CN114840921A - Method for designing cooling blade of high-pressure turbine at outlet of combustion chamber - Google Patents
Method for designing cooling blade of high-pressure turbine at outlet of combustion chamber Download PDFInfo
- Publication number
- CN114840921A CN114840921A CN202210425143.1A CN202210425143A CN114840921A CN 114840921 A CN114840921 A CN 114840921A CN 202210425143 A CN202210425143 A CN 202210425143A CN 114840921 A CN114840921 A CN 114840921A
- Authority
- CN
- China
- Prior art keywords
- temperature
- blade
- design
- cooling
- outlet
- 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.)
- Granted
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
-
- 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
- G06F2119/08—Thermal analysis or thermal 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
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Geometry (AREA)
- Theoretical Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Evolutionary Computation (AREA)
- Computer Hardware Design (AREA)
- General Engineering & Computer Science (AREA)
- Pure & Applied Mathematics (AREA)
- Mathematical Optimization (AREA)
- Mathematical Analysis (AREA)
- Computational Mathematics (AREA)
- Automation & Control Theory (AREA)
- Aviation & Aerospace Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the field of design of engine blades, and relates to a method for designing a cooling blade of a high-pressure turbine in consideration of the outlet temperature of a combustion chamber, which comprises the steps of selecting a state point with the highest cooling effect as a design point to provide enough margin for subsequent design, then carrying out grid division on the outlet section of the combustion chamber, counting the outlet temperature distribution of the combustion chamber at the design point, giving the inlet design temperature of the high-pressure turbine at different radiuses, carrying out appearance design, gas side pneumatic and heat exchange calculation according to the temperature, then carrying out region division on the cooling blade of the high-pressure turbine, calculating the heat exchange coefficients required by the cold air sides at different regions, obtaining the corresponding relation between the inner cavity cooling requirement of the cooling blade and the outlet temperature of the combustion chamber, carrying out targeted design according to the heat exchange coefficients at the cold air sides at different regions, carrying out quantitative analysis and failure rate calculation according to the uncertainty temperature distribution probability function of the combustion chamber outlet, And judging, and finally optimizing the blade structure to complete the design.
Description
Technical Field
The application belongs to the field of engine blade design, and particularly relates to a high-pressure turbine cooling blade design method considering the outlet temperature of a combustion chamber.
Background
Due to the limited space and millisecond-level mixing time in the main combustion chamber, complex unsteady flow, uneven oil-gas mixing and other factors, the actual outlet temperature of the combustion chamber is uneven, and local high-temperature hot spots appear, which far exceed the melting point of the downstream turbine blade high-temperature alloy. The deterioration of the combustor exit temperature profile will result in turbine blade overheating, erosion or damage, etc., and is critical to turbine blade reliability and life. The research shows that: when the combustor exit temperature distribution exceeds the ideal temperature distribution by 30K, the life of the turbine component will be reduced by 50%.
In order to evaluate the nonuniformity of the temperature distribution at the outlet of the combustion chamber of an aircraft engine and the difference between the nonuniformity and the expected temperature distribution shape of a turbine in the prior art, a combustion chamber outlet nonuniformity coefficient OTDF and an outlet radial temperature distribution coefficient RTDF are often adopted to respectively represent the overall uniformity and the radial uniformity of the temperature distribution of combustion gas at the outlet of the combustion chamber.
OTDF＝(T 4,max -T 4 )/(T 4 -T 3 )………………………………(1)
RTDF＝(T 4,avr -T 4 )/(T 4 -T 3 )………………………………(2)
Wherein:
T 4,max -maximum temperature of combustion gas at the outlet of the combustion chamber.
T 4 -average total temperature of combustion gas at the outlet of the combustion chamber.
T 3 -average total temperature of combustor inlet gas flow.
T 4,avr -the total temperature is averaged in the circumferential direction to a rear radial maximum.
In the current design of turbine cooling blades, for guide (rotor) blades, the temperature of combustion chamber outlet gas is generally averaged along the circumferential direction to obtain radial temperature distribution, and on the basis, the influence of OTDF (RTDF) is corrected on all radial gas recovery temperatures in a multiplying factor or increment mode.
T p ＝OTDF(T 4i -T 3 )+T 4i (guide vane) … … … … … … … … … … … … (3)
T p ＝RTDF(T 4 -T 3 )+T 4i (rotor) … … … … … … … … … … … … (4)
Wherein:
T p -blade inlet peak temperature.
T 4i The average total gas temperature at the inlet of the blade (guide vane)/the relative average total gas temperature at the inlet (rotor).
T g -gas recovery temperature.
T′ g -corrected gas recovery temperature
According to the gas temperature T 'after considering RTDF and OTDF' g The turbine blade is designed and the temperature field is evaluated according to the temperature.
According to the existing high-pressure turbine blade design and evaluation method, the consideration of OTDF and RTDF is not radially distinguished, the design pertinence is not strong, and the blade cooling air is not reasonably utilized, so that the following problems can be brought: under the condition of limited cold air consumption, the wall temperature of the hot spot position is overhigh; or in order to ensure that the wall temperatures of all positions of the blades meet the requirements, the amount of cold air is increased, and the efficiency of the engine is further influenced.
Therefore, how to take the non-uniformity of the outlet temperature of the combustion chamber into consideration is a problem to be solved by carrying out the targeted design of the cooling blades of the high-pressure turbine so as to improve the cold air utilization rate of the blades.
Disclosure of Invention
The application aims to provide a design method of a high-pressure turbine cooling blade considering the temperature of an outlet of a combustion chamber, and the method is used for solving the problems that in the prior art, the temperature field of the blade is only distinguished in the circumferential direction, so that the radial design cannot meet the cooling requirements of radial positions, and cold air cannot be reasonably utilized.
The technical scheme of the application is as follows: a method of designing a high pressure turbine cooling blade taking into account combustor exit temperature, comprising: calculating the cooling effect eta required by each state in the envelope, and selecting the state with the highest required cooling effect as a design point; the method comprises the following steps of carrying out grid division on the outlet section of a combustion chamber in the axial direction and the radial direction, and counting the outlet temperature distribution of the combustion chamber at a design point to obtain a distribution interval, a probability function and a mean value; according to the temperature distribution of each radius outlet of the combustion chamber, the design temperature of the inlet of the high-pressure turbine under the corresponding radius is given; designing the appearance of the blade, calculating aerodynamic and heat exchange data of a gas side, dividing the wall surface of the back side of a three-dimensional blade body basin of the blade into a plurality of areas along the radial direction and chord direction, and calculating the heat exchange coefficient required by the cold air side of each area of the blade; respectively designing cooling structures of all regions according to heat exchange coefficients required by a cold air side, selecting cooling characteristics corresponding to all regions, and completing calculation of the cold air consumption of the blades and a wall surface temperature field; judging whether the amount of cold air of the blade and the wall surface temperature meet the design requirements at the design point, and if so, establishing a response surface; if not, the design of the blade cooling structure is carried out again; according to a design point combustion chamber outlet temperature distribution probability function, carrying out quantitative analysis on the uncertainty of the temperature of the wall surface of the blade on the basis of a response surface, carrying out analysis on the mean value, the variance and the failure rate of a target value, judging whether the failure rate meets the design requirement, and if so, executing the next step; if not, the design of the blade cooling structure is carried out again, and the calculation is repeated; and (4) repeatedly optimizing the blade structure, and seeking the optimal solution of the highest wall temperature, the temperature difference and the cold air consumption of the blade to complete the design.
Preferably, the combustion chamber outlet cross section is gridded, and the specific method for counting the combustion chamber outlet temperature distribution at the design point is as follows: dividing the outlet section of the combustion chamber into grid points of n1 Xm 1 along the radial direction and the circumferential direction respectively; the gas temperature of each node is represented by Tij, i represents the radial direction, and j represents the circumferential direction; counting the temperature distribution of the nodes from 1 to j in the circumferential direction with the same radius; judging the distribution characteristics of the materials, and establishing a distribution curve; and obtaining a distribution interval, a probability function and a mean value.
Preferably, the calculation method of the gas side aerodynamic and heat exchange data comprises the following steps: and (3) calculating aerodynamic parameters of flow surfaces S2 and S1 according to the radial distribution of the designed temperature of the inlet of the high-pressure turbine, and calculating the heat exchange outside the blade by combining the profile parameters of the blade to obtain the heat exchange coefficient outside the blade and the temperature of the heat insulation wall.
Preferably, the method for calculating the heat exchange coefficient required by the cold air side comprises the following steps:
setting the wall thickness delta, the heat conductivity coefficient lambda and the outer wall temperature T of the blade w1 Inner wall temperature T w2 Temperature T of gas side g Heat transfer coefficient h of gas side g Temperature T of cold gas side c Heat transfer coefficient h of cold gas side c Outer wall area A g Inner wall area A c And a heat conduction area A e To obtain
h c ＝((1+A′)λh g (T g -T w1 ))/(λ(A′+A′ 2 )(T w1 -T c )-2A′δh g (T g -T w1 ))。
Preferably, the design requirements of the amount of cold air and the wall temperature are as follows: blade air usage < defined target; the highest outer wall temperature of the blade plus the margin < the initial melting temperature of the material; blade wall mean temperature < material allowable temperature.
Preferably, the failure rate criterion is: p ═ np/N < 3%: n is the total simulation times, and np is the times that the wall temperature of the blade does not meet the requirements.
Preferably, the calculation method of the blade temperature difference is as follows:
preferably, the specific division method of the blade area is as follows: according to the shape and the size of the blade profile and the combination of heat exchange coefficient distribution, the blade is divided into n2 Xm 2 small blocks along the radial direction and chord direction of the blade, wherein n2 is 3-7, and m2 is 4-10.
Preferably, the method for calculating the cooling effect η is as follows:
wherein, T c The inlet temperature of the cold air is set for each state; t is 4i The average total temperature of the gas at the outlet of the combustion chamber in each state; t is w Is the long-term allowable temperature of the metal material.
Preferably, each radius selects a circumferential temperature value with a standard deviation of ± 2 σ as the high pressure turbine inlet design temperature.
The application relates to a method for designing a high-pressure turbine cooling blade considering the outlet temperature of a combustion chamber, which is characterized in that when the high-pressure turbine cooling blade is designed, a state point with the highest cooling effect is selected as a design point to provide enough margin for subsequent design, then the outlet cross section of the combustion chamber is subjected to grid division, the outlet temperature distribution of the combustion chamber at the design point is counted, the inlet design temperature of the high-pressure turbine at different radiuses is given, the shape design, gas side pneumatic and heat exchange calculation are carried out according to the temperature, then the high-pressure turbine cooling blade is subjected to region division, the heat exchange coefficients required by the cold air sides at different regions are calculated, and thus the corresponding relation between the inner cavity cooling requirement of each small region of the cooling blade and the outlet temperature of the combustion chamber is obtained, the targeted design can be carried out according to the heat exchange coefficients at the cold air sides at different regions, and the cooling requirements at all the positions in the circumferential direction and the radial direction can be adapted, and judging the cold air consumption and the wall surface temperature of each area, carrying out uncertainty quantitative analysis and failure rate calculation and judgment according to a combustion chamber outlet temperature distribution probability function, and finally optimizing the blade structure until an optimal scheme is found to finish the design. The method can reasonably utilize the cold air of the blade and meet the cooling requirement of the blade.
Drawings
In order to more clearly illustrate the technical solutions provided by the present application, the following briefly introduces the accompanying drawings. It is to be expressly understood that the drawings described below are only illustrative of some embodiments of the invention.
FIG. 1 is a schematic overall flow diagram of the present application;
FIG. 2 is a schematic view of a thermal conductivity model of the present application without thermal barrier coating and without pore wall;
FIG. 3 is a schematic view of the distribution of the heat transfer coefficient of the vane gas to the outside;
FIG. 4 is a schematic diagram of the vane gas relative static pressure distribution of the present application.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application.
A design method of a high-pressure turbine cooling blade considering the temperature of an outlet of a combustion chamber is characterized in that the temperature distribution of different radiuses of the outlet of the combustion chamber is counted, the design temperature of an inlet of the high-pressure turbine with the corresponding radius is set, the heat exchange coefficient required by cold gas measurement of each partition of the blade is calculated, a targeted cooling structure design is carried out, and a corresponding temperature field evaluation method and a failure rate judgment method are provided to ensure that the blade meets the cooling requirement. The method can design a more reasonable cooling structure and save the cooling amount.
As shown in fig. 1, the method comprises the following steps:
step S100, calculating the cooling effect eta required by each state in the envelope, and selecting the state with the highest required cooling effect as a design point;
the calculation method of the cooling effect eta comprises the following steps:
wherein, T c The inlet temperature of the cold air is set for each state; t is 4i The average total temperature of the gas at the outlet of the combustion chamber in each state; t is w Is the long-term allowable temperature of the metal material.
Step S200, carrying out grid division on the outlet section of the combustion chamber according to the axial direction and the radial direction, and counting the outlet temperature distribution of the combustion chamber at a design point to obtain a distribution interval, a probability function and a mean value;
the full ring or sector combustor exit temperature is obtained through testing (combustor or engine testing) or calculation.
The measuring of outlet section temperature under the combustion chamber state wherein can realize with the rotatory displacement mechanism of special design, this mechanism installs more than 3 thermocouple rakes respectively on rotatable carousel, and every rake radially distributes the several measuring point, realizes axial direction's measurement through rotatory, can effectively measure combustion chamber outlet section temperature through this mode.
The method is characterized in that the measured data of the rotary displacement mechanism is collated to obtain the probability distribution rule and mathematical description of the outlet temperature field of the combustion chamber component, and the specific method comprises the following steps: the outlet section of the combustion chamber is divided into grid points of n1 multiplied by m1 along the radial direction and the circumferential direction respectively. The gas temperature at each grid point is characterized by T ij I represents a radial direction; j denotes the circumferential direction. And counting the temperature distribution of the nodes from 1 to j in the circumferential direction with the same radius. Judging the distribution characteristics of the materials, and establishing a distribution curve. And obtaining a distribution interval, a probability function and a mean value.
The temperature distribution of the outlets of the combustion chambers with the same radius is counted, so that a foundation is provided for the radial design of the blades; and the probability function is established to provide input for subsequent uncertain quantitative analysis of the temperature field.
Step S300, according to the temperature distribution of each radius outlet of the combustion chamber, giving the design temperature of the inlet of the high-pressure turbine under the corresponding radius;
the circumferential temperature value with the standard deviation of +/-2 sigma is selected as the design temperature of the inlet of the high-pressure turbine for each radius, namely the outlet temperature of a combustion chamber on 95.44% of the radius is not higher than the value, so that the cooling requirements of all positions in the circumferential direction can be basically met after the cooling structure is designed.
S400, designing the appearance of the blade, calculating aerodynamic and heat exchange data of a gas side, dividing the back wall surface of a three-dimensional blade body basin of the blade into a plurality of areas along the radial direction and chord direction, and calculating a heat exchange coefficient required by the cold air side of each area of the blade;
the method for calculating the heat exchange coefficient comprises the following steps: and (3) calculating aerodynamic parameters of flow surfaces S2 and S1 according to the radial distribution of the designed temperature of the inlet of the high-pressure turbine, and calculating the heat exchange outside the blade by combining the profile parameters of the blade to obtain the heat exchange coefficient outside the blade and the temperature of the heat insulation wall. The calculation of the external heat exchange refers to calculation by an integral method, a differential method or a criterion formula method. The calculation can also be solved by adopting a three-dimensional gas-heat coupling method.
According to the shape and the size of the blade profile and by combining the distribution of heat exchange coefficients, the back side wall surface of the three-dimensional blade body basin of the blade is divided into n2 Xm 2 small blocks in the radial direction and the chord direction, and n 2: 3-7, m 2: 4-10.
Then, calculating the heat exchange coefficient of the cold air side of each small block according to the calculation method, taking the design of no insulating coating and no air film hole as an example:
as shown in FIG. 2, the wall thickness and the thermal conductivity are represented as δ and λ, and the outer wall and inner wall temperatures are represented as T w1 And T w2 The temperature and heat transfer coefficient of the gas side is T g And h g The temperature and heat transfer coefficient of the cold gas side is T c And h c The outer wall, the inner wall and the heat conducting area are respectively A g 、A c 、A e . And preliminarily setting the wall thickness of the blade (reference: 1.3mm of guide vane and 0.8mm of rotor) according to the design and processing level of the existing blade. Will T w1 Set to the long-term service temperature of the blade material. T is c The value gives a temperature increase of 0-80K on the basis of the cold air inlet temperature in the flow direction. Then h can be solved by c The relation with A'.
Setting: a. the e ＝(A g +A c )/2…………………………(8)
A′＝A c /A g …………………………………(9)
h c ＝((1+A′)λh g (T g -T w1 ))/(λ(A′+A′ 2 )(T w1 -T c )-2A′δh g (T g -T w1 ))……(10)
The required heat exchange coefficient of the cold air side of all the areas can be calculated by adopting the formula (10), the value represents the cooling requirement of the blade aiming at the circumferential and radial temperature distribution of the outlet of the combustion chamber, and therefore, the data can be used for carrying out targeted cooling design on each area.
Step S500, respectively designing cooling structures of all regions according to heat exchange coefficients required by the cold air side, selecting cooling characteristics corresponding to all regions, and completing calculation of the cold air consumption of the blades and a wall surface temperature field;
for different heat exchange coefficients, the design of the cooling structure is also different, such as the diameter, the row number and the position of the air film holes, the number and the diameter of the cooling holes on the cold air guide pipe, the shape and the position of the inner cavity rib and the like can be changed along with the difference of the heat exchange coefficients of the cold air side.
Step S600, judging whether the amount of cold air of the blade and the wall surface temperature meet the design requirements at the design point, and if so, establishing a response surface; if not, repeating the step S500, and designing the cooling structure of the blade again until the requirement is met;
the design requirements of the cold air consumption and the wall surface temperature are as follows:
blade air usage < defined target;
the temperature of the highest outer wall of the blade + the margin < the initial melting temperature of the material, wherein the margin is generally selected from 100 and 180K;
blade wall mean temperature < material allowable temperature.
Step S700, according to a design point combustion chamber outlet temperature distribution probability function, carrying out quantitative analysis on uncertainty of temperature of a blade wall surface on the basis of a response surface, carrying out analysis on a mean value, a variance and a failure rate of a target value, judging whether the failure rate meets design requirements or not, and if so, executing the next step; if not, the design of the blade cooling structure is carried out again, and the calculation is repeated;
the failure rate judgment standard is as follows: p ═ np/N < 3%: n is the total simulation times, and np is the times that the wall temperature of the blade does not meet the requirements.
According to the method, the uncertainty quantitative analysis of the wall surface temperature is carried out according to the distribution probability function of the temperature of the outlet of the combustion chamber, and the failure rate is judged, so that the cooling structure design of the blade can be further ensured to meet the circumferential and radial distribution of the outlet of the combustion chamber.
And step S800, repeatedly optimizing the blade structure, and seeking the optimal solution of the highest wall temperature, the temperature difference and the cold air consumption of the blade to complete the design.
1) Min leaf blade maximum wall temperature.
2) Min temperature difference.
3) Min dosage of cold air.
Structural optimization refers to varying the size and number of selected cooling features or modifying the cooling features for each location.
When designing the cooling blade of the high-pressure turbine, selecting a state point with the highest cooling effect as a design point to provide enough margin for subsequent design, then carrying out grid division on the outlet section of the combustion chamber, counting the temperature distribution of the outlet of the combustion chamber at the design point, giving out the inlet design temperature of the high-pressure turbine at different radiuses, carrying out appearance design, gas side pneumatic and heat exchange calculation according to the temperature, then carrying out region division on the cooling blade of the high-pressure turbine, calculating the heat exchange coefficients required by the cold air sides of different regions, thus obtaining the corresponding relation between the inner cavity cooling requirement of each small region of the cooling blade and the outlet temperature of the combustion chamber, carrying out targeted design according to the heat exchange coefficients of the cold air sides at different regions, adapting to the cooling requirements at circumferential and radial positions, and then carrying out judgment on the cold air consumption and wall surface temperature of each region, and carrying out uncertainty quantitative analysis and failure rate calculation and judgment according to the outlet temperature distribution probability function of the combustion chamber, and finally optimizing the blade structure until an optimal scheme is found to finish the design. The method can reasonably utilize the cold air of the blade and meet the cooling requirement of the blade.
As a specific embodiment, the design of the guide vane of the high-pressure turbine taking into account the temperature non-uniformity at the outlet of the combustor is performed according to the method described above, which is described in detail below.
1) The cooling effect required for each state is calculated from equation (6), and as shown in table 1, state 2 is selected as the blade design point.
TABLE 1 required cooling effect for each state
| Status of state | Required cooling effect | Status of state | Required cooling effect |
| State 1 | 0.25 | State 8 | 0.21 |
| State 2 | 0.51 | State 9 | 0.25 |
| State 3 | 0.45 | State 10 | 0.18 |
| State 4 | 0.47 | State 11 | 0.36 |
| State 5 | 0.40 | State 12 | 0.30 |
| State 6 | 0.35 | State 13 | 0.48 |
| State 7 | 0.28 | State 14 | 0.12 |
2) And (4) counting the temperature distribution of the outlet of the combustion chamber at the design point to obtain a distribution interval, a probability function and an average value. The combustion chamber outlet cross section (sector) is divided into grid points of 5 × 50 in the radial direction and the circumferential direction. And (4) obtaining the probability distribution rule and the mathematical description of the temperature field of the outlet of the combustion chamber component. As shown in table 2:
TABLE 2 combustor exit temperature field distribution
| Probability distribution | Distribution interval, K | Mean value, K | Standard deviation of | |
| Radius 1 | Normal distribution | 1553-1985 | 1820 | 10 |
| Radius 2 | Normal distribution | 1665-2060 | 1900 | 10.1 |
| Radius 3 | Normal distribution | 1711-2105 | 1950 | 10.2 |
| Radius 4 | Normal distribution | 1700-2093 | 1925 | 10 |
| Radius 5 | Normal distribution | 1608-2030 | 1860 | 9.8 |
3) And (4) giving the design temperatures of the inlets of the guide vanes of the high-pressure turbine with different radiuses. As shown in table 3:
TABLE 3 high pressure turbine Inlet design temperature
| Design inlet temperature，K | |
| Radius 1 | 1856 |
| Radius 2 | 1935 |
| Radius 3 | 1984 |
| Radius 4 | 1960 |
| Radius 5 | 1915 |
4) And calculating the pneumatic and heat exchange data of the gas side to obtain the heat exchange coefficients required by the cold air sides of all the parts of the blade.
The aerodynamic parameter calculation of the flow surface of S2 and S1 is carried out according to the radial distribution of the designed temperature of the inlet of the guide vane of the high-pressure turbine, the external heat exchange calculation of the vane is carried out by adopting an integral method in combination with the profile parameter of the vane, and the curve corresponding to the section with the radius of 3, which is obtained by the external heat exchange coefficient of the vane and the temperature of the heat insulation wall, is shown in figures 3-4.
According to the shape and the size of the blade profile and by combining the distribution of heat exchange coefficients, the back side wall surface of the three-dimensional basin of the blade is divided into 3 multiplied by 8 small blocks along the radial direction and the chord direction.
The wall thickness of the blade is preliminarily given to be 1.3 mm. Will T w1 Set to the long-term allowable temperature 1373K of the blade material. Then, the formula (10) can be used to solve each small block h when no thermal insulation coating and no air film hole are obtained c The relationship with A' is shown in Table 4 after the simplification of the middle section. T is c The value gives a temperature increase of 0-80K on the basis of the cold air inlet temperature in the flow direction.
Table 4 Cold gas side Heat transfer coefficients for Cross sections
| Chordal block sequence | Relation, W/(m) 2 ·K) |
| 1 | h c ＝116.6×10 3 /9.38A′ |
| 2 | h c ＝28.27×10 3 /9.38A′ |
| 3 | h c ＝49.13×10 3 /9.38A′ |
| 4 | h c ＝55.95×10 3 /9.38A′ |
| 5 | h c ＝96.74×10 3 /9.38A′ |
| 6 | h c ＝86.98×10 3 /8.93A′ |
| 7 | h c ＝96.11×10 3 /8.48A′ |
| 8 | h c ＝99.38×10 3 /8.25A′ |
5) And designing a cooling structure of each section according to the heat exchange coefficient of the cold air side, selecting cooling characteristics and finishing the calculation of the cold air consumption of the blades and the wall surface temperature field. In order to further reduce the temperature of the hot spot of the blade, a film hole is added on the surface of the blade.
6) And judging whether the amount of the blade cold air and the wall surface temperature meet the following requirements at the design point. And if not, repeating the step 5 until the requirement is met, and establishing a Kriging response surface. Wherein the margin is selected to be 150K.
7) And carrying out quantitative analysis on the uncertainty of the temperature of the wall surface of the blade on the basis of the response surface according to the probability function of the temperature distribution of the outlet of the combustion chamber at the design point.
8) And optimizing the blade structure by adopting an SMODE algorithm, repeating 5-8, and seeking the following target optimal solution. As shown in tables 5-7.
TABLE 5 Final Structure of the blade
TABLE 6 leaf relative air consumption cold air usage
| Serial number | Relative gas consumption% |
| 1 | 2.05 |
| 2 | 0.73 |
| 3 | 2.03 |
| 4 | 3.34 |
| 5 | 2.96 |
| 6 | 5.22 |
| 7 | 4.64 |
| 8 | 4.64 |
| General assembly | 11.11 |
TABLE 7 temperature of the outer wall of the middle section
| Serial number | 1 | 2 | 3 | 4 |
| Wall temperature, K | 1255-1300 | 1170-1210 | 1170-1250 | 1210-1260 |
| Serial number | 5 | 6 | 7 | 8 |
| Wall temperature, K | 1210-1260 | 1210-1255 | 1255-1280 | 1280-1310 |
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (10)
1. A method of designing a high pressure turbine cooling blade that takes into account combustor exit temperature, comprising:
calculating the cooling effect eta required by each state in the envelope, and selecting the state with the highest required cooling effect as a design point;
the method comprises the following steps of (1) carrying out grid division on the outlet section of a combustion chamber in the axial direction and the radial direction, and counting the temperature distribution of the outlet of the combustion chamber at a design point to obtain a distribution interval, a probability function and a mean value;
according to the temperature distribution of each radius outlet of the combustion chamber, the design temperature of the inlet of the high-pressure turbine under the corresponding radius is given;
designing the appearance of the blade, calculating aerodynamic and heat exchange data of a gas side, dividing the back wall surface of a three-dimensional blade body basin of the blade into a plurality of areas along the radial direction and chord direction, and calculating the heat exchange coefficient required by the cold air side of each area of the blade;
respectively designing cooling structures of all regions according to heat exchange coefficients required by a cold air side, selecting cooling characteristics corresponding to all regions, and completing calculation of the cold air consumption of the blades and a wall surface temperature field;
judging whether the amount of cold air of the blade and the wall surface temperature meet the design requirements at the design point, and if so, establishing a response surface; if not, the design of the blade cooling structure is carried out again;
according to a design point combustion chamber outlet temperature distribution probability function, carrying out quantitative analysis on the uncertainty of the temperature of the wall surface of the blade on the basis of a response surface, carrying out analysis on the mean value, the variance and the failure rate of a target value, judging whether the failure rate meets the design requirement, and if so, executing the next step; if not, the design of the blade cooling structure is carried out again, and the calculation is repeated;
and (4) repeatedly optimizing the blade structure, and seeking the optimal solution of the highest wall temperature, the temperature difference and the cold air consumption of the blade to complete the design.
2. The method for designing a cooling blade of a high pressure turbine in consideration of the outlet temperature of a combustor as claimed in claim 1, wherein the outlet cross section of the combustor is gridded, and the specific method for counting the outlet temperature distribution of the combustor at the design point is as follows: dividing the outlet section of the combustion chamber into grid points of n1 Xm 1 along the radial direction and the circumferential direction respectively; the gas temperature of each node is represented by Tij, i represents the radial direction, and j represents the circumferential direction; counting the temperature distribution of the nodes from 1 to j in the circumferential direction with the same radius; judging the distribution characteristics of the materials, and establishing a distribution curve; and obtaining a distribution interval, a probability function and a mean value.
3. The method of claim 1, wherein the gas side aerodynamics and heat transfer data are calculated by: and (3) calculating aerodynamic parameters of flow surfaces S2 and S1 according to the radial distribution of the designed temperature of the inlet of the high-pressure turbine, and calculating the heat exchange outside the blade by combining the profile parameters of the blade to obtain the heat exchange coefficient outside the blade and the temperature of the heat insulation wall.
4. The method for designing a cooling blade of a high pressure turbine in consideration of the outlet temperature of a combustor according to claim 1, wherein the method for calculating the heat exchange coefficient required on the cold gas side comprises:
setting the wall thickness delta, the heat conductivity coefficient lambda and the outer wall temperature T of the blade w1 Inner wall temperature T w2 Temperature T of gas side g Heat transfer coefficient h of gas side g Temperature T of cold gas side c Heat transfer coefficient h of cold gas side c Outer wall area A g Inner wall area A c Heat conducting area A e To obtain
h c ＝((1+A′)λh g (T g -T w1 ))/(λ(A′+A′ 2 )(T w1 -T c )-2A′δh g (T g -T w1 ))。
5. The method of claim 1, wherein the design requirements for cooling air usage and wall temperature are as follows:
blade air usage < defined target;
the highest outer wall temperature of the blade plus the margin < the initial melting temperature of the material;
blade wall mean temperature < material allowable temperature.
6. The method of designing a high pressure turbine cooling blade taking into account the combustor exit temperature as recited in claim 1, wherein the failure rate criteria are: p ═ np/N < 3%: n is the total simulation times, and np is the times that the wall temperature of the blade does not meet the requirements.
7. A method of designing a high pressure turbine cooling vane that takes into account the combustor exit temperature as set forth in claim 1, wherein the vane temperature differential is calculated by:
8. the method for designing a cooling blade of a high pressure turbine in consideration of the outlet temperature of a combustor as claimed in claim 1, wherein the specific division method of the blade region is as follows: according to the shape and the size of the blade profile and the combination of heat exchange coefficient distribution, the blade is divided into n2 Xm 2 small blocks along the radial direction and chord direction of the blade, wherein n2 is 3-7, and m2 is 4-10.
9. A method of designing a cooling blade for a high pressure turbine taking into account the temperature at the outlet of the combustor as set forth in claim 1, wherein the cooling effect η is calculated by:
wherein, T c The inlet temperature of the cold air is set for each state; t is 4i The average total temperature of the gas at the outlet of the combustion chamber in each state; t is w Is the long-term allowable temperature of the metal material.
10. A method of designing a high pressure turbine cooling blade that takes into account combustor exit temperature as recited in claim 1, wherein: and selecting a circumferential temperature value with standard deviation of +/-2 sigma as the design temperature of the high-pressure turbine inlet for each radius.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210425143.1A CN114840921B (en) | 2022-04-21 | 2022-04-21 | Design method of high-pressure turbine cooling blade at outlet of combustion chamber |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210425143.1A CN114840921B (en) | 2022-04-21 | 2022-04-21 | Design method of high-pressure turbine cooling blade at outlet of combustion chamber |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN114840921A true CN114840921A (en) | 2022-08-02 |
| CN114840921B CN114840921B (en) | 2023-10-20 |
Family
ID=82565509
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210425143.1A Active CN114840921B (en) | 2022-04-21 | 2022-04-21 | Design method of high-pressure turbine cooling blade at outlet of combustion chamber |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114840921B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116502469A (en) * | 2023-06-25 | 2023-07-28 | 中国航发四川燃气涡轮研究院 | Turbine blade temperature correction method and device based on temperature test data |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20170101871A1 (en) * | 2015-10-08 | 2017-04-13 | Rolls-Royce Deutschland Ltd & Co Kg | Blade for fluid flow machine, turbo fan engine and method for manufacturing a blade |
| CN110489863A (en) * | 2019-08-20 | 2019-11-22 | 成立航空技术有限公司 | The determination method of aero-engine primary combustor chamber exit temperature field index |
| CN112507586A (en) * | 2020-12-02 | 2021-03-16 | 中国航发沈阳发动机研究所 | Rapid assessment method for two-dimensional temperature and strength of turbine air cooling blade |
-
2022
- 2022-04-21 CN CN202210425143.1A patent/CN114840921B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20170101871A1 (en) * | 2015-10-08 | 2017-04-13 | Rolls-Royce Deutschland Ltd & Co Kg | Blade for fluid flow machine, turbo fan engine and method for manufacturing a blade |
| CN110489863A (en) * | 2019-08-20 | 2019-11-22 | 成立航空技术有限公司 | The determination method of aero-engine primary combustor chamber exit temperature field index |
| CN112507586A (en) * | 2020-12-02 | 2021-03-16 | 中国航发沈阳发动机研究所 | Rapid assessment method for two-dimensional temperature and strength of turbine air cooling blade |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116502469A (en) * | 2023-06-25 | 2023-07-28 | 中国航发四川燃气涡轮研究院 | Turbine blade temperature correction method and device based on temperature test data |
| CN116502469B (en) * | 2023-06-25 | 2023-09-05 | 中国航发四川燃气涡轮研究院 | Turbine blade temperature correction method and device based on temperature test data |
Also Published As
| Publication number | Publication date |
|---|---|
| CN114840921B (en) | 2023-10-20 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN114840921A (en) | Method for designing cooling blade of high-pressure turbine at outlet of combustion chamber | |
| US7670112B2 (en) | Turbine blade with cooling and with improved service life | |
| RU2616743C2 (en) | Gas turbine engine | |
| CN115688327A (en) | Turbine blade temperature field calculation method with subareas and variable-thickness thermal barrier coating | |
| TW201920829A (en) | Turbine blade and gas turbine | |
| Chowdhury et al. | A predictive model for preliminary gas turbine blade cooling analysis | |
| CN113357054B (en) | High-temperature gas channel design method | |
| CN113167123A (en) | Turbine blade with coupled serpentine channels | |
| CN113343494B (en) | Aircraft engine bearing cavity thermal performance simulation calculation model correction method | |
| CN109948231B (en) | Method and device for analyzing thermal cycle parameters of engine | |
| Vinogradov et al. | Robust optimization of the HPT blade cooling and aerodynamic efficiency | |
| CN111428184B (en) | Method for calculating core size of plate-fin heat exchanger | |
| CN113874600B (en) | Turbine blade with serpentine passage | |
| JP7118597B2 (en) | Method for manufacturing internal ribs | |
| CN114282323B (en) | Flow distribution prediction method for turbine blade laminate cooling structure | |
| CN115859739B (en) | Method for judging local spalling damage degree of thermal barrier coating of combustion engine and aeroengine | |
| RU2680636C1 (en) | Cooling system of a multi-circuit gas turbine unit | |
| CN110059372A (en) | A kind of objective design method of the shell-and-tube heat exchanger based on differential evolution algorithm | |
| CN110427661A (en) | A kind of high efficient heat exchanging structure adaptive optimum design method based on variational method | |
| CN111783306B (en) | Turbine blade cold air quantity and cold effect characteristic analysis method | |
| Pis’Mennyi et al. | Optimization of the ribbing of a new heat exchange surface of flat-oval tubes | |
| Sivakumar et al. | Influence of Different Rib Height on Heat Transfer Augmentation in Rectangular Convergent/Divergent Channels With Continuous Ribs on Bottom Surface | |
| Cunha et al. | Analysis of airfoil trailing edge heat transfer and its significance in thermal-mechanical design and durability | |
| CN113188585A (en) | Freezing station sensor fault diagnosis method based on few redundant sensors | |
| CN114741817B (en) | Turbine cooling blade design method with thermal barrier coating |
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 |