Turbine: Research progress on cooling structure of air-cooled turbine blades for aircraft engines

An efficient cooling structure is the key to avoiding thermal damage to air-cooled turbine blades, which directly affects blade cooling efficiency and aircraft engine stability. However, the efficient cooling structure leads to a more complex coherent effect between the mainstream and cold air flow, and the development of efficient cooling structures has always been constrained by processing technology. This article divides turbine blades into leading edge, middle chord, and trailing edge regions from the perspective of controlling cold air flow. It focuses on the research progress of cooling structures for air-cooled turbine blades in the past decade and the aerodynamic heat transfer characteristics under turbine rotation, including vortex impingement cooling, film hole shape, enhanced heat transfer structure inside the trailing edge, and partition rib shape. On this basis, the cooling effects of various structures were compared, and the defects of each cooling structure were pointed out. Finally, the future development direction and challenges of air-cooled turbine blades are proposed.

The design of turbine blade cooling structure is an important part of aircraft engine manufacturing. With the continuous advancement of research on high-temperature resistant materials and the adoption of more complex and efficient cooling structures, the working temperature that turbine blades can withstand has reached around 2200K. As of now, advanced cooling technologies include divergent cooling, film cooling, impact cooling, internal enhanced convection cooling, laminate cooling, and thermal barrier coatings. During the 12th Five Year Plan period, the effective temperature drop in domestic hanging experiments has reached 700K. On the other hand, turbine blades are prone to creep and elongation after being washed away by high-temperature gas, leading to the fracture of other components of the turbine. Therefore, in order to improve engine efficiency and avoid accidents, in-depth research on cooling technology for aircraft engine turbine blades is of great significance.

This article summarizes the research status of turbine blade cooling structures in the past decade, and summarizes and analyzes previous research results. It is worth noting that the large size and low operating temperature of gas turbine blades result in a cooling structure design that focuses on relatively simple, low-cost, and high degree of freedom directions. The cold air flow rate of F-class and G-class gas turbines accounts for 16% to 20% of the inlet flow rate of the compressor, and the diameter of the gas film holes on the gas turbine blades is 0.5-1mm. The internal cooling structure of aircraft engine turbine blades is more compact, with the proportion of cold air flow in aircraft engines reaching 20% to 30%, and there are many air film holes with a diameter of only 0.1-0.8mm. Although there are significant differences in the design details of cooling structures between gas turbines and aircraft engines, their design principles are universal. Therefore, several research works on the cooling structure of gas turbines have been discussed, and these summarized results can not only serve as the basis for the study of gas thermal coupling of turbine blades, but also as a reference for improving cooling technology, thereby providing support for the experimental and design of gas thermal coupling of air-cooled blades in aircraft engines.

Although divergent cooling has high cooling efficiency, the manufacturing of divergent cooling structures is difficult and their strength is low, which makes them unsuitable for engineering applications; The cooling principle of laminated boards is similar to that of divergent cooling. Cold air passes through microchannels in the laminated boards, absorbs heat, and is discharged from the air film; Thermal barrier coatings belong to the field of high-temperature materials and are not within the scope of this study; In addition, the cooling and heat transfer in the blade tip region of the turbine can be found in the literature, which is not included in the research content of this article. Therefore, in addition to the cooling techniques mentioned above, this article focuses on vortex impingement cooling, film hole parameters, enhanced heat transfer structures inside the trailing edge, and expansion plate shapes, as shown in Table 1. Due to geometric limitations, different cooling structures are used in different regions of the blade. As shown in Figure 1, impact cooling and film cooling are widely used in the leading edge region of blades; Film cooling and internally enhanced convective cooling are used in the mid chord region, which includes the suction surface and pressure surface; Due to the thin trailing edge, a combination of split slit air film cooling or inclined impact cooling with internal enhanced convective cooling is generally used. In addition, due to the differences in the working environment of turbine moving and stationary blades, the aerodynamic heat transfer characteristics of the cooling structure (especially the internal cooling channels) in each region of the moving blades under rotating conditions need to be analyzed separately.

This article divides the blades into leading edge, middle chord, and trailing edge regions, and focuses on discussing the influence of cooling structure, cooling structure parameters, and rotational effects on cooling efficiency in each region. The mechanisms and shortcomings of each structure are identified, and targeted suggestions are proposed. Finally, the development trend of cooling structure for air-cooled turbine blades in aircraft engines is discussed.

Innovation points of the paper

This article summarizes the research status of turbine blade cooling structures in the past decade, and summarizes and analyzes previous research results. The blade is divided into leading edge, middle chord, and trailing edge regions, and the influence of cooling structure, cooling structure parameters, and rotational effects on cooling efficiency in each region is emphasized. The influence mechanism and shortcomings of each structure are identified, and targeted suggestions are proposed. Finally, the development trend of cooling structure for air-cooled turbine blades in aviation engines is discussed.

Development Trends

In the 1950s, the United States first used investment casting to manufacture hollow turbine blades for aircraft engines. Subsequently, the manufacturing process of turbine blades continued to improve, from surplus machining to no surplus machining, and now to the best single crystal no surplus hollow blade machining. The complexity of the cooling structure also increased accordingly. However, improving the thrust to weight ratio of aircraft engines is an eternal pursuit of scientists. The constantly increasing working temperature forces the cooling structure of turbine blades to become more complex, while also placing higher demands on the manufacturing process of turbine blades. At the end of the 20th century, casting and cooling technology and 3D additive manufacturing technology emerged successively. The innovation of casting technology on the one hand reduced the processing cost of aircraft engine parts. For example, in July 2020, a lightweight aircraft engine 3D printed in Russia passed flight tests, and compared with conventional manufacturing, its production time was shortened to 1/20 of the original, and the cost was reduced to 1/2 of the original. It is planned to start mass production in 2021; On the other hand, making the processing of more complex and narrow internal flow channels feasible, the research direction of turbine cooling has begun to tilt towards composite cooling structures, such as confined space impact/vortex+air film+turbulence ribs, micro ribs/pin fins+pits, etc. Therefore, with the advancement of processing technology, the future development trend of air-cooled turbine blades is still to optimize the cooling structure to adapt to the constantly increasing working temperature; In addition, exploring the aerodynamic heat transfer characteristics of composite cooling structures in confined spaces under static and rotating states will be one of the main research directions; However, there are still challenges and difficulties. Specifically divided into the following three points. 

1. In terms of internal cooling, it is necessary to improve the internal cooling structure, control the flow of cold air reasonably, and balance the uniform distribution of cooling efficiency while improving cooling efficiency. For example, structures such as vortex cooling, serpentine channels, and microscale channels have advantages in this regard, but the efficiency of vortex cooling is lower than that of impingement cooling. Therefore, in the future, optimization research can be conducted based on the combination of the two to achieve complementary advantages, and a composite cooling structure can be formed with film cooling and turbulence ribs; The aerodynamic losses of serpentine and microscale channels are relatively large. In the future, the internal cooling channel that balances cooling efficiency and aerodynamic losses can be improved by using the gas thermal coupling method. Moreover, the emergence of 3D printing technology makes it feasible to arrange composite structures of protrusions, pits, and ribs in narrow channels.
2. Optimizing the shape and arrangement of external cooling structures in the leading edge, middle chord, and trailing edge regions of the blade remains a research hotspot, aiming to increase the coverage area of the air film and enhance its wall adhesion ability. For example, compared with traditional cylindrical holes, cooling structures such as irregular holes and shallow slot holes produce air films that are more widely distributed in the spanwise direction and extend longer in the upward direction. However, in the stagnation zone of the leading edge, the position of the air film holes needs to be carefully selected, which will affect the ratio of cold air flow to the suction and pressure surfaces, and thus affect the degree of air film coverage at the leading edge; Moreover, in the area near the blade shoulder and hub, due to the influence of vortices, the air film is easily lifted from the wall, resulting in high thermal stress; In addition, the reverse pressure gradient at larger curvature points on the suction surface can easily lead to film separation, which has not been effectively solved to date.
3. In the process of optimizing the cooling structure, the overall outlet area of the air film holes increases, and there are more hollow chambers inside the blades. In the absence of high-temperature resistant materials and limited manufacturing processes, improving cooling efficiency while reducing cold air volume and ensuring blade structural strength remains a difficult point in turbine blade cooling. It is worth considering that optimizing composite cooling under the premise of the same cold air flow rate and temperature is a paradox. Taking the impact+film composite cooling structure as an example, the heat capacity of the cold air flow is fixed. If the impact cooling efficiency is high, the temperature of the cold air flow extracted by the film holes will increase, and the film cooling efficiency will decrease. Therefore, in the optimization of composite cooling, emphasis should be placed on the allocation of cooling efficiency in order to achieve the best comprehensive cooling efficiency. In addition, due to the higher difficulty and cost of rotating experiments compared to stationary experiments, research on the aerodynamic heat transfer performance of rotating blade cooling structures, especially vortex cooling, composite cooling, and internal cooling of the trailing edge, is not yet sufficient. To avoid or utilize the influence of Coriolis force and buoyancy force, the cooling structure in the moving blades needs to be optimized separately, and as the thrust to weight ratio of the aircraft engine increases, the working temperature of the moving blades will inevitably increase, forcing the application of more complex cooling structures.

Summary

1. The use of impact cooling in the internal cooling structure of the leading edge can achieve a high heat transfer coefficient, but the distribution is uneven, while vortex cooling can effectively reduce the high temperature zone in the leading edge region of the blade. Therefore, the complementary advantages of vortex cooling and impact cooling have become the focus of improving the cooling efficiency of the leading edge. With the innovation of processing technology, the cooling channels and turbulence generators in the middle of the string are becoming more complex, and their detailed optimization remains a research hotspot. The relative position of the nozzle and rib inside the trailing edge, the position of the cold air outlet, and the arrangement and shape of the pin fins all deeply affect the cooling efficiency. Under the premise of the same size and shape, the convex part obtains higher cooling efficiency than the concave part, but the aerodynamic loss of the concave part is smaller.
2. The design principles for the outer air film holes of the leading edge and middle chord are: increasing the area of the air cooling outlet, reducing the momentum of the cold air flow outlet, increasing the coverage area of the air film, and reducing the mixing loss between the cold air flow and the mainstream. Installing expansion holes on the outer edge of the tail can increase cooling efficiency, but there is still research potential in the shape, size, position, and arrangement of the expansion holes.
3. Excessive hole spacing leads to a decrease in air film coverage, while insufficient spacing requires more cooling air. Therefore, there exists an optimal hole spacing for specific cooling structures. Moreover, improving the shape of the split lip plate and reducing its thickness can increase cooling efficiency. One of the future research directions is to use efficient and low-cost genetic algorithms to find the optimal hole arrangement or splitting parameters.
4. Due to the influence of rotation, the average cooling efficiency of the leading and trailing edges increases, while the cooling efficiency of the pressure surface decreases. However, further research is needed to investigate the effect of rotation on the cooling efficiency of the suction surface. Moreover, due to the difficulty and high cost of rotational experiments, research on the aerodynamic heat transfer mechanism of rotating blade cooling structures, especially vortex cooling and composite cooling, is not yet sufficient.
5. With the development of aviation engines and the innovation of manufacturing processes, the cooling structure of air-cooled turbine blades is evolving towards optimizing single cooling structures and composite cooling structures. However, potential issues such as aerodynamic losses and blade structural strength have also become prominent.