG. 4 is a comparative schematic illustration of inlet guide vane profiles of three aspects of the present invention;
FIG. 5 is a plot of a comparison of isentropic Mach number distributions for the airfoil surfaces of inlet guide vanes according to three aspects of the present invention;
FIG. 6 is a comparison graph of the total pressure loss coefficient distribution downstream of the inlet guide vane airfoil for three aspects of the present invention;
FIG. 7 is a graph comparing the total pressure loss coefficient versus inlet angle of attack for the inlet guide vane airfoil of three aspects of the present invention;
FIG. 8 is a schematic structural view of an inlet guide vane airfoil in a gas turbine in accordance with a preferred embodiment of the present invention.
Illustration of the drawings:
1. leaf basin; 2. leaf back; 3. a leading edge; 4. a trailing edge; 5. an inlet construction angle; 6. an outlet airflow angle; 7. and (4) an inlet attack angle.
Detailed Description
The embodiments of the invention will be described in detail below with reference to the drawings, but the invention can be practiced in many different ways, which are defined and covered by the following.
FIG. 1 is a block flow diagram of the steps of a method of designing a gas turbine inlet guide vane airfoil in accordance with a preferred embodiment of the present invention; FIG. 2 is an isentropic Mach number distribution plot for a fundamental airfoil surface of the present invention; FIG. 3 is a schematic illustration of step c of the gas turbine inlet guide vane profile design method of the preferred embodiment of the present invention; FIG. 4 is a comparative schematic of inlet guide vane profiles of three aspects of the present invention; FIG. 5 is a plot of a comparison of isentropic Mach number distributions for inlet guide vane airfoil surfaces of three aspects of the present invention; FIG. 6 is a comparison graph of the distribution of the total pressure loss coefficient downstream of the inlet guide vane airfoil for three versions of the present invention; FIG. 7 is a graph comparing the total pressure loss coefficient versus inlet angle of attack for the inlet guide vane airfoil of three aspects of the present invention; FIG. 8 is a schematic view of the inlet guide vane airfoil in the gas turbine of the preferred embodiment of the present invention.
As shown in fig. 1, the method for designing the gas turbine inlet guide vane profile of the present embodiment includes the following steps: a. acquiring characteristic parameters of the blade profile of the inlet guide blade and the isentropic Mach number distribution condition of the surface of the blade profile of the inlet guide blade; b. acquiring an inlet attack angle 7 of the inlet guide vane profile and a low-load region of the inlet guide vane profile surface according to characteristic parameters of the inlet guide vane profile and the isentropic Mach number distribution condition of the inlet guide vane profile surface, wherein the low-load region of the inlet guide vane profile surface is positioned at the front edge 3 of the inlet guide vane profile; c. setting a preset proportion and a preset angle according to design requirements, cutting off the axial width of the preset proportion in the leading edge 3 of the inlet guide vane profile, and simultaneously increasing the inlet attack angle 7 of the inlet guide vane profile by the preset angle; d. obtaining the amplification factor of the blade profile of the inlet guide vane according to a preset proportion; e. and amplifying the blade profile of the inlet guide blade according to the amplification factor of the blade profile of the inlet guide blade, and carrying out parametric reconstruction to keep the thickness of an inlet attack angle 7 and a tail edge 4 of the amplified blade profile of the inlet guide blade unchanged so as to obtain the blade profile of the inlet guide blade of the gas turbine required by the design requirement. In the prior art, zweifel load coefficients of inlet guide vanes of gas turbines currently in service are generally between 0.6 and 0.8, and loads borne by single inlet blades are relatively low. Specifically, the design method of the gas turbine inlet guide vane profile obtains the distribution condition and the characteristic parameters of the isentropic Mach number of the inlet guide vane profile surface through motion simulation, and obtains a low-load region and an inlet attack angle 7 of the inlet guide vane profile surface; a low-load area of a front edge 3 of an inlet guide vane profile is cut off by adopting a mode of cutting off the front edge 3 to construct a high-load inlet guide vane profile, an inlet attack angle 7 of the inlet guide vane profile is increased to increase the front load of the inlet guide vane profile, and in order to keep the axial installation space of the vanes in the aeroengine unchanged, the inlet guide vane profile is amplified and subjected to parameter reconstruction, so that the pneumatic parameters on the surface of the inlet guide vane are reasonably distributed, the pneumatic loss is less, after the inlet guide vane is amplified, the blade grid distance of a turbine is amplified according to a corresponding proportion, and the number of the blades of the turbine is correspondingly reduced according to the proportion, so that the cold air volume for the inlet guide vane is reduced, and the manufacturing cost is reduced. The invention can design, process and manufacture the gas turbine inlet guide vane in various aeroengines according to the design method of the gas turbine inlet guide vane profile, under the condition that the total cold air quantity is not changed, because the number of the vanes is reduced, the cold air quantity used for a single inlet guide vane is increased, the metal wall temperature of the inlet guide vane is reduced, researches show that the metal wall temperature is reduced by 25K, and the service life of the inlet guide vane can be increased by 2 times; or, under the condition of maintaining the temperature of the metal wall of the inlet guide vane unchanged, the total gas consumption of the inlet guide vane is reduced due to the reduction of the number of the vanes, so that the transmission efficiency of the gas turbine and the comprehensive performance of the aero-engine are improved. It should be understood that gas turbine inlet guide vane blade rows differ significantly from other blade rows, primarily in that: 1. the outlet airflow angle 6 is close to the axial direction, and the channel convergence is much higher than that of other blade rows: 2. the inner and outer flow passages are commonly contraction passages, and the convergence of the whole cascade passage is further enhanced. While other blade discharge ducts typically employ expanding channels to accommodate the reduction in gas density. Therefore, due to the difference of the geometrical characteristics, a larger inlet-outlet area ratio of the inlet guide vane is caused, so that the inlet mach number of the turbine inlet guide vane is lower and is generally about 0.1. And the low Mach number flow is relatively insensitive to the geometry of the leading edge 3 of the inlet guide vane, and the blade profile loss is relatively insensitive to the change of the inlet attack angle 7, so that the inlet attack angle 7 can be increased to increase the load of the leading edge 3 of the inlet guide vane.
As shown in fig. 1 and fig. 2, in this embodiment, step a specifically includes: carrying out motion simulation analysis on the assembly and operation conditions of the inlet guide vane; acquiring the axial width, the inlet structure angle 5, the outlet airflow angle 6 and the thickness of a tail edge 4 of an inlet guide vane profile; and acquiring the distribution condition of the isentropic Mach number of the blade profile surface of the inlet guide vane. Specifically, the inner section of an inlet guide vane of a gas turbine of an engine in service in the prior art is set as a basic vane profile (scheme A), and the basic vane profile is subjected to motion simulation analysis to obtain the axial width, the inlet structure angle 5, the outlet airflow angle 6 and the thickness of a trailing edge 4 of the basic vane profile and obtain the isentropic Mach number distribution condition of the surface of the basic vane profile.
As shown in fig. 2, in this embodiment, step b specifically includes: acquiring an inlet attack angle 7 according to an inlet structure angle 5 and an outlet airflow angle 6 of the inlet guide vane profile, wherein the inlet attack angle 7 is the angle difference between the inlet structure angle 5 and the outlet airflow angle 6; and acquiring a low-load region of the surface of the blade profile of the inlet guide vane according to the distribution condition of the entropy Mach number of the surface of the blade profile of the inlet guide vane, wherein the low-load region of the surface of the blade profile of the inlet guide vane is positioned at the front edge 3 of the blade profile of the inlet guide vane. In particular, it can be seen from FIG. 2 that the loads in the region of the leading edge 3 of the surface of the base blade profile are low, and that the Zweifel load coefficient of the base blade profile is 0.64. And the Zweifel load coefficients of the blade type schemes B and C which are newly designed by adopting the invention are respectively 0.96 and 1.28, and the corresponding blade number can be respectively reduced by 33 percent and 50 percent compared with the scheme A (24 sheets).
As shown in fig. 3, in the present embodiment, the inlet attack angle 7 of the inlet guide vane profile is increased by a preset angle, which specifically includes: and keeping the outlet airflow angle 6 of the inlet guide vane profile unchanged, so that the inlet structure angle 5 of the inlet guide vane profile is increased by a preset angle.
As shown in fig. 4, in this embodiment, step e specifically includes: amplifying the blade profile of the inlet guide blade according to the amplification factor of the blade profile of the inlet guide blade; and carrying out parametric reconstruction on the blade profile of the inlet guide blade, so that the thicknesses of an inlet structure angle 5, an outlet airflow angle 6 and a tail edge 4 of the blade profile of the inlet guide blade after amplification are consistent with the thicknesses of the inlet structure angle 5, the outlet airflow angle 6 and the tail edge 4 of the blade profile of the inlet guide blade before amplification.
In this embodiment, step d specifically includes: obtaining a calculation formula according to a preset proportion and a magnification factor; the calculation formula is as follows: 1/(1-x) = y (wherein x is a preset ratio and y is a magnification); and substituting the preset proportion into a calculation formula, and obtaining the magnification factor through calculation.
As shown in fig. 3 and 4, in the present embodiment, the preset ratio is set to 17% according to design requirements, the preset angle is 40 ° to 60 °, and preferably, the preset angle is 50 °, the scheme B is set.
In another embodiment, as shown in fig. 3 and 4, the preset ratio is set to 41.5% according to design requirements, the preset angle is 40-60 °, and preferably, the scheme C is set when the preset angle is 50 °.
As shown in FIG. 3 and FIG. 4, the scheme B and the scheme C are respectively realized by increasing the inlet attack angle 7 and the truncation of the leading edge 3 and axial scale enlargement, the leading edge 3 of the scheme B and the scheme C truncates 17 percent and 41.5 percent of the axial width respectively, the inlet structure angle 5 is increased by 50 degrees, after the leading edge 3 is truncated, the remaining part at the rear part is enlarged and parameterized and reconstructed to be consistent with the characteristic parameters of the axial width, the thickness of the trailing edge 4, the outlet airflow angle 6 and the like of the basic blade profile, the blade grid distance is correspondingly enlarged in proportion, and the number of blades can be correspondingly reduced.
As shown in fig. 5, the front zone load is significantly increased for case B and case C compared to case a.
As shown in fig. 6, the peak loss is greatest for case C, but less for case B, because the forward peak of the high-lift airfoil suction surface mach number causes the length of the back 2 diffuser to increase, resulting in increased airfoil loss for a single blade. But the number of the blades in the scheme A is the largest, and the high-loss area in the grid pitch influenced by the blade profile boundary layer and the trail accounts for the largest percentage and accounts for about 83.6 percent of the whole channel; the high loss area fraction for case B and case C was about 58.1% and 43.9%. The total pressure loss coefficients of the three scheme cascade channels are obtained by integration along the horizontal axis and are respectively 6.96%, 5.39% and 4.37%. The analysis result shows that the high-load blade profile constructed by adopting the mode of cutting off the front edge 3 and increasing the inlet attack angle 7 can reduce the blade cascade loss while reducing the number of blades.
As shown in FIG. 7, the increase in the flow angle results in an increase in the total pressure loss, which is relatively large in the high-load blade profile (case B and case C), but generally, the change in the flow angle in this range has little effect on the total pressure loss. This corresponds to low inlet mach numbers that are not as sensitive to angle of attack.
As shown in fig. 8, the gas turbine of the present embodiment is designed and manufactured by using the above-described gas turbine inlet guide vane profile design method. Specifically, the gas turbine includes inlet guide vanes including a bucket 1, a bucket back 2, a leading edge 3, a trailing edge 4, an inlet configuration angle 5, an outlet flow angle 6, and an inlet angle of attack 7.
The aircraft engine of the embodiment comprises the gas turbine. Particularly, the inlet guide vane of the gas turbine has high load and few blades, and can reduce the cold air consumption and the manufacturing and production cost of the aircraft engine.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. A design method for the blade profile of the inlet guide vane of a gas turbine is characterized in that,
the method comprises the following steps:
a. acquiring characteristic parameters of the blade profile of the inlet guide blade and the isentropic Mach number distribution condition of the surface of the blade profile of the inlet guide blade;
b. acquiring an inlet attack angle (7) of the inlet guide vane profile and a low-load region of the inlet guide vane profile surface according to characteristic parameters of the inlet guide vane profile and the isentropic Mach number distribution condition of the inlet guide vane profile surface, wherein the low-load region of the inlet guide vane profile surface is positioned at the front edge (3) of the inlet guide vane profile;
c. setting a preset proportion and a preset angle according to design requirements, cutting off the axial width of the preset proportion in the front edge (3) of the inlet guide vane profile, and simultaneously increasing the inlet attack angle (7) of the inlet guide vane profile by the preset angle;
d. acquiring the amplification factor of the blade profile of the inlet guide vane according to a preset proportion;
e. and amplifying the blade profile of the inlet guide blade according to the amplification factor of the blade profile of the inlet guide blade, and carrying out parametric reconstruction to keep the thickness of an inlet attack angle (7) and a tail edge (4) of the amplified blade profile of the inlet guide blade unchanged so as to obtain the blade profile of the inlet guide blade of the gas turbine required by the design requirement.
2. The gas turbine inlet guide vane airfoil design method of claim 1,
the step a specifically comprises the following steps:
performing motion simulation analysis on the assembly and operation conditions of the inlet guide vane;
acquiring the axial width, the inlet structure angle (5), the outlet airflow angle (6) and the thickness of a tail edge (4) of an inlet guide vane profile;
and acquiring the distribution condition of the isentropic Mach number of the blade profile surface of the inlet guide vane.
3. The gas turbine inlet guide vane airfoil design method of claim 2,
the step b is specifically as follows:
acquiring an inlet attack angle (7) according to an inlet structure angle (5) and an outlet airflow angle (6) of the inlet guide vane profile, wherein the inlet attack angle (7) is the angle difference between the inlet structure angle (5) and the outlet airflow angle (6);
and acquiring a low-load region of the surface of the blade profile of the inlet guide vane according to the distribution condition of the entropy Mach number of the surface of the blade profile of the inlet guide vane, wherein the low-load region of the surface of the blade profile of the inlet guide vane is positioned at the front edge (3) of the blade profile of the inlet guide vane.
4. The gas turbine inlet guide vane airfoil design method of claim 3,
the angle of attack (7) of the inlet guide vane profile is increased by a preset angle, and the angle is specifically as follows:
and keeping an outlet airflow angle (6) of the inlet guide vane profile unchanged, and increasing an inlet structure angle (5) of the inlet guide vane profile by a preset angle.
5. The gas turbine inlet guide vane airfoil design method of claim 4,
the step e is specifically as follows:
amplifying the blade profile of the inlet guide blade according to the amplification factor of the blade profile of the inlet guide blade;
and carrying out parametric reconstruction on the blade profile of the inlet guide blade, so that the thickness of an inlet structure angle (5), an outlet airflow angle (6) and a trailing edge (4) of the blade profile of the inlet guide blade after amplification is consistent with the thickness of the inlet structure angle (5), the outlet airflow angle (6) and the trailing edge (4) of the blade profile of the inlet guide blade before amplification.
6. The gas turbine inlet guide vane airfoil design method of any one of claims 1-5,
the step d is specifically as follows:
obtaining a calculation formula according to a preset proportion and a magnification factor;
the calculation formula is as follows: 1/(1-x) = y (wherein x is a preset ratio and y is a magnification);
and substituting the preset proportion into a calculation formula, and obtaining the magnification factor through calculation.
7. The method of designing a gas turbine inlet guide vane airfoil as claimed in any one of claims 1 to 5, wherein the predetermined ratio is 17% and the predetermined angle is 40 ° to 60 °.
8. The method of designing a gas turbine inlet guide vane airfoil as claimed in any one of claims 1 to 5, wherein the predetermined ratio is 41.5% and the predetermined angle is 40 ° to 60 °.
9. A gas turbine, characterized in that the gas turbine is designed and manufactured according to the gas turbine inlet guide vane profile design method of any one of claims 1 to 8.
10. An aircraft engine comprising a gas turbine according to claim 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110936959.6A CN113836633B (en) | 2021-08-16 | 2021-08-16 | Gas turbine inlet guide vane profile design method, gas turbine and aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110936959.6A CN113836633B (en) | 2021-08-16 | 2021-08-16 | Gas turbine inlet guide vane profile design method, gas turbine and aircraft engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113836633A CN113836633A (en) | 2021-12-24 |
| CN113836633B true CN113836633B (en) | 2022-12-02 |
Family
ID=78960721
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110936959.6A Active CN113836633B (en) | 2021-08-16 | 2021-08-16 | Gas turbine inlet guide vane profile design method, gas turbine and aircraft engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113836633B (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114542216B (en) * | 2022-02-25 | 2024-06-14 | 中国航发沈阳发动机研究所 | Turbine support plate blade design method with supporting and guiding functions and blade |
| CN115098959B (en) * | 2022-05-29 | 2023-05-12 | 中国船舶重工集团公司第七0三研究所 | Gas turbine high-pressure turbine guide vane design method |
| 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 |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104895618A (en) * | 2015-04-10 | 2015-09-09 | 中国科学院工程热物理研究所 | Super-high load low pressure turbine blade, high load low pressure turbine and aviation gas turbine engine |
| CN106050319A (en) * | 2016-06-14 | 2016-10-26 | 中国科学院工程热物理研究所 | Large-attack-angle containment turbine blade used for aero gas turbine engine |
-
2021
- 2021-08-16 CN CN202110936959.6A patent/CN113836633B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104895618A (en) * | 2015-04-10 | 2015-09-09 | 中国科学院工程热物理研究所 | Super-high load low pressure turbine blade, high load low pressure turbine and aviation gas turbine engine |
| CN106050319A (en) * | 2016-06-14 | 2016-10-26 | 中国科学院工程热物理研究所 | Large-attack-angle containment turbine blade used for aero gas turbine engine |
Non-Patent Citations (1)
| Title |
|---|
| 高负荷涡轮扇形叶栅变攻角气动试验研究;李国强等;《节能技术》;20191126(第06期);全文 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113836633A (en) | 2021-12-24 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN113836633B (en) | Gas turbine inlet guide vane profile design method, gas turbine and aircraft engine | |
| US6881029B2 (en) | Casing, a compressor, a turbine, and a combustion turbine engine including such a casing | |
| EP1939399B1 (en) | Axial flow turbine assembly | |
| US8573941B2 (en) | Tandem blade design | |
| EP1505302B1 (en) | Compressor airfoil | |
| EP2746536A1 (en) | Rotor stage of a turbine | |
| CN210859342U (en) | Gas compressor and air guide groove flow guide control structure thereof | |
| US20070297904A1 (en) | Compressor Of A Gas Turbine And Gas Turbine | |
| CN114718659B (en) | Turbine blade tip clearance flow control method coupling radial ribs and circumferential grooves | |
| CN114444331B (en) | Stage characteristic matching method of multistage axial flow compressor | |
| CN115186398B (en) | Method for determining key angle parameters of inlet guide vane modeling of axial flow compressor | |
| CN113283198A (en) | Method, system and terminal for optimizing treatment of compressor casing and improving stability margin | |
| CN109113870B (en) | Diffuser, compressor and gas turbine | |
| Jin et al. | Design of a highly loaded transonic two-stage fan using swept and bowed blading | |
| CN115929694A (en) | Centrifugal compressor diffuser and centrifugal compressor | |
| CN115263436A (en) | Transonic turbine rotor blade, turbine rotor and turbine | |
| CN115169032A (en) | Design method of gas turbine working blade with blade tip groove and winglet composite structure | |
| GB2572793A (en) | Turbine component | |
| CN114547990A (en) | Coupling type casing design method for improving flow of blade top of gas compressor | |
| Nayak et al. | Criteria for selection of solidity in design of contra rotating fan stage | |
| CN111120406A (en) | Method for increasing air volume of centrifugal fan | |
| US20240003353A1 (en) | Compressor for an engine | |
| CN113586164B (en) | High-load high-pressure turbine rotor blade suitable for medium-thrust aero-engine | |
| KR20110054368A (en) | Flow stabilization structure of compressor for a gas turbine engine | |
| CN113898472B (en) | Gas compressor retrofitting method and gas compressor |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |