阅读说明：本技术 一种涡轮叶片尾缘折线式排气劈缝结构 (Turbine blade trailing edge broken line type exhaust split structure ) 是由 吕东 周亦胄 王晓放 孔星傲 王楠 于 2019-11-29 设计创作，主要内容包括：本发明属于航空发动机涡轮冷却技术领域,涉及一种涡轮叶片尾缘折线式排气劈缝结构。一种涡轮叶片尾缘折线式排气劈缝结构,包括空心涡轮叶片、内腔冷气通道、尾缘排气劈缝通道和尾缘劈缝隔肋；所述空心涡轮叶片内部设有内腔冷气通道,供低温冷却气体在叶片内部流动,对叶片进行冷却。空心涡轮叶片尾缘设有并排排列的尾缘劈缝隔肋,并排排列的尾缘劈缝隔肋之间形成尾缘排气劈缝通道。本发明通过将尾缘排气劈缝设计成倾斜折线式排气,减小冷却气在劈缝内的转折角,且冷气由一次转折变为两次或多次转折,从而减小叶片内腔冷气流动阻力和损失,可以使流阻减小约19％。(The invention belongs to the technical field of aircraft engine turbine cooling, and relates to a turbine blade trailing edge broken line type exhaust split structure. A broken-line exhaust split structure of the tail edge of a turbine blade comprises a hollow turbine blade, an inner cavity cold air channel, a tail edge exhaust split channel and a tail edge split partition rib; the hollow turbine blade is internally provided with an inner cavity cold air channel for low-temperature cooling gas to flow inside the blade to cool the blade. The tail edges of the hollow turbine blades are provided with tail edge slit partition ribs which are arranged side by side, and tail edge exhaust slit channels are formed between the tail edge slit partition ribs which are arranged side by side. The invention designs the tail edge exhaust cleft into the inclined broken line type exhaust, reduces the turning angle of the cooling air in the cleft, and the cooling air is turned twice or for multiple times from one turning, thereby reducing the flow resistance and loss of the cooling air in the inner cavity of the blade and reducing the flow resistance by about 19 percent.)

1. A broken-line exhaust split structure of the tail edge of a turbine blade is characterized by comprising a hollow turbine blade (1), an inner cavity cold air channel (2), a tail edge exhaust split channel (3) and a tail edge split partition rib (4);

an inner cavity cold air channel (2) is arranged in the hollow turbine blade (1), tail edge split joint partition ribs (4) which are arranged side by side are arranged at the tail edge of the hollow turbine blade (1), a tail edge exhaust split joint channel (3) is formed between the tail edge split joint partition ribs (4) which are arranged side by side, the structure of each tail edge split joint partition rib (4) forms a broken line shape, the shape of each broken line is controlled by a partition rib central line (5), the partition rib central line (5) is formed by connecting at least two line segments with different angles, at least one section of each line segment is inclined relative to the horizontal plane, the width of each tail edge split joint partition rib (4) is symmetrically distributed along the partition rib central line (5), and the included angles between the cold air inlet end and the cold air outlet end of the hollow turbine blade (1) and the horizontal plane are an incident angle ∠ A1 and an.

2. The turbine blade trailing edge fold line exhaust cleft structure of claim 1 wherein the incidence angle ∠ a1 and the exit angle ∠ a2 are ∠ a1> ∠ a 2.

3. The turbine blade trailing edge broken line type exhaust split structure as claimed in claim 1 or 2, wherein the incident angle ∠ A1 is 15-45 °, the emergence angle ∠ A2 is 0-30 °, and the cold air turning angle ∠ A is 90 ° - ∠ A1.

Technical Field

The invention belongs to the technical field of aircraft engine turbine cooling, and relates to a turbine blade trailing edge broken line type exhaust split structure.

Background

Increasing the turbine front gas temperature can greatly improve the efficiency of aircraft engines and gas turbines, but the current turbine front gas temperature is far beyond the limit that the materials used can bear, so that the development of more effective turbine blade cooling technology is urgent. At present, the turbine blade is generally designed in a hollow mode, heat is taken away through enhanced convection heat exchange of cooling air in the turbine blade, and an air film is formed to cover and isolate fuel gas for heating when the turbine blade is discharged, so that the turbine blade is a main solution for the cooling problem of the turbine blade, and the turbine blade is mainly focused and pursued by the requirements of larger internal heat exchange area, smaller cold air flow resistance, higher heat exchange efficiency, larger air film coverage area, smaller structural strength damage and the like.

The tail edge area of the turbine blade is heated by combustion gas on the two sides of the basin side and the back side of the blade, and is structurally thin and difficult to form a hollow cooling structure, so that the tail edge area is an area difficult to cool in the blade, is an area with high wall surface temperature and easy ablation in work, and is a difficult problem to be mainly solved in the blade cooling design. At present, a half-open horizontal exhaust split structure is usually adopted for cooling the tail edge of the blade, the structure can convert cold air flowing in a cooling channel in the blade along the radial direction into cold air flowing along the chord direction, after strengthened convection cooling is formed on the wall surface of the channel and the rib structure, the cold air is exhausted from a narrow gap (called a split gap) at the edge of the basin side of the blade, and an air film is formed on the part of the tail edge to cover and isolate heating of fuel gas, and the typical structure is shown in figure 1 and is basically characterized in that the split gap horizontally exhausts the air. The tail edge slot-splitting cooling structure has large flow resistance and low cooling effect, and also has certain damage to the structural strength of the blade.

Disclosure of Invention

Aiming at the defects of the existing horizontal exhaust tail edge split seam cooling technology, the broken line type exhaust split seam structure of the tail edge of the turbine blade is provided, and the broken line type exhaust split seam structure of the tail edge of the turbine blade is designed to be the broken line type exhaust structure, so that the turning angle of cooling gas can be effectively reduced, the flow resistance and loss are reduced, the covering area of a gas film is increased, the temperature level of the blade is reduced, the load resistance of the structure is improved, and the casting technology performance of the blade is improved.

The technical scheme of the invention is as follows:

a broken-line exhaust split structure of the tail edge of a turbine blade comprises a hollow turbine blade, an inner cavity cold air channel, a tail edge exhaust split channel and a tail edge split partition rib;

the hollow turbine blade is internally provided with an inner cavity cold air channel for low-temperature cooling gas to flow inside the blade to cool the blade. The tail edges of the hollow turbine blades are provided with tail edge slit partition ribs which are arranged side by side, and tail edge exhaust slit channels are formed among the tail edge slit partition ribs which are arranged side by side so as to exhaust cooling air out of the blades and simultaneously carry out air film covering cooling on the tail edges of the blades. The structure of the tail edge slot partition rib can increase the heat exchange area in the blade, and guide the cooling air in the inner cavity of the blade to make the flowing direction of the cooling air turn.

The structure of the tail edge slit partition rib forms a broken line shape, the shape of the broken line shape is controlled by a partition rib central line, the partition rib central line is formed by connecting at least two line segments with different angles, at least one segment of the line segment is inclined relative to a horizontal plane, the width of the tail edge slit partition rib is symmetrically distributed along the partition rib central line, the included angles between the connected inclined line segments and the horizontal plane at the cold air inlet end and the cold air outlet end of the hollow turbine blade are an incident angle ∠ A1 and an emergent angle ∠ A2 respectively, the relationship between the two angles is ∠ A1 to ∠ A2, and the included angle between the flow direction of cold air after entering the tail edge slit and the flow direction before entering the tail edge slit, namely the cold air turning angle ∠ A is smaller than 90.

Further, the incident angle ∠ A1 can be 15-45 degrees, the emergence angle ∠ A2 can be 0-30 degrees, and the cold air turning angle ∠ A is 90- ∠ A1.

In the original structure, as shown in fig. 1, the center line of the partition rib is a horizontal straight line, the air inlet/outlet angles ∠ a1 and ∠ a2 of the cold air in the tail edge cleft are both 0 °, the cold air turning angle ∠ a is about 90 °, the turning angle is too large, the flow loss is large, the air film coverage area is small, and the strength damage is large.

The principle of the invention is as follows:

the inclined broken line type structure of the invention reduces the flow resistance and loss of the cold air in the inner cavity of the blade:

compared with the original horizontal exhaust tail edge slot structure, the invention has the beneficial effects that the cold air flow resistance and loss are reduced. For the horizontal exhaust tail edge slot structure, the cooling gas needs to complete 90-degree direction turning at high speed in a narrow space, and such flow can form approximate step flow on the leeward side of the end part of the partition rib to generate low-speed vortex, as shown in fig. 3, the vortex flow not only generates energy dissipation due to the strong friction thereof, but also extrudes the main flow to force the main flow to generate additional direction turning and energy loss. The driving of the cold air flow needs to extract energy from the whole engine, so the high resistance flow can increase the power consumption of the whole engine, thereby reducing the efficiency. When the partition ribs of the split seam are changed into an inclined broken line type structure from the horizontal state, the turning angle of the cooling gas in the split seam is changed into an acute angle from an approximate right angle, and the numerical value is reduced by more than 35%. And the cooling gas needs to be bent twice or more in the broken line type channel, and compared with the method of completing the bending once, the angle of each bending is smaller. Under the same flow, because the turning flow becomes more moderate, especially the leeward side of the partition rib does not form obvious steps any more, the vortex can not be formed, the flow becomes smooth, the flow resistance and loss are reduced, and the whole efficiency of the engine is improved.

2, the inclined broken line type structure improves the gas film covering effect of the tail edge area of the blade:

the exhaust tail edge slit has a gas film covering direction approximately parallel to the gas combustion, and the gas combustion cannot deflect radially under the condition of being clamped by the gas combustion, so that the gas film covering width of a single slit is approximately the same as the actual width of the slit, and the rib area between two adjacent slits is almost uncovered by the gas film, the area which cannot be covered by the gas film is called a gas film covering dead area, the cooling effect is poor, high-temperature ablation is easily caused, the gas film outflow direction of the oblique fold line type exhaust tail edge slit forms a certain included angle with the gas combustion flowing direction, which can be approximately considered as ∠ A2, because the gas combustion has the clamping effect on the gas film outflow, the cold gas is deflected in the flowing direction and gradually parallel to the gas combustion flowing direction, and the covering of the partial rib area is realized, so that the covering width of the actual gas film covering width is increased relative to the gas combustion film outflow direction, and the covering safety of the blade can be completely improved under the condition that the safety covering of the two adjacent slits is not increased, and the safety of the blade can be completely improved.

3, the inclined broken line type structure improves the load resistance of the hollow structure at the tail edge of the blade:

the turbine blades are subjected in operation mainly to the following loads: centrifugal loads caused by high-speed rotation, aerodynamic loads exerted by the gas flow, vibratory loads due to vibrations, which exert a tendency to deform in tension, torsion and bending on the blade base body and to generate corresponding stresses, and thermal stresses due to thermal expansion non-uniformities. These stresses couple together and act on the component alternately over time, and after exceeding the limits that the material can withstand, failure can occur. The blade design requires the use of a minimum of materials optimized to withstand these loads, avoiding the presence of high stress areas and damage to the blade. The tail edge crack structure is a vulnerable area with higher stress level in the turbine blade, the vulnerable area is firstly positioned at the thinnest part in the blade profile, and the hollow structure of the crack greatly weakens the strength, so that stress concentration can be generated at the opening of the crack. As shown in fig. 5(b), the stress level of the tail edge split structure shows a periodic law along the leaf height direction under the action of the radial tensile load, and shows a peak of the stress at the split opening and a trough at the partition rib. For a horizontal exhaust tail edge split, the periodicity is more pronounced and the peak value is higher because of the high degree of hollowing of the material at the tail edge, lack of effective support and reinforcement at the cross-sectional location shown in fig. 5 (a). When the inclined broken line type exhaust tail edge slit structure is adopted, as shown in fig. 5(c), the partition rib structure can be connected with the two sides of the basin back of the blade at any section position, so that the tail edge structure is strengthened, the stress level is improved, the load resistance of the structure is improved, and the safety and the reliability of the whole machine are improved. As can be seen from a comparison of the stress distribution along the blade height of the trailing edge shown in FIG. 5(b), the peak stress of the oblique broken line exhaust trailing edge cleft structure is significantly less than that of the horizontal exhaust pattern.

4, the inclined broken line type structure improves the blade casting manufacturability:

the turbine blade is a part with higher manufacturing cost, usually adopts zero-allowance precision casting, has lower qualification rate, and particularly needs to adopt a directional solidification process for a monocrystalline turbine blade (the molten metal is crystallized into a crystal grain during solidification, and has better mechanical property at high temperature and gas corrosion resistance compared with a polycrystalline blade), namely the molten metal is gradually cooled to a solidification state from bottom to top. Since the volume of the metal is reduced from liquid to solid, the liquid metal must be ensured to have better fluidity in the casting shell during the structural design and the corresponding manufacturing process design of the blade, and the liquid metal can be timely supplemented to the space generated after solidification shrinkage, which is called feeding. If the feeding path of the molten metal is blocked and the feeding is insufficient, dense fine holes are generated when the molten metal in the corresponding region is solidified, which is called a loosening phenomenon. The looseness can cause the mechanical property of the material and the load resistance of the blade to be greatly reduced, belongs to a serious quality problem and must be scrapped. For the horizontal exhaust tail edge slot structure, as shown in the left side of fig. 6, because the liquid-solid interface of metal can take a slightly downward concave shape because the temperature of the outer side of the blade is lower than that of the inner side, a part of feeding paths of the rib structure can be blocked, so that the horizontal rib structure is easy to cause a loosening problem, and the blade has poor casting manufacturability. For the oblique broken line type exhaust tail edge slot splitting structure, as shown in the right side of fig. 6, the rib separating structure forms a certain included angle with the liquid-solid boundary, and the oblique rib separating structure just forms a molten metal feeding channel, so that the generation of looseness is avoided, and the casting manufacturability of the blade is improved.

The invention has the beneficial effects that:

the invention provides a turbine blade broken line type exhaust tail edge split seam cooling structure, which reduces the turning angle of cooling air in a split seam by designing the tail edge exhaust split seam into inclined broken line type exhaust, and changes the cold air from one turning to two or more turning, thereby reducing the cold air flow resistance and loss of the blade inner cavity and reducing the flow resistance by about 19%. Utilize the carminative gas film of slope broken line formula to flow out direction and gas air current direction and become certain contained angle, make the gas go out to the gas film and have the parcel and hold the effect and improve blade tail edge district gas film covering effect, can make the heat transfer reinforcing about 10%. The load-resisting capacity of the hollow structure of the blade tail edge is improved by designing the partition ribs into an inclined broken line type structure, and compared with a horizontal exhaust tail edge split seam cooling structure, the load-resisting capacity of the blade tail edge structure can be improved by about 20%. In addition, the inclined rib structures can form molten metal feeding channels, so that the blade casting manufacturability is improved, and the casting yield can be improved by about 14%.

Drawings

FIG. 1(a) is a view showing a horizontal exhaust cleft at the trailing edge of a conventional turbine blade.

FIG. 1(b) is a C-C sectional view of a horizontal exhaust cleft structure of the tail edge of the prior turbine blade.

FIG. 1(c) is a partial enlarged view of a prior art turbine blade trailing edge horizontal exhaust slot structure.

FIG. 2(a) is a view showing a broken-line exhaust cleft at the trailing edge of a turbine blade.

FIG. 2(b) is a partial enlarged view of a broken line type exhaust cleft structure at the trailing edge of the turbine blade.

FIG. 3 is a comparison diagram of cooling gas flow conditions in two tail edge cleavage slit structures.

FIG. 4 is a comparison graph of the gas film covering effect of two tail edge slit structures.

FIG. 5(a) is a schematic sectional view of a blade profile of a conventional turbine blade trailing edge horizontal exhaust slot structure.

FIG. 5(b) is a graph showing the comparison of stress levels of two types of tail edge split structures.

FIG. 5(c) is a schematic view of the blade profile cross-sectional shape of the turbine blade trailing edge fold line type exhaust cleft structure.

FIG. 6 is a comparative directional solidification casting process between two types of tail edge slot structures.

In the figure, 1 is a hollow turbine blade, 2 is an inner cavity cold air channel, 3 is a tail edge exhaust split gap channel, 4 is a tail edge split gap partition rib, 5 is a partition rib central line, 6 is an incidence angle ∠ A1, 7 is an emergence angle ∠ A2, 8 is a cold air turning angle ∠ A, 9 is a cold air channel partition wall, 10 is an air film coverage area, 11 is an air film coverage width, 12 is a high stress area at a split gap opening, and 13 is a low stress area at the split gap partition rib.

Detailed Description

In order that the present invention may be more readily and clearly understood, a more particular description of the invention briefly described above will be rendered by reference to specific embodiments that are illustrated in the appended drawings.