A gas turbine engine compressor tip clearance control structure and method of designing the same

Abstract

The present application relates to the technical field of gas turbine engine compressor tip clearance control, and discloses a gas turbine engine compressor tip clearance control structure and a design method thereof, the structure comprising an inner casing, an outer casing and a reinforcing rib arranged between the inner casing and the outer casing; the inner casing is a whole ring split structure, and the split seam penetrates the whole inner casing along the axial direction; the compressor tip clearance control structure can maintain the compressor tip clearance at a low level under multiple working conditions and ensure the structural safety.

Description

Technical Field

[0001] This invention relates to the field of gas turbine engine compressor tip clearance control technology, and discloses a gas turbine engine compressor tip clearance control structure and its design method. Background Technology

[0002] The compressor is a core component of a gas turbine engine, and its performance has a significant impact on the overall engine performance. The blade tip clearance of the compressor rotor has a substantial influence on the compressor's performance; excessive blade tip clearance leads to a significant decrease in compressor performance. Studies show that for every 1% increase in blade length in tip clearance, efficiency decreases by approximately 1.5%; and for every 1% decrease in efficiency, fuel consumption increases by approximately 2%. It is understood that the impact of blade tip clearance on fuel consumption accounts for approximately 67% of the total losses from blade profile and clearance sealing. Therefore, during the design phase, blade tip clearance should be minimized while ensuring engine safety.

[0003] Currently, in engineering practice, the common method for designing compressor tip clearance is to select the worst operating condition as the design condition, and determine the cold clearance based on the rotor-stator deformation difference under this condition and the design clearance. However, the rotor-stator deformation difference under other operating conditions will deviate from the design condition, which will result in a larger hot clearance under these conditions, leading to a decrease in compressor performance. As the maximum operating temperature of compressors continues to increase, this problem becomes increasingly prominent and has become one of the key factors restricting the improvement of compressor performance. Summary of the Invention

[0004] The purpose of this invention is to provide a gas turbine engine compressor tip clearance control structure and its design method, which can maintain the compressor tip clearance at a low level under multiple operating conditions, thereby improving the compressor performance.

[0005] To achieve the above-mentioned technical effects, the technical solution adopted by the present invention is as follows:

[0006] A gas turbine engine compressor tip clearance control structure includes:

[0007] The inner casing has an inner surface that is a flow channel surface and faces the rotor blade tip, and two reinforcing ribs are provided on the outer surface. The inner casing has a ring-shaped split structure, and the cut slit runs through the entire inner casing axially.

[0008] The outer casing is coaxially fitted on the outside of the inner casing, and its inner surface is in contact with the reinforcing ribs of the inner casing.

[0009] Furthermore, an abrasive coating is provided on the inner surface of the inner casing at a position opposite to the blade tip.

[0010] Furthermore, the inner surface of the inner casing is provided with a groove for installing the wear-resistant coating, and there are at least two reinforcing ribs, which are symmetrically distributed along the axial center plane of the groove.

[0011] Furthermore, the projection points of all the reinforcing ribs on the groove divide the groove into multiple segments along the engine axis.

[0012] To achieve the above technical effects, the present invention also provides a design method for a compressor tip clearance control structure of a gas turbine engine, used to obtain the compressor tip clearance control structure, comprising:

[0013] Based on the temperature distribution in the compressor flow channel of the aero-engine under preset operating conditions, determine the temperature value at the tip of the compressor rotor blades under the preset operating conditions.

[0014] Based on the flow channel dimensions corresponding to the compressor rotor blade tip, the given inner casing thickness and the radial height of the reinforcing ribs are designed, and the radius of the inner surface of the outer casing is obtained through analysis.

[0015] Based on the temperature value at the compressor rotor blade tip under the preset operating conditions, the radius of the inner surface of the outer casing, the thermal expansion coefficient of the inner casing, the thermal expansion coefficient of the outer casing, and the number of slits in the inner casing, the slit width of the inner casing is obtained by analysis.

[0016] Furthermore, the slit width of the inner casing is based on... Analysis yielded, among which The slit width of the inner casing. This refers to the number of slits in the inner casing. The radius of the inner surface of the outer casing. This refers to the temperature value at the compressor rotor blade tip under preset operating conditions. For temperature The coefficient of thermal expansion of the inner casing, For temperature The coefficient of thermal expansion of the outer casing, The ambient temperature.

[0017] Furthermore, it also includes:

[0018] Construct a three-dimensional simulation model that includes the inner casing, outer casing, and compressor rotor blades;

[0019] Using the first tip clearance value of a given compressor rotor blade at the compressor's maximum temperature as a constraint, the thickness of the outer casing as an adjustment variable, and the cold-state tip clearance design value of the rotor blade as an input condition, the range of adjustment variables that satisfy the compressor rotor blade tip clearance being near the first tip clearance value at the maximum temperature is obtained by analyzing the three-dimensional simulation model. This range is the design value range of the outer casing thickness.

[0020] Furthermore, the design value of the cold clearance at the rotor blade tip is based on... Analysis yielded, among which This is the design value for the cold clearance at the rotor blade tip. This represents the flow channel radius at the position corresponding to the compressor rotor blade tip. For rotor blade materials in The coefficient of thermal expansion at that time, for The expected value of the tip gap at temperature, and It is greater than or equal to the design value of the first blade tip clearance.

[0021] Compared with the prior art, the beneficial effects of the present invention are: through the above-mentioned structure and its design method, the present invention can achieve the technical goal of maintaining the compressor tip clearance at a low level under multiple operating conditions, and ensure the structural safety of the compressor. Attached Figure Description

[0022] Figure 1 This is a schematic diagram of the compressor tip clearance control structure of the gas turbine engine in the embodiment;

[0023] Figure 2 This is a schematic diagram of the inner casing structure in the embodiment;

[0024] Figure 3 This is a schematic diagram of the cross-section of the outer casing in the embodiment;

[0025] Figure 4 The diagram shows the flow channel surface of the inner casing of the present invention, the flow channel surface of the conventional casing, and the radial deformation of a typical rotor blade tip as a function of temperature within the flow channel in the embodiments.

[0026] Figure 5 This is a schematic diagram illustrating the change in blade tip clearance of each structure as a function of temperature within the flow channel, as shown in the embodiment.

[0027] Among them, 1. Inner casing; 2. Reinforcing ribs; 3. Slits; 4. Outer casing; 5. Wear-resistant coating; 6. Rotor blades. Detailed Implementation

[0028] The present invention will now be described in further detail with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the above-described subject matter of the present invention to the following embodiments; all technologies implemented based on the content of the present invention fall within the scope of the present invention.

[0029] Example

[0030] See Figures 1 to 5 A design method for a compressor tip clearance control structure of a gas turbine engine, wherein the compressor tip clearance control structure includes:

[0031] The inner casing 1 has an inner surface that forms a flow channel and faces the rotor blade tip. Two reinforcing ribs 2 are provided on the outer surface of the inner casing 1, and a wear-resistant coating 5 is provided on the inner surface of the inner casing 1. The inner casing 1 has a fully ring-segmented structure, with a slit 3 axially penetrating the entire inner casing 1, dividing it into three sections (see details). Figure 2 The inner casing 1 is made of a material with a coefficient of thermal expansion similar to that of the compressor rotor material.

[0032] The outer casing 4 is coaxially sleeved on the outside of the inner casing 1, and its inner surface is in contact with the reinforcing rib 2 of the inner casing 1. The outer casing 4 is a ring structure and is made of a material with a lower coefficient of thermal expansion than the compressor rotor material.

[0033] In this embodiment, when the temperature inside the flow channel... At lower temperatures, because the outer surface of the reinforcing rib 2 of the inner casing 1 is in contact with the inner surface of the outer casing 4, and the coefficient of thermal expansion of the inner casing 1 is higher than that of the outer casing 4, the radial thermal expansion of the inner casing 1 is limited by the outer casing 4. Simultaneously, because the inner casing 1 has a segmented structure with three circumferential slits 3, its thermal deformation is not completely limited. At this time, the inner casing 1 will undergo circumferential thermal expansion, reducing the width of its slits 3, while in the radial direction, it will not generate compressive stress with the outer casing 4. The equivalent coefficient of thermal expansion of the inner surface of the inner casing 1 is close to that of the outer casing 4 material. Since the coefficient of thermal expansion of the outer casing 4 material is less than that of the rotor material, the radial thermal deformation of the inner surface of the inner casing 1 is less than the radial thermal deformation of the rotor blade tip, and the radial clearance at the blade tip will decrease. As the temperature inside the flow channel increases... As the temperature inside the flow channel increases, the inner casing 1 will continue to expand circumferentially, causing the width of its slit 3 to continuously decrease. When the size increases to a certain value, the width of the slit 3 will decrease to 0, and the thermal expansion of the inner casing 1 will be completely restricted. When the temperature inside the flow channel... As the temperature continues to rise, the inner casing 1's thermal expansion is completely restricted, resulting in compressive stress in the radial direction and on the outer casing 4. At this point, because the outer casing 4 applies compressive stress to the inner casing 1, the radial thermal deformation of the inner casing 1 will be less than its radial thermal deformation in its free state. Its equivalent coefficient of thermal expansion on its inner surface will be between the coefficients of thermal expansion of the inner casing 1 and the outer casing 4, and lower than the coefficient of thermal expansion of the rotor material. Therefore, the blade tip radial clearance will continue to decrease. When the temperature inside the flow channel reaches its maximum value, the blade tip radial clearance decreases to the design value.

[0034] The trend of radial dimension of the inner casing 1 flow channel surface with the temperature inside the flow channel in this embodiment is shown below. Figure 4 The dashed line represents the trend of typical rotor blade tip radial dimension variation with flow channel temperature. Figure 4 Midpoint line, typical traditional casing flow channel surface radial dimension variation with flow channel temperature is shown in the figure. Figure 4 The solid line in the middle indicates that the traditional casing has a single-layer structure, and its material is the same as that of the outer casing 4.

[0035] Depend on Figure 4 It can be seen that when the temperature in the flow channel At lower temperatures, because the equivalent linear expansion coefficient of the inner casing 1 flow channel surface is close to the thermal expansion coefficient of the outer casing 4 material, the radial deformation of the inner casing 1 flow channel surface in this range is basically the same as the radial deformation of the conventional casing flow channel surface, and there is a certain difference compared with the radial deformation of the rotor blade tip. When the temperature inside the flow channel is higher than the temperature at which the width of the slit 3 will decrease to 0, the equivalent thermal expansion coefficient of the inner casing 1 flow channel surface is between the thermal expansion coefficients of the inner casing 1 and the outer casing 4 materials. Therefore, in this range, the radial deformation of the inner casing 1 flow channel surface will be greater than the radial deformation of the conventional casing flow channel surface but less than the radial deformation of the rotor blade tip. According to the commonly used blade tip clearance design method, the cold blade tip clearance is determined by formula... Confirmed, among which This refers to the cold blade tip gap. The radial deformation difference between the casing flow channel surface and the rotor blade tip when the temperature inside the flow channel is at its maximum. This is the design value for the tip clearance when the temperature inside the flow channel is at its maximum. (Due to the nature of this embodiment...) The value is smaller than that of the traditional structure, therefore the cold tip clearance of this embodiment will be smaller than that of the traditional structure.

[0036] The trend of tip clearance variation with flow channel temperature in this embodiment is shown below. Figure 5 The dashed line represents the trend of tip clearance variation with flow channel temperature in a typical conventional structure. Figure 5 The solid line in the figure represents the cold blade tip clearance. Sure.

[0037] Depend on Figure 5As can be seen, the cold-state tip clearance of the structure of this invention is reduced compared with the traditional structure, consistent with the previous description. The slope of the tip clearance variation curve is determined by the difference between the equivalent thermal expansion coefficient of the flow channel surface and the thermal expansion coefficient of the rotor material. When the difference is large, the radial deformation of the rotor tip and the radial deformation of the flow channel surface are significantly different, resulting in a larger change in the tip radial clearance, and therefore a larger curve slope. When the difference is small, the radial deformation of the rotor tip and the radial deformation of the flow channel surface are relatively small, resulting in a smaller change in the tip radial clearance, and therefore a smaller curve slope. As mentioned earlier, when the temperature inside the flow channel is lower than the temperature that reduces the width of the slit 3 to 0, the equivalent thermal expansion coefficient of the flow channel surface of the inner casing 1 of the compressor tip clearance control structure in this embodiment is close to the thermal expansion coefficient of the outer casing 4 material, and significantly different from the thermal expansion coefficient of the rotor material. Therefore, in this range, the slope of the tip clearance variation curve is large, and very close to the slope of the tip clearance variation curve of the traditional structure (because the casing material of the traditional structure is the same as that of the outer casing 4). When the temperature inside the flow channel is between the maximum operating temperature and the temperature at which the width of slit 3 is reduced to 0, the equivalent thermal expansion coefficient of the flow channel surface in this embodiment is between the thermal expansion coefficients of the inner casing 1 and the outer casing 4, and the difference from the thermal expansion coefficient of the rotor material is small. Therefore, in this range, the slope of the blade tip clearance variation curve of the structure of this invention is small. Compared with the conventional structure, the blade tip clearance of the structure of this invention decreases significantly when the temperature inside the flow channel is equal to the temperature at which the width of slit 3 is reduced to 0, achieving the technical goal of maintaining a low level of compressor blade tip clearance under multiple operating conditions. At the same time, the cold clearance of the structure of this invention is still maintained at a large level, which can ensure structural safety.

[0038] The compressor tip clearance control structure of this embodiment can effectively reduce the tip clearance when the temperature inside the flow channel is equal to the temperature at which the width of the slit 3 is reduced to 0. However, this increases the risk of rotor-stator rubbing during the transition state (operating state where the temperature inside the flow channel drops rapidly). Transition state rubbing is caused by a mismatch in the thermal inertia of the rotor and stator. Because the rotor has a larger thermal inertia and the stator has a smaller thermal inertia, when the temperature inside the flow channel drops too quickly, the stator temperature drops faster than the rotor temperature. This causes the radial contraction speed of the stator flow channel surface to be greater than that of the rotor tip, thus leading to rotor-stator rubbing. Since the slope of the tip clearance change curve is small when the temperature inside the flow channel is between the maximum operating temperature and the temperature at which the width of the slit 3 is reduced to 0, this exacerbates the risk of rotor-stator rubbing during the transition state. In view of this, this embodiment further proposes targeted improvement measures, one of which is to spray a thermal expansion coating on the A, B, and C surfaces of the outer casing 4 (see details...). Figure 3 Secondly, reinforcing ribs 2 are provided on the outer surface of the inner casing 1 to reduce the contact area between the inner casing 1 and the outer casing 4 (see details). Figure 1These two measures reduce heat loss from the outer casing 4, thus providing insulation. When the temperature inside the flow channel drops rapidly, the temperature of the inner casing 1 also drops rapidly. The outer casing 4, due to less heat loss, experiences a slower temperature drop, resulting in a temperature difference between the two casings in a short period. Since the outer casing 4 is at a higher temperature and the inner casing 1 is at a lower temperature, the restrictive effect of the outer casing 4 on the radial deformation of the inner casing 1 weakens. This causes the radial dimension of the flow channel surface of the inner casing 1 to tend to increase, thereby reducing the radial contraction rate of the inner flow channel surface and lowering the risk of transitional rubbing. These two measures have achieved good results in this embodiment.

[0039] In summary, the present invention can maintain a low level of compressor tip clearance under multiple operating conditions while ensuring high structural safety.

[0040] In this embodiment, the inner surface of the inner casing 1 is provided with a groove for installing the wear-resistant coating 5. There are at least two reinforcing ribs 2, which are symmetrically distributed along the axial center plane of the groove. The specific shape of the reinforcing ribs 2 can be designed in different styles according to actual needs, such as rectangular or trapezoidal, to better adapt to the structural characteristics of the inner casing 1. Simultaneously, the groove requires high machining precision to ensure its surface flatness, avoiding any impact on the installation quality and performance of the wear-resistant coating 5 due to surface unevenness. The reinforcing ribs 2 are made of the same material as the inner casing 1, possessing good strength and heat resistance, and can maintain stable performance under high temperature and high pressure environments, providing reliable support for the inner casing 1.

[0041] The reinforcing ribs 2 should be distributed as evenly as possible along the axial direction to improve the axial uniformity of the radial deformation of the inner casing 1. Therefore, in this embodiment, the projection points of all the reinforcing ribs 2 on the groove divide the groove into multiple segments along the engine axial direction, so that each segment of the groove can be more uniformly subjected to pressure and thermal load, effectively avoiding local stress concentration and improving the stability and durability of the entire structure.

[0042] The design method for the compressor tip clearance control structure of the gas turbine engine in this embodiment includes:

[0043] Step 1: Based on the temperature distribution in the compressor flow channel of the aero-engine under preset operating conditions, determine the temperature value at the tip of the compressor rotor blade under the preset operating conditions.

[0044] Step 2: Based on the flow channel dimensions corresponding to the compressor rotor blade tip, design the given thickness of the inner casing 1 and the radial height of the reinforcing rib 2, and analyze to obtain the radius of the inner surface of the outer casing 4.

[0045] Step 3: Based on the temperature value at the compressor rotor blade tip under the preset operating conditions, the radius of the inner surface of the outer casing 4, the thermal expansion coefficient of the inner casing 1, the thermal expansion coefficient of the outer casing 4, and the number of slits 3 in the inner casing 1, the width of the slits 3 in the inner casing 1 is analyzed and obtained; wherein, in this embodiment, the width of the slits 3 in the inner casing 1 is based on... Analysis yielded, among which The width of the slit 3 in the inner casing 1. The number of slits 3 in the inner casing 1. The radius of the inner surface of the outer casing 4 is [missing information]. This refers to the temperature value at the compressor rotor blade tip under preset operating conditions. For temperature The coefficient of thermal expansion of the inner casing 1 is as follows: For temperature The coefficient of thermal expansion of the outer casing 4 is as follows: The ambient temperature.

[0046] Step 4: Using the given first tip clearance value of compressor rotor blade 6 at the compressor's maximum temperature as a constraint, the thickness of the outer casing 4 as an adjustment variable, and the cold-state tip clearance design value of rotor blade 6 as an input condition, the three-dimensional simulation model is used to analyze and obtain the range of adjustment variables that satisfy the compressor rotor blade tip clearance being near the first tip clearance value at the maximum temperature. This range is the design value range of the outer casing 4 thickness.

[0047] The thickness of the outer casing 4 is designed to adjust or ensure the cold clearance. In this embodiment, the design value of the cold clearance at the tips of the 6 rotor blades is based on... Analysis yielded, among which This is the design value for the cold-state clearance of the 6 rotor blade tips. This represents the flow channel radius at the position corresponding to the compressor rotor blade tip. For rotor blade 6 material in The coefficient of thermal expansion at that time, for The expected value of the tip gap at temperature, and It is greater than or equal to the design value of the first blade tip clearance.

[0048] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A compressor tip clearance control structure for a gas turbine engine, characterized in that, include: The inner casing has an inner surface that is a flow channel surface and is opposite to the rotor blade tip, and at least two reinforcing ribs are provided on the outer surface; the inner casing has a ring-splitting structure, and the cut slit runs through the entire inner casing axially. The outer casing is coaxially fitted outside the inner casing, and its inner surface is in contact with the reinforcing ribs of the inner casing. The outer casing has a ring structure, and the thermal expansion coefficient of the inner casing is higher than that of the outer casing.

2. The gas turbine engine compressor tip clearance control structure according to claim 1, characterized in that, An abrasive coating is provided on the inner surface of the inner casing at a position opposite to the blade tip.

3. The gas turbine engine compressor tip clearance control structure according to claim 2, characterized in that, The inner surface of the inner casing is provided with a groove for installing the wear-resistant coating, and the reinforcing ribs are symmetrically distributed along the axial center plane of the groove.

4. The gas turbine engine compressor tip clearance control structure according to claim 3, characterized in that, The projection points of all the reinforcing ribs on the groove divide the groove into multiple segments along the engine axis.

5. A design method for a compressor tip clearance control structure of a gas turbine engine, used to obtain the compressor tip clearance control structure according to any one of claims 1-4, characterized in that, include: Based on the temperature distribution in the compressor flow channel of the aero-engine under preset operating conditions, determine the temperature value at the tip of the compressor rotor blades under the preset operating conditions. Based on the flow channel dimensions corresponding to the compressor rotor blade tip, the given inner casing thickness and the radial height of the reinforcing ribs are designed, and the radius of the inner surface of the outer casing is obtained through analysis. Based on the temperature value at the compressor rotor blade tip under the preset operating conditions, the radius of the inner surface of the outer casing, the thermal expansion coefficient of the inner casing, the thermal expansion coefficient of the outer casing, and the number of slits in the inner casing, the slit width of the inner casing is obtained by analysis.

6. The design method for the compressor tip clearance control structure of a gas turbine engine according to claim 5, characterized in that, The slit width of the inner casing is based on Analysis yielded, among which The slit width of the inner casing. This refers to the number of slits in the inner casing. The radius of the inner surface of the outer casing. This refers to the temperature value at the compressor rotor blade tip under preset operating conditions. For temperature The coefficient of thermal expansion of the inner casing, For temperature The coefficient of thermal expansion of the outer casing, The ambient temperature.

7. The design method for the compressor tip clearance control structure of a gas turbine engine according to claim 6, characterized in that, Also includes: Construct a three-dimensional simulation model that includes the inner casing, outer casing, and compressor rotor blades; Using the first tip clearance value of a given compressor rotor blade at the compressor's maximum temperature as a constraint, the thickness of the outer casing as an adjustment variable, and the cold-state tip clearance design value of the rotor blade as an input condition, the range of adjustment variables that satisfy the compressor rotor blade tip clearance being near the first tip clearance value at the maximum temperature is obtained by analyzing the three-dimensional simulation model. This range is the design value range of the outer casing thickness.

8. The design method for the compressor tip clearance control structure of a gas turbine engine according to claim 7, characterized in that, The design value of the cold clearance at the rotor blade tip is based on Analysis yielded, among which This is the design value for the cold clearance at the rotor blade tip. This represents the flow channel radius at the position corresponding to the compressor rotor blade tip. For rotor blade materials in The coefficient of thermal expansion at that time, for The expected value of the tip gap at temperature, and It is greater than or equal to the design value of the first blade tip clearance.

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