CN115203862B - Air film hole design method based on volume density distribution - Google Patents
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- G06F30/17—Mechanical parametric or variational design
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
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- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
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- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
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Abstract
The invention relates to the technical field of turbine blade cooling, and discloses a method for designing air film holes based on volume density distribution, which comprises the steps of calculating the temperature field and the stress field of a single turbine blade without the air film holes, and determining the number of the air film holes according to the hole type of the air film holes and cold air parameters introduced into the single turbine blade; distribution is based on temperature field, stress field division of turbine bladesnThe volume domain is subjected to volume domain gas film hole distribution design based on balanced temperature and volume domain gas film hole distribution design based on balanced stress; the problem of large radial temperature distribution difference can be fully avoided by the aid of the film hole distribution based on the balanced temperature, and the problem that a plurality of film holes are gathered in a high-stress area can be avoided by the aid of the film hole distribution based on stress, so that strength of the turbine blade is improved; the turbine blade air film hole comprehensive distribution scheme which gives consideration to cooling and strength is obtained by comprehensively considering the temperature and stress distribution characteristics of the turbine blade and balancing the requirements of temperature uniformity and low stress level through weight distribution, and the practical engineering application is met.
Description
Technical Field
The invention relates to the technical field of turbine blade cooling, and discloses a volume density distribution-based film hole design method.
Background
At present, the turbine blade of the advanced aero-engine adopts air film cooling to improve the temperature resistance, after cooling air in the inner cavity of the turbine blade passes through an air film hole, the cooling air is attached to the outer surface of the blade body to protect a turbine blade substrate from being ablated by high-temperature and high-pressure fuel gas, and the design of the air film hole is very critical to the temperature resistance of the turbine blade.
A large amount of numerical values and experimental researches are carried out on the structural form and the effect of air film cooling at home and abroad, but research objects mainly focus on a single air film hole, and the research on the distribution of the air film hole on the surface of the turbine blade is very little. Compared with the turbine blade, the diameter of the film hole is small (generally less than 1 mm), the cooling range is limited, dozens or even hundreds of film holes are often needed for a single turbine blade, and therefore, the distribution of the film holes is very important for the temperature uniformity and the stress level of the whole turbine blade.
At present, the air film hole distribution design method at home and abroad basically adopts a row distribution method (as shown in figure 1), namely, a row of air film holes are distributed along the blade profile (circumferential direction) according to the temperature field of the turbine blade. This method can improve the temperature unevenness in the circumferential direction of the blade profile, but cannot sufficiently consider the temperature unevenness in the radial direction of the turbine blade and the strength level of the entire turbine blade. The temperature of the blade body is uneven in the radial direction and the circumferential direction, and the air film holes designed in a row distribution mode can only solve the problem that the temperature is uneven along the circumferential direction of the blade profile, and the problem that the radial temperature distribution is uneven cannot be fully considered. Meanwhile, the distribution of an air film hole at the same position of the turbine blade can cause the loss of a large amount of matrix materials, for example, a row of 20 air film holes with the diameter of 0.4mm is arranged in the turbine blade with the blade height of 30mm, the total length of the air film hole is 8mm, and the total length of the air film hole accounts for about 27% of the whole turbine blade height, which means that the bearing capacity of the position is greatly reduced.
In conclusion, the distribution design research of the air film holes on the surface of the turbine blade is very little at home and abroad; meanwhile, the conventional row distribution design method cannot fully consider the integral temperature uniformity and the strength level of the turbine blade, and a film hole distribution design method which fully considers the radial temperature unevenness, the circumferential temperature unevenness and the stress unevenness of the turbine blade is lacked, so that the purposes of improving the temperature uniformity of the turbine blade and improving the strength of the turbine blade are achieved.
Disclosure of Invention
The invention aims to provide a volume density distribution-based film hole design method, which can comprehensively consider the temperature and stress distribution characteristics of a turbine blade, balance the requirements of temperature uniformity and low stress level through weight distribution, obtain a turbine blade film hole comprehensive distribution scheme which has both cooling and strength, and meet the practical engineering application.
In order to realize the technical effects, the invention adopts the technical scheme that:
a method for designing a gas film hole based on volume density distribution comprises the following steps:
step 1, respectively calculating a temperature field and a stress field of a single turbine blade without an air film hole according to the working condition requirement of the turbine blade of the aero-engine;
step 2, designing the hole pattern of the air film holes according to the temperature field, introducing the flow of the cold air inside the single turbine blade and the flow of the single hole, and determining the number of the air film holesn；
Step 3, dividing the temperature field and the stress field of the single turbine blade intonThe volume domain is respectively subjected to volume domain film hole distribution design based on balanced temperature and volume domain film hole distribution design based on balanced stress aiming at a single turbine blade;
and 4, respectively giving different weights to the volume domain based on the equilibrium temperature and the volume domain based on the equilibrium stress according to the working condition requirement of the turbine blade and the distribution characteristics of the temperature field and the stress field of different parts, balancing the nonuniformity of the temperature and the stress, and obtaining a turbine blade air film hole distribution scheme considering both the temperature field and the stress field.
Further, in the step 1, heat transfer calculation of the turbine blade without the film holes is carried out according to the temperature, pressure and flow rate of the gas at the inlet and the outlet of the turbine blade, the temperature, pressure, flow rate and flow rate of the cold air, so as to obtain the temperature field of the turbine blade without the film holes.
Further, in the step 1, according to the temperature field, the rotating speed, the pressure and the load of the turbine blade without the air film holes, the strength of the turbine blade without the air film holes is calculated, and the stress field of the turbine blade without the air film holes is obtained.
Further, in step 2, the number of film holes = single turbine blade cold air flow/single hole flow.
Further, the design flow of the volume domain gas film hole distribution based on the equilibrium temperature in the step 3 is as follows:
dividing the turbine blades intonIndividual volume domainV k Each volume domainV k The total temperature in the interior isT m ；
Will be provided withnThe air film holes are distributed in each volume areaV k In the inner part, the volume area of a single air film hole is realizedV k The total temperature in the reactor is the same;
wherein the content of the first and second substances,
，T m is the average temperature of the individual film hole coverage,T t the total temperature of the entire turbine blade body,T i is a unit volumedvThe temperature of (2).
Further, the volume domain gas film hole distribution design flow based on the equilibrium stress in the step 3 is as follows:
by dividing the turbine blades intonIndividual volume domainV s Each volume domainV s Total internal stress isσ m ；
Will be provided withnThe air film holes are distributed in each volume areaV s In the volume region where the single air film hole is positionedV s The total internal stress is the same;
wherein the content of the first and second substances,
，σ m is the average stress covered by a single film hole,σ t the total amount of stress for the entire turbine blade body,σ i is a unit volumedvOf the stress of (c).
Further, according to the turbine blade air film hole distribution scheme obtained in the step 4, temperature evaluation of the turbine blade with the air film holes and strength evaluation of the turbine blade with the air film holes are respectively carried out, and the turbine blade air film hole distribution scheme is determined to be qualified after the design requirements are met; otherwise, adjusting the hole type of the air film hole or the flow of the cold air introduced into the single turbine blade and the single-hole flow parameters, and repeating the step 2-4 until the temperature evaluation and the strength evaluation are met and the design requirements are met.
Further, the temperature evaluation method comprises the following steps: and performing heat transfer calculation on the turbine blade with the film hole to obtain a temperature field of the turbine blade with the film hole, wherein the highest temperature is lower than the long-term use temperature of the material, and the requirement on the temperature is judged to be met.
Further, the strength evaluation method comprises the following steps: and acquiring a stress field of the turbine blade with the film hole, and judging that the maximum stress of the turbine blade is lower than the allowable stress value in design use, wherein the maximum stress of the turbine blade is judged to meet the requirement.
Compared with the prior art, the invention has the following beneficial effects:
1. according to the invention, the problem of large radial temperature distribution difference can be fully avoided based on the air film hole distribution of the equilibrium temperature, and the problem that a plurality of air film holes are gathered in a high stress area can be avoided based on the air film hole distribution of the stress, so that the strength of the turbine blade is improved; the turbine blade air film hole comprehensive distribution scheme which gives consideration to cooling and strength is obtained by comprehensively considering the temperature and stress distribution characteristics of the turbine blade and balancing the requirements of temperature uniformity and low stress level through weight distribution, and the practical engineering application is met.
2. The invention adopts the design of the film holes with volume density distribution according to the radial and circumferential temperature distribution and stress distribution characteristics of the turbine blade, establishes a standardized design flow so as to achieve the purposes of improving the temperature uniformity of the turbine blade and improving the strength of the turbine blade, has strong comprehensiveness and great flexibility, can be used for the distribution design of the film holes of the turbine guide turbine blade and the turbine rotor turbine blade, can also be popularized and used for the design of the film holes of structures such as a flame tube, a spray pipe and the like of a combustion chamber of an aircraft engine, and has great engineering application value.
Drawings
FIG. 1 is a schematic view of a conventional turbine blade film hole distribution;
FIG. 2 is a flow chart of a method for designing a gas film hole based on a bulk density distribution in example 2;
FIG. 3 is an exemplary diagram of a high pressure turbine rotor blade obtained by the film hole design considering only the temperature unevenness in example 2.
10, turbine blades; 20. a gas film hole; 11. the middle part of the front edge; 12. a trailing edge; 13. a leading edge root; 14. a leading edge tip.
Detailed Description
The present invention will be described in further detail with reference to the following examples and the accompanying drawings. It should be understood that the scope of the above-described subject matter is not limited to the following examples, and any techniques implemented based on the disclosure of the present invention are within the scope of the present invention.
Example 1
Referring to fig. 1, a method for designing a gas film hole based on bulk density distribution includes the following steps:
step 1, respectively calculating the temperature field and the stress field of a single turbine blade 10 without an air film hole according to the working condition requirement of the turbine blade 10 of the aero-engine;
step 2, designing and introducing cold into the interior of the single turbine blade 10 according to the temperature field and the hole type of the film hole 20The flow rate of gas and the flow rate of a single hole are determined, and the number of 20 holes of a gas film hole is determinedn；
Step 3, dividing the temperature field and the stress field of the single turbine blade 10 intonThe volume domain is respectively designed for the volume domain film hole 20 distribution based on the equilibrium temperature and the volume domain film hole 20 distribution based on the equilibrium stress aiming at the single turbine blade 10;
and 4, respectively giving different weights to the volume domain based on the equilibrium temperature and the volume domain based on the equilibrium stress according to the working condition requirement of the turbine blade 10 and the distribution characteristics of the temperature field and the stress field of different parts, balancing the nonuniformity of the temperature and the stress, and obtaining the distribution scheme of the gas film holes 20 of the turbine blade 10 considering both the temperature field and the stress field.
In the embodiment, according to the radial and circumferential temperature distribution and stress distribution characteristics of the turbine blade 10, the design of the film holes 20 with volume density distribution is adopted, the blade body is divided into volume areas with different volume sizes according to the number of the film holes 20, the total temperature or the total stress of each volume area is equal, namely the temperature of the cooling area of each film hole 20 is equal, so as to realize temperature equalization, or the stress of the covering area of each film hole 20 is equal, so that the aggregation of a plurality of film holes 20 in a high stress area is avoided, so as to improve the strength of the turbine blade 10. In the embodiment, the problem of large radial temperature distribution difference can be fully avoided by the film hole 20 distribution based on the equilibrium temperature, and the film hole 20 distribution based on the stress can avoid the phenomenon that a plurality of film holes 20 are gathered in a high-stress area, so that compared with the film hole distribution of the traditional turbine blade, the film hole distribution related by the method can improve the strength of the turbine blade 10; by comprehensively considering the temperature and stress distribution characteristics of the turbine blade 10 and by weight distribution to balance the requirements of temperature uniformity and low stress level, a comprehensive distribution scheme of the turbine blade 10 film holes 20 with both cooling and strength is obtained, and the practical engineering application is met.
In the embodiment, in addition to stress concentration caused by the film holes 20 formed on the turbine blade 10, temperature stress is also generated due to temperature gradient caused by non-uniform temperature; however, for the distribution of the weight, generally, the temperature unevenness can be considered preferentially, so in the engineering design of the turbine blade 10, the volume domain distribution weight ratio based on the equilibrium temperature and the equilibrium stress is performed according to the following formula of 7 to 6.
Example 2
Referring to fig. 1-3, the present embodiment illustrates aspects of the present invention in terms of a certain type of turbine blade film hole distribution design. The method specifically comprises the following steps:
1) Calculating the temperature field of the gas film hole-free turbine blade 10
According to the temperature, pressure and flow rate of the gas at the inlet and the outlet of the turbine blade 10, and the temperature, pressure, flow rate and flow rate of the cold air, the heat transfer calculation of the turbine blade 10 without the film holes is carried out, and the temperature field (temperature distribution cloud chart) of the turbine blade 10 without the film holes is obtained.
2) Determining the number of air film holesnStep (ii) of
Designing the film hole pattern according to the temperature field of the turbine blade 10, introducing the flow rate of the cold air in the single turbine blade 10, the flow rate of the single hole and the like, combining the film hole pattern and the cooling efficiency, primarily designing the film hole parameters, and determining the number of the film holesn(ii) a In this embodiment, the number of film holes = single turbine blade cold air flow/single hole flow.
3) Calculating a volume domainV k Total internal temperatureT m
Calculating the total temperature of the entire turbine blade 10 based on the temperature field of the turbine blade 10
(ii) a Wherein the content of the first and second substances,T i is a unit volumedvThe temperature of (a) is set to be,T t the total temperature of the entire turbine blade 10; according to the total temperature of the turbine blade 10T t Number of pores of the gas filmnCalculating the average temperature of the coverage of the single air film holeT m =T t /n。
4) Obtaining bulk density distribution based on equilibrium temperature
Divide the turbine blade 10 intonIndividual volume domainV k Each volume domainV k Total internal temperature isT m (ii) a Will be provided withnThe air film holes are distributed in each volume areaV k In the inner part, the volume region where a single air film hole is positioned is realizedV k The total amount of the temperature in the chamber is the same.
5) Calculating stress field of a gas film hole-free turbine blade 10
According to the temperature field, the rotating speed and the pressure of the non-film-hole turbine blade 10 and the load of the turbine blade 10, the strength of the non-film-hole turbine blade 10 is calculated, and the stress field (stress distribution cloud chart) of the non-film-hole turbine blade 10 is obtained.
6) Calculating a volume domainV s Total amount of internal stressσ m
From the stress field of the turbine blade 10, the total stress of the entire turbine blade 10 is calculated
(ii) a Wherein the content of the first and second substances,σ i is a unit volumedvThe stress of (a) is greater than (b),σ t the total amount of stress for the entire turbine blade 10; according to the total stress of the turbine blade 10σ t Number of pores of the gas filmnCalculating the average stress of single gas film hole coverage
。
7) Obtaining bulk density distribution based on equilibrium stress
Divide the turbine blade 10 intonIndividual volume domainV s Each volume domainV s The total amount of internal stress isσ m ；
Will be provided withnThe air film holes are distributed in each volume areaV s In the volume region where the single air film hole is positionedV s The total internal stress is the same;
8) Determining a gas film pore distribution scheme
Obtaining the volume domain based on the equilibrium temperature from the step 4) and the step 7) respectivelyV k And volume domain based on equilibrium stressV s By dispensing a volumeV k AndV s different weights are used for balancing the temperature uniformity and the stress magnitude, and the turbine blade 10 air film hole distribution scheme giving consideration to both the temperature field and the stress field is obtained.
In actual engineering design, if the turbine blade 10 has a large temperature field non-uniformity and a low stress level, a volume region based on equilibrium temperature can be setV k Weight of 100%, and volume domain based on equilibrium stressV s And if the weight is 0, obtaining a gas film hole design scheme only considering temperature unevenness only through the steps 1) to 4), namely the temperature of each gas film hole cooling area is equal, so as to achieve the purpose of temperature homogenization. As shown in FIG. 1 for a conventional film hole distribution, the film holes 20 distribution on the turbine blade 10 are designed in a row, failing to adequately account for radial temperature non-uniformity and reduce the stress level of the turbine blade 10. And fig. 3 shows a high-pressure turbine working blade, the middle 11 of the front edge of the turbine blade 10 is generally a high-temperature region, the tail edge 12 of the turbine blade 10 is a low-temperature region, and the root 13 of the front edge of the turbine blade 10 and the tip 14 of the front edge of the turbine blade 10 are middle-temperature regions, by performing the step 1) to 4) of the embodiment, the design of the gas film holes only considering the temperature unevenness is performed, the dense gas film holes can be distributed in the middle 11 of the front edge of the turbine blade 10 in the high-temperature region, the volume covered by a single gas film hole is small, while the temperature of the tail edge 12 of the turbine blade 10 is low, the gas film holes are distributed sparsely, the volume covered by a single gas film hole is large, the root 13 of the front edge of the turbine blade 10 in the middle-temperature region and the tip 14 of the front edge of the turbine blade 10 are between, the dense high-temperature gas film holes are distributed sparsely, and the gas film holes are distributed in the low-temperature region, and the whole turbine blade 10 presents a self-coordination phenomenon of the temperature distribution and the gas film hole distribution.
If the turbine blade 10 has a relatively high stress level and a relatively uniform temperature, a volume region based on balanced stress may be providedV s Weight 100% and volume based on equilibrium temperatureV k If the weight is 0, the steps 3) to 4) can be skipped, and the gas film hole design scheme considering the stress distribution, namely the stress phase of the coverage area of each gas film hole, can be obtained only through the steps 5) to 7)And the like, to avoid the accumulation of multiple film holes in high stress regions to increase the strength of the turbine blade 10.
For other types of turbine blades 10, the degree of non-uniformity of the temperature field and the stress level of the turbine blade 10 have an effect, and then the turbine blade 10 is properly adjusted according to the actual degree of non-uniformity of the temperature field and the stress distribution condition in combination with the requirements of the working conditionsV k AndV s and (4) weight distribution. Generally, temperature unevenness can be considered preferentially, so in the engineering design of the turbine blade 10, the volume distribution weight ratio based on the equilibrium temperature and the equilibrium stress is performed according to the following steps of 7.
9) Temperature assessment of a turbine blade 10 with film holes
In the embodiment, the temperature field of the turbine blade 10 with the film hole is obtained through heat transfer calculation of the turbine blade 10 with the film hole, and if the highest temperature is lower than the long-term use temperature of the material, the requirement on the temperature is judged to be met, and then strength evaluation is carried out; if the highest temperature is higher than or equal to the long-term use temperature of the material, adjusting the hole type of the air film hole or the flow of the cold air introduced into the single turbine blade 10 and the single-hole flow parameters, and repeating the steps 2) to 8) until the temperature requirement is met.
10 Strength evaluation of a turbine blade 10 with film holes
And (3) acquiring a stress field of the turbine blade 10 with the film hole, adjusting the film hole type or the flow rate of cold air introduced into the single turbine blade 10 and the single-hole flow parameter if the maximum stress of the turbine blade 10 is lower than the design, and repeating the steps 2) to 9) until the strength requirement is met.
The present invention is not limited to the above preferred embodiments, and any modifications, equivalent substitutions and improvements made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (9)
1. A method for designing a gas film hole based on volume density distribution is characterized by comprising the following steps:
step 1, respectively calculating the temperature field and the stress field of a single turbine blade without an air film hole according to the working condition requirement of the turbine blade of the aero-engine;
step 2, performing air film hole pattern design according to the temperature field, introducing the flow of cold air in the single turbine blade and the flow of a single hole, and determining the number of the air film holesn；
Step 3, dividing the temperature field and the stress field of the single turbine blade intonThe volume domain is used for respectively carrying out volume domain gas film hole distribution design based on balanced temperature and volume domain gas film hole distribution design based on balanced stress on a single turbine blade;
and 4, respectively giving different weights to the volume domain based on the equilibrium temperature and the volume domain based on the equilibrium stress according to the working condition requirement of the turbine blade and the distribution characteristics of the temperature field and the stress field of different parts, balancing the nonuniformity of the temperature and the stress, and obtaining a turbine blade air film hole distribution scheme considering both the temperature field and the stress field.
2. The method for designing the gas film holes based on the bulk density distribution according to claim 1, wherein in step 1, the temperature field of the turbine blade without the gas film holes is obtained by performing heat transfer calculation on the turbine blade without the gas film holes according to the temperature, the pressure and the flow rate of the gas at the inlet and the outlet of the turbine blade and the temperature, the pressure, the flow rate and the flow rate of the cold gas.
3. The method for designing the film holes based on the bulk density distribution according to claim 1, wherein in step 1, the stress field of the non-film hole turbine blade is obtained by performing the strength calculation of the non-film hole turbine blade according to the temperature field, the rotating speed, the pressure and the load of the non-film hole turbine blade.
4. The method as claimed in claim 1, wherein the number of the film holes = single turbine blade cold air flow/single hole flow in step 2.
5. The method for designing the gas film hole based on the bulk density distribution according to claim 1, wherein the flow of designing the gas film hole distribution based on the volume domain of the equilibrium temperature in step 3 is as follows:
dividing the turbine blades intonIndividual volume domainV k Each volume domainV k The total temperature in the interior isT m ；
Will be provided withnThe air film holes are distributed in each volume areaV k In the inner part, the volume area of a single air film hole is realizedV k The total temperature in the furnace is the same;
wherein, the first and the second end of the pipe are connected with each other,
，T m is the average temperature of the individual film hole coverage,T t the total temperature of the entire turbine blade body,T i is a unit volumedvThe temperature of (2).
6. The method for designing the gas film hole based on the bulk density distribution according to claim 1, wherein the flow of designing the gas film hole distribution based on the volume domain of the equilibrium stress in the step 3 is as follows:
by dividing the turbine blades intonIndividual volume domainV s Each volume domainV s The total amount of internal stress isσ m ；
Will be provided withnThe air film holes are distributed in each volume areaV s In the volume region where the single air film hole is positionedV s The total internal stress is the same;
wherein the content of the first and second substances,
，σ m is the average stress covered by a single film hole,σ t the total amount of stress for the entire turbine blade body,σ i is a unit volumedvOf the stress of (c).
7. The method for designing the gas film holes based on the volume density distribution according to the claim 1, wherein according to the turbine blade gas film hole distribution scheme obtained in the step 4, the temperature evaluation of the turbine blade with the gas film holes and the strength evaluation of the turbine blade with the gas film holes are respectively carried out, and the turbine blade gas film hole distribution scheme is determined to be qualified after the design requirements are met; otherwise, adjusting the air film hole type or the flow of the cold air introduced into the single turbine blade and the single-hole flow parameters, and repeating the step 2-4 until the temperature evaluation and the strength evaluation are met and the design requirements are met.
8. The method for designing a gas film hole based on bulk density distribution according to claim 7, wherein the temperature evaluation method comprises: and performing heat transfer calculation on the turbine blade with the film hole to obtain a temperature field of the turbine blade with the film hole, wherein the highest temperature is lower than the long-term use temperature of the material, and the requirement on the temperature is judged to be met.
9. The method for designing the gas film hole based on the bulk density distribution according to claim 7, wherein the strength evaluation method comprises the following steps: and acquiring a stress field of the turbine blade with the film hole, and judging that the maximum stress of the turbine blade is lower than the allowable stress value in design use, wherein the maximum stress of the turbine blade is judged to meet the requirement.
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Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107725115A (en) * | 2017-04-28 | 2018-02-23 | 中国航发湖南动力机械研究所 | The aerofoil profile air film hole and electrode of aero-engine hot-end component |
| 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 |
| CN112780355A (en) * | 2021-02-25 | 2021-05-11 | 哈尔滨工业大学 | Supersonic turbine blade's cooling film hole distribution structure that diverges |
| CN114151139A (en) * | 2021-10-20 | 2022-03-08 | 中国航发四川燃气涡轮研究院 | Method for simulating flow of air film hole cold air layer on surface of turbine blade by adopting permeation model |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US11359301B2 (en) * | 2016-01-13 | 2022-06-14 | David Roberts Winn | Transparent and colorless hardcoating films for optical materials with a tunable index of refraction and scratch resistance, as formed from anodic aluminum films |
-
2022
- 2022-09-14 CN CN202211112716.1A patent/CN115203862B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107725115A (en) * | 2017-04-28 | 2018-02-23 | 中国航发湖南动力机械研究所 | The aerofoil profile air film hole and electrode of aero-engine hot-end component |
| 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 |
| CN112780355A (en) * | 2021-02-25 | 2021-05-11 | 哈尔滨工业大学 | Supersonic turbine blade's cooling film hole distribution structure that diverges |
| CN114151139A (en) * | 2021-10-20 | 2022-03-08 | 中国航发四川燃气涡轮研究院 | Method for simulating flow of air film hole cold air layer on surface of turbine blade by adopting permeation model |
Non-Patent Citations (1)
| Title |
|---|
| 双排气膜孔冷却热弹耦合特性的分析;王湛;《推进技术》;20170930;第38卷(第9期);2038-2045 * |
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