Gas supply cooling system and method of designing the same

Abstract

The application discloses a kind of air cooling system and its design method, including sequentially arranged along axis first turbine, turbine guider, second turbine, further include the first flow guide disc, second flow guide disc sequentially arranged along axis in turbine guider radial inner side;Turbine guider includes upper edge plate, guider blade and lower edge plate, upper edge plate is opened with through hole, guider blade is opened with front edge cooling flow channel, lower edge plate is opened with through hole, the axis of through hole is inclined with the tangent of turbine guider at through hole place, to make the airflow that flows out through hole have with first turbine rotating direction same direction speed component;Lower edge plate is distributed with seal ring, first flow guide disc, seal ring and lower edge plate jointly enclose front stage cooling seal cavity, second flow guide disc, seal ring and lower edge plate jointly enclose rear stage cooling seal cavity;It can reduce cold air in high radius area wind resistance temperature rise, improve the heat exchange capacity of this area, effectively reduce the temperature level of wheel disc.

Description

Technical Field

[0001] This invention relates to the field of turbine engine technology, and in particular, to an air supply cooling system. Furthermore, this invention also relates to a design method for an air supply cooling system. Background Technology

[0002] The turbine disk is a core component of a gas turbine engine. Its harsh operating environment—high temperature, high pressure, and high speed—places high demands on turbine disk design, with high temperature significantly impacting its strength and lifespan. For example... Figure 1 As shown, in existing turbine disk cooling schemes, cold air is usually introduced into the disk cavity from the lower radius region of the disk, and the cold air flows radially to cool the disk surface. In this type of design, the temperature of the cold air increases along the path. During the cooling process of the cold air flowing radially along the disk, the wind resistance temperature rise is large due to the difference between the circumferential speed of the cold air and the local rotational speed of the end face of the disk being cooled, which is extremely unfavorable for cooling the high radius region of the disk.

[0003] Therefore, a pre-swirl design can be used to reduce the air resistance and temperature rise of the cooling air, thereby improving the cooling effect of the turbine disk. Pre-swirl designs can be divided into high-position pre-swirl and low-position pre-swirl based on the height of the pre-swirl gas outlet radius. Low-position pre-swirl is beneficial for reducing air leakage and performance loss, and its pre-swirl structure is small in size; however, the reduction in air temperature rise is not significant, and it consumes a large amount of turbine disk pump power. High-position pre-swirl has its gas outlet located in the high-radius region of the turbine disk cavity, resulting in a larger size of its pre-swirl structure and the structure supporting it. Furthermore, high-position pre-swirl is close to the turbine flow channel, leading to less space for designing a pressurization sealing device and difficulty in controlling leakage. Therefore, there is currently no technical solution using a high-position pre-swirl design for cooling the turbine disk. Summary of the Invention

[0004] This invention provides an air supply cooling system and its design method to solve the technical problem of poor cooling effect in the high radius area of ​​the disc in the prior art.

[0005] According to one aspect of the present application, a gas supply cooling system is provided, comprising a first stage turbine, a turbine guide vane, a second stage turbine arranged in sequence along an axial direction, and further comprising a first flow guide disc and a second flow guide disc arranged in sequence along the axial direction and located radially inside the turbine guide vane; the turbine guide vane comprises an upper edge plate, a guide vane blade, and a lower edge plate, the upper edge plate is provided with a through hole for introducing cold gas, the guide vane blade is provided with a leading edge cooling flow channel in communication with the through hole, and the lower edge plate is provided with a through hole in communication with the leading edge cooling flow channel, the axial center line of the through hole is obliquely arranged with respect to the tangent line of the turbine guide vane at the through hole, so that the airflow flowing out of the through hole has a velocity component in the same direction as the rotation direction of the first stage turbine; the lower edge plate is provided with a sealing ring, the first flow guide disc, the sealing ring, and the lower edge plate jointly enclose a front stage inter-stage cooling sealing cavity, the front stage inter-stage cooling sealing cavity is in communication with a gas flow channel, the second flow guide disc, the sealing ring, and the lower edge plate jointly enclose a rear stage inter-stage cooling sealing cavity, and the rear stage inter-stage cooling sealing cavity is in communication with the gas flow channel.

[0006] Further, the axial center line of the through hole 231 forms a pre-rotation angle a with the tangent line of the turbine guide vane 2 at the through hole 231, and the pre-rotation angle a is 30-60°.

[0007] Further, the first flow guide disc comprises a first shaft ring portion extending in the axial direction and arranged opposite to the inner ring surface of the sealing ring, the second flow guide disc comprises a second shaft ring portion extending in the axial direction and arranged opposite to the inner ring surface of the sealing ring, and the first shaft ring portion and the second shaft ring portion are overlapped; the first flow guide disc, the first shaft ring portion, the second shaft ring portion, and the second flow guide disc enclose an internal flow guide disc cavity b; a plurality of grid teeth are arranged on the first shaft ring portion and the second shaft ring portion respectively and matched with the gap of the sealing ring.

[0008] According to another aspect of the present application, a design method of a gas supply cooling system is also provided, which is suitable for the above-mentioned gas supply cooling system and comprises the following steps:

[0009] S1, determining the rotor speed Ng, the number of guide vane blades N, the cold gas flow limit value m0_max, the material temperature limit T1_max of the first flow guide disc, and the material temperature limit T2_max of the second flow guide disc;

[0010] S2, determining the minimum sealing cold gas flow m1_min of the front stage inter-stage and the minimum sealing cold gas flow m2_min of the rear stage inter-stage according to the guide vane front stage inter-stage rotating structure, the guide vane rear stage inter-stage rotating structure, and the inter-stage gas parameters to meet the gas sealing;

[0011] S3, determining the flow m0 passing through the through hole according to the flow continuity principle;

[0012] S4, determining the hole diameter D of the through hole;

[0013] S5, determining a pre-swirl angle a of the through hole;

[0014] S6, calculating actual sealing cold air flow m1 between the front stage and actual sealing cold air flow m2 between the rear stage, and determining whether m1 >= m1_min and m2 >= m2_min are satisfied;

[0015] S7, if the actual sealing cold air flow does not satisfy the minimum sealing flow requirement, repeating S6 and S7 until m1 >= m1_min and m2 >= m2_min are satisfied;

[0016] S8, determining the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc, and determining whether T1 <= T1_max and T2 <= T2_max are satisfied;

[0017] S9, if the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc do not satisfy the design requirement, increasing the flow m0 through the through hole, and repeating S4-S8 until T1 <= T1_max and T2 <= T2_max are satisfied;

[0018] S10, outputting the design results, i.e., the hole diameter D of the through hole, the pre-swirl angle a of the through hole, the flow m0 through the through hole, the actual sealing flow m1 between the front stage, the actual sealing flow m2 between the rear stage, the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc.

[0019] Further, step S4 specifically comprises calculating the hole diameter D of the through hole according to the total flow m0 through the through hole, the number N of the guide vane, the number n of the through hole arranged at the bottom of each guide vane and the cold air parameters.

[0020] Further, step S5 specifically comprises: obtaining the outlet radius Rout of the through hole, and determining the pre-swirl angle a of the through hole according to the outlet radius Rout of the through hole and the rotor speed Ng.

[0021] Further, step S6 specifically comprises: carrying out fluid dynamics calculation according to the sealing structure parameters of the front stage structure of the guide vane, the sealing structure parameters of the rear stage structure of the guide vane, the radial gap parameters of the grid and the sealing ring, the rotor speed Ng, the hole diameter D of the through hole, the number n of the through hole at the bottom of each guide vane, the pre-swirl angle a, the inter-stage gas static pressure, the inter-stage gas total temperature, the cold air inlet total pressure, the cold air inlet total temperature, determining the actual sealing cold air flow m1 between the front stage and the actual sealing cold air flow m2 between the rear stage, and determining whether m1 >= m1_min and m2 >= m2_min are satisfied.

[0022] Further, the adjusting the radial gap between the labyrinth and the seal ring in step S7 specifically comprises: when m1 < m1_min, reducing the radial gap between the labyrinth and the seal ring; when m2 < m2_min, increasing the radial gap between the labyrinth and the seal ring; until m1 >= m1_min and m2 >= m2_min are satisfied.

[0023] Further, step S8 specifically comprises: according to M1, M2, the temperature corresponding to M1, the pressure corresponding to M1, the temperature corresponding to M2 and the pressure corresponding to M2 determined in S7, calculating the heat exchange boundary of the first and second flow guide discs, performing temperature analysis of the first and second flow guide discs to obtain the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc.

[0024] Further, step S9 further comprises: when the flow m0 through the through hole increases to the initial cold gas flow limit value m0_max, and the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc still do not satisfy the design requirements, increasing the value of the cold gas flow limit value m0_max, and repeating S4-S8 until T1 <= T1_max and T2 <= T2_max are satisfied.

[0025] The present application has the following beneficial effects:

[0026] The air supply cooling system of the present application cools the guide vane when the cold gas flows through the leading edge cooling flow channel, and is forced to generate a velocity component in the first turbine rotation direction when entering the front stage interstage cooling seal cavity through the through hole, so that the air temperature sensed by the rotor is reduced and the cooling efficiency is improved; the cold gas is divided into two parts in the front stage interstage cooling seal cavity, one part is discharged forward into the gas flow channel between the first turbine and the guide vane, and this part of cold gas completes the front stage interstage gas sealing and the cooling of the first flow guide disc; the other part is discharged backward through the entry into the rear stage interstage cooling seal cavity, and then discharged into the flow channel between the second turbine and the guide vane after sealing and cooling the rear stage; the outlet radius Rout of the through hole is relatively high (the radial height from the rotor structure axis), which is in the high radius area of the turbine disc rear hot environment, can reduce the cold gas wind resistance temperature rise in this area, improve the heat exchange capacity of this area, effectively reduce the temperature level of the turbine disc, and at the same time, since the through hole is integrated on the lower edge plate, the structure is simplified, the adverse structural factors of the weight caused by the large size of the traditional high position pre-rotation structure are avoided, the pre-rotation radius of the cold gas is improved, the cold gas temperature rise is reduced, combined with the seal flow path design, the potential of the cooling air is fully utilized, so that the air utilization efficiency is improved and the performance loss is reduced.

[0027] In addition to the purposes, features and advantages described above, the present application has other purposes, features and advantages. The present application will be further described below with reference to the drawings. BRIEF DESCRIPTION OF DRAWINGS

[0028] The accompanying drawings, which form a part of this application, are included to provide a further understanding of the application and are incorporated in and constitute a part of this specification. The illustrations are shown schematically in the drawings where:

[0029] Figure 1 is a schematic diagram of the cold air flow direction of the air supply cooling system in the prior art;

[0030] Figure 2 is a schematic diagram of the structure of the air supply cooling system of the preferred embodiment of the present application Figure 1 ;

[0031] Figure 3 is a schematic diagram of the structure of the air supply cooling system of the preferred embodiment of the present application Figure 2 ;

[0032] Figure 4 is a schematic diagram of the cold air flow direction of the air supply cooling system of the preferred embodiment of the present application;

[0033] Figure 5 is a schematic diagram of the pre-rotation of the air guide hole of the air supply cooling system of the preferred embodiment of the present application;

[0034] Figure 6 is a schematic diagram of the pre-rotation analysis of the air guide hole of the air supply cooling system of the preferred embodiment of the present application;

[0035] Fig. 7(a) is a relative total temperature distribution of the inter-stage seal cooling air in the prior art; Fig. 7(b) is a relative total temperature distribution of the inter-stage seal cooling air in the preferred embodiment of the present application.

[0036] Legend:

[0037] 1, first stage turbine; 101, first side lobe; 2, turbine guide vane; 21, upper shroud; 211, through hole; 22, guide vane blade; 221, leading edge cooling flow channel; 23, lower shroud; 231, via hole; 3, second stage turbine; 301, second side lobe; 4, first flow guide disc; 401, first lobe; 402, first collar portion; 403, louver; 5, second flow guide disc; 501, second lobe; 502, second collar portion; 6, seal ring; la, front inter-stage cooling seal cavity; lb, front inter-stage seal buffer cavity; 2a, gas flow channel; 3a, rear inter-stage cooling seal cavity; 3b, rear inter-stage seal buffer cavity; 4a, first gap passage; 4b, internal ventilation disc cavity; 5a, second gap passage. DETAILED DESCRIPTION

[0038] The embodiments of the present application will be described in detail with reference to the drawings, but the present application can be implemented in various different ways as defined and covered by the following.

[0039] Please refer to Figures 2 to 6The air supply cooling system of the embodiment is suitable for a turbine engine, and the air supply cooling system comprises a first-stage turbine 1, a turbine guide vane 2, and a second-stage turbine 3 arranged in sequence along an axial direction of the turbine engine, and further comprises a first flow guide disc 4 and a second flow guide disc 5 arranged in sequence along the axial direction and located radially inside the turbine guide vane 2. The turbine guide vane 2 comprises an upper edge plate 21, a guide vane blade 22, and a lower edge plate 23. The upper edge plate 21 is provided with a through hole 211 for introducing cold air. The guide vane blade 22 is provided with a leading edge cooling flow channel 221 in communication with the through hole 211. The lower edge plate 23 is provided with a through hole 231 in communication with the leading edge cooling flow channel 221. An axial center line of the through hole 231 is arranged obliquely to a tangent of the turbine guide vane 2 at the through hole 231, so that the airflow flowing out of the through hole 231 has a velocity component in the same direction as the rotation direction of the first-stage turbine 1. The lower edge plate 23 is provided with a seal ring 6. The first flow guide disc 4, the seal ring 6, and the lower edge plate 23 jointly enclose a front-stage inter-stage cooling seal cavity 1a. The front-stage inter-stage cooling seal cavity 1a is in communication with a gas flow channel 2a. The second flow guide disc 5, the seal ring 6, and the lower edge plate 23 jointly enclose a rear-stage inter-stage cooling seal cavity 3a. The rear-stage inter-stage cooling seal cavity 3a is in communication with the gas flow channel 2a.

[0040] The air supply cooling system of the embodiment comprises the first-stage turbine 1 and the second-stage turbine 3, both of which comprise turbine discs and working blades. The turbine guide vane 2 and the seal ring 6 are static structures. The first-stage turbine 1, the second-stage turbine 3, the first flow guide disc 4, and the second flow guide disc 5 are rotor structures. The first flow guide disc 4 and the second flow guide disc 5 are collectively referred to as turbine discs. The cold air cools the guide vane blade 22 when flowing through the leading edge cooling flow channel 221. The cold air is forced to generate a velocity component in the rotation direction (circumferential direction) of the first-stage turbine 1 when passing through the through hole 231 and entering the front-stage inter-stage cooling seal cavity 1a, thereby reducing the difference between the outlet air circumferential velocity of the through hole 231 and the local velocity. As shown in FIGS. 7(a) and 7(b), the cold air temperature rise can be reduced by 60-100 K. The air temperature sensed by the rotor can be reduced, and the cooling efficiency can be improved. The cold air in the front-stage inter-stage cooling seal cavity 1a is divided into two parts. One part is discharged forward into the gas flow channel 2a between the first-stage turbine 1 and the guide vane blade 22. This part of the cold air completes the front-stage gas sealing and the cooling of the first flow guide disc 4. The other part is discharged backward into the flow channel between the second-stage turbine 3 and the guide vane blade 22 after sequentially entering the rear-stage inter-stage cooling seal cavity 3a and completing the rear-stage sealing and the cooling of the second flow guide disc 5. The outlet radius Rout of the through hole is relatively high (the radial height from the axial center line of the rotor structure), which is in the high-radius region of the turbine disc where the thermal environment is poor. The cold air wind resistance temperature rise in this region can be reduced, the heat exchange capacity of this region can be improved, and the temperature level of the turbine disc can be effectively reduced. Since the through hole 231 is integrated on the lower edge plate 23, the structure is simplified, and the adverse structural factors caused by the large size of the traditional high-position pre-rotation structure are avoided. The pre-rotation radius of the cold air can be increased, the cold air temperature rise can be reduced, and the potential of the cooling air can be fully utilized by combining the seal flow path design, thereby improving the air utilization efficiency and reducing the performance loss.

[0041] like Figure 5 As shown, in this embodiment, the centerline of the through-hole 231 and the tangent of the turbine guide 2 at the through-hole 231 form a pre-rotation angle α, which is 30° to 60°. This ensures that the cold air flowing out of the outlet of the through-hole 231 has sufficient circumferential velocity, thereby reducing [the cold air's temperature]. Optionally, the pre-rotation angle α can be any one of 30°, 40°, 45°, 50°, and 60°. It is understood that the number of through-holes 231 is unlimited, and the pre-rotation angle α of each through-hole 231 is the same.

[0042] like Figure 3 As shown, in this embodiment, the first guide plate 4 includes a first collar portion 402 that extends axially and is arranged opposite to the inner ring surface of the sealing ring 6, and the second guide plate 5 includes a second collar portion 502 that extends axially and is arranged opposite to the inner ring surface of the sealing ring 6. The first collar portion 402 and the second collar portion 502 overlap. The first guide plate 4, the first collar portion 402, the second collar portion 502 and the second guide plate 5 form an internal ventilation cavity 4b. The first collar portion 402 and the second collar portion 502 are respectively provided with a plurality of grates 403 that are in clearance fit with the sealing ring 6. The grates 403 and the sealing ring 6 directly form a sealing gap. Under high temperature conditions, the stress is greatest in the low radius area. If high temperature gas enters the low radius area, it may cause damage to the first guide plate 4 and the second guide plate 5. The internal ventilation cavity 4b can prevent gas from reaching the low radius area, which can effectively reduce the temperature and stress level of the wheel and improve the service life and strength reserve of the wheel. Optionally, a portion of the second collar portion 502 extends radially inward to the first collar portion 402, causing them to overlap radially to form an overlapping structure. A very small amount of airflow can flow into / out of the internal ventilation disc cavity 4b from the overlapping point of the first collar portion 402 and the second collar portion 502, allowing the low-radius area of ​​the disc to receive cool air. Optionally, the front guide plate 4 is provided with two grates 403, and the second guide plate 5 is provided with two grates 403, with the four grates 403 spaced axially. By adjusting the radial clearance between the grates 403 on the first collar portion 402 and the sealing ring 6 and / or adjusting the radial clearance between the grates 403 on the second collar portion 502 and the sealing ring 6, the distribution of cool air flow between the pre-stage intercooling sealing cavity 1a and the post-stage intercooling sealing cavity 3a can be achieved.

[0043] like Figure 3As shown, in the embodiment, the first flow guide disc 4 is provided with first protrusions 401 protruding towards the inner annular surface of the turbine guide vane 2, and the first protrusions 401 and the turbine guide vane 2 form a first gap passage 4a, and the airflow in the front inter-stage cooling sealing cavity 1a is discharged through the first gap passage 4a; the second flow guide disc 5 is provided with second protrusions 501 protruding towards the inner annular surface of the turbine guide vane 2, and the second protrusions 501 and the turbine guide vane 2 form a second gap passage 5a, and the airflow in the rear inter-stage cooling sealing cavity 3a is discharged through the second gap passage 5a; by adjusting the size of the first gap passage 4a and the second gap passage 5a, the minimum air flow rate for meeting the front and rear inter-stage sealing can be adjusted, and the risk of gas backflow can be reduced.

[0044] As shown, Figure 3 In the embodiment, the first turbine 1, the first flow guide disc 4 and the lower edge plate 23 form a front inter-stage sealing buffer cavity 1b, the front inter-stage cooling sealing cavity 1a and the front inter-stage sealing buffer cavity 1b are communicated through the first gap passage 4a, the first turbine 1 is provided with first side protrusions 101 protruding towards the lower edge plate 23, the first side protrusions 101 and the lower edge plate 23 form a first exhaust gap, and the airflow in the front inter-stage sealing buffer cavity 1b is discharged into the gas flow passage 2a through the first exhaust gap; the second flow guide disc 5, the lower edge plate 23 and the second turbine 3 form a rear inter-stage sealing buffer cavity 3b, the rear inter-stage cooling sealing cavity 3a and the rear inter-stage sealing buffer cavity 3b are communicated through the second gap passage 5a, the turbine blades of the second turbine 3 have second side protrusions 301 protruding towards the lower edge plate 23, the second side protrusions 301 and the lower edge plate 23 form a second exhaust gap, and the airflow in the rear inter-stage sealing buffer cavity 3b enters the gas flow passage 2a through the second exhaust gap; the first side protrusions 101 and the second side protrusions 301 both extend towards the lower edge plate 23, and the risk of gas backflow can be reduced.

[0045] A design method of a gas supply cooling system, which is suitable for the above-mentioned gas supply cooling system, and comprises the following steps:

[0046] S1, determining the rotor speed Ng, the number N of guide vane blades 22, the cold gas flow rate limit value m0_max, the material temperature limit T1_max of the first flow guide disc 4 and the material temperature limit T2_max of the second flow guide disc 5;

[0047] S2, determining the front inter-stage minimum sealing cold gas flow rate m1_min and the rear inter-stage minimum sealing cold gas flow rate m2_min meeting the gas sealing according to the guide vane front inter-stage rotating and static structure, the guide vane rear inter-stage rotating and static structure and the inter-stage gas parameters, wherein the rotating and static structure refers to the radial relative position and the axial relative position of the rotor and the stator, and specifically refers to the radial dimension, the radial gap and the axial overlap dimension;

[0048] S3, determining the flow m0 through the through hole 231 = m1_min + m2_min according to the flow continuity principle;

[0049] S4, determining the hole diameter D of the through hole 231;

[0050] S5, determining the pre-rotation angle a of the through hole 231;

[0051] S6, calculating the actual sealing cold gas flow m1 between the front stage and the actual sealing cold gas flow m2 between the rear stage, and determining whether m1 >= m1_min and m2 >= m2_min are satisfied;

[0052] S7, if the actual sealing cold gas flow does not satisfy the minimum sealing flow requirement, repeating S6 and S7 until m1 >= m1_min and m2 >= m2_min are satisfied;

[0053] S8, determining the temperature T1 of the first flow guide disc 4 and the temperature T2 of the second flow guide disc 5, and determining whether T1 <= T1_max and T2 <= T2_max are satisfied;

[0054] S9, if the temperature T1 of the first flow guide disc 4 and the temperature T2 of the second flow guide disc 5 do not satisfy the design requirement, increasing the flow m0 through the through hole 231, and repeating S4-S8 until T1 <= T1_max and T2 <= T2_max are satisfied;

[0055] S10, outputting the design results, i.e. the hole diameter D of the through hole 231, the pre-rotation angle a of the through hole 231, the flow m0 through the through hole 231, the actual sealing flow m1 between the front stage, the actual sealing flow m2 between the rear stage, the temperature T1 of the first flow guide disc 4 and the temperature T2 of the second flow guide disc 5.

[0056] The design method sets the through hole 231 on the lower edge plate 23, the outlet of the through hole 231 is in the disc rear high radius area (i.e. the inlet area of the front stage sealing buffer cavity 1b), on the basis of satisfying the minimum sealing flow, the pre-rotation angle a is iteratively adjusted, so that the outlet airflow of the through hole 231 has a velocity component consistent with the rotation direction of the rotor, thereby reducing the cold gas temperature rise, which can reduce the cold gas temperature rise by 60-100K, improve the cold gas quality, and finally realize the sealing cooling function with the minimum cold gas flow by gradually adjusting the cold gas flow and distribution, which can reduce the thermal load and cold air flow demand of the turbine disc rear high radius area, effectively reduce the disc temperature and stress level, and improve the disc service life, strength reserve and engine performance.

[0057] In this embodiment, step S4 specifically comprises calculating the hole diameter D of the orifice according to the total flow m0 through the orifice 231, the number N of the guide vanes 22, the number n of the orifices 231 arranged at the bottom of each guide vane 22, and the cold air parameters, wherein the cold air parameters include the total pressure before the orifice, the total temperature, and the static pressure after the orifice, wherein the total pressure before the orifice is an input parameter determined by the upstream system and does not need to be considered in this method, and the static pressure after the orifice is an automatically circulating and iterating parameter in the implementation process of this method.

[0058] In this embodiment, step S5 specifically comprises obtaining the outlet radius Rout of the orifice 231, and determining the pre-swirl angle a of the orifice according to the outlet radius Rout of the orifice 231 and the rotational speed Ng of the rotor.

[0059] In order to maximize the heat exchange intensity in the high-radius area after the disc and reduce the conduction of heat to the low position of the wheel disc, the outlet of the orifice 231 should be arranged in the high-radius area as much as possible. In order to reduce the number of parts and reduce the weight, the high-position orifice 231 is designed to be integrated with the turbine guide vane 2 in this embodiment, and the outlet radius Rout of the orifice 231 is determined accordingly. According to the rotational speed Ng of the rotor and the outlet radius Rout of the orifice 231, the rotational speed of the rotor at this radius (referred to as the local speed) can be calculated by the rotational speed calculation formula V=2*π*Ng*Rout / 60, and the unit of the obtained value is m / s;

[0060] As shown in the velocity vector relationship of the rotating and stationary system, Figure 6 As shown in the velocity vector relationship of the rotating and stationary system,

[0061] According to the cosine theorem,

[0062] w 2 = c 2 + u 2 - 2cu cos θ (1)

[0063] According to the energy equation,

[0064] The absolute total temperature of the fluid

[0065]

[0066] Wherein T is the static temperature of the air, and Cp is the specific heat capacity of the air at constant pressure.

[0067] The relative total temperature of the fluid

[0068]

[0069] As can be seen from the above formula, when θ=0°, the relative total temperature of the fluid is the smallest, and when θ=180°, the relative total temperature of the fluid is the largest.

[0070] Therefore, according to the principle of minimum difference between the outlet air circumferential velocity of the through hole 231 and the local velocity, the local velocity V is used as the initial value of the outlet air circumferential velocity of the through hole 231 to determine the pre-swirl angle α, and if the pre-swirl angle α cannot be achieved in the structure, the pre-swirl angle is gradually reduced until it can be achieved in the structure.

[0071] In this embodiment, step S6 specifically comprises: according to the inter-stage structure sealing structure parameters of the front stage of the guide vane, the sealing structure parameters of the rear stage of the guide vane, the radial gap parameters of the grid teeth 403 and the sealing ring 6, the rotor speed Ng, the hole diameter D of the through hole 231, the number n of the through hole 231 at the bottom of each guide vane 22, the pre-swirl angle α, the inter-stage gas static pressure, the inter-stage gas total temperature, the cold air inlet total pressure, and the cold air inlet total temperature, fluid dynamics calculation is carried out to determine the actual sealing cold air flow rate m1 of the front inter-stage and the actual sealing cold air flow rate m2 of the rear inter-stage, and to determine whether m1≧m1_min and m2≧m2_min are satisfied.

[0072] In this embodiment, in step S7, adjusting the radial gap of the grid teeth 403 and the sealing ring 6 specifically comprises: when m1 < m1_min, reducing the radial gap of the grid teeth 403 and the sealing ring 6; when m2 < m2_min, increasing the radial gap of the grid teeth 403 and the sealing ring 6; until m1≧m1_min and m2≧m2_min are satisfied.

[0073] In this embodiment, step S8 specifically comprises: according to M1, M2, the temperature corresponding to M1, the pressure corresponding to M1, the temperature corresponding to M2, and the pressure corresponding to M2 determined in S7, calculating the heat exchange boundary of the first and second flow guide discs 4 and 5, performing temperature analysis of the first and second flow guide discs 4 and 5, and obtaining the temperature T1 of the first flow guide disc 4 and the temperature T2 of the second flow guide disc 5.

[0074] In this embodiment, step S9 further comprises: when the flow rate m0 passing through the through hole 231 is increased to the initial cold air flow rate limit value m0_max, and the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc still do not satisfy the design requirements, the value of the cold air flow rate limit value m0_max is increased, and steps S4-S8 are repeated until T1≦T1_max and T2≦T2_max are satisfied.

[0075] The above only describes the preferred embodiments of the present application and is not intended to limit the present application. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present application shall be included in the protection scope of the present application.

Claims

1. A method for designing a gas cooling system, characterized by, The gas supply cooling system comprises a first stage turbine (1), a turbine guide vane (2) and a second stage turbine (3) arranged in sequence along an axial direction, and further comprises a first flow guide disc (4) and a second flow guide disc (5) arranged in sequence along the axial direction and located radially inside the turbine guide vane (2); The turbine guide vane (2) comprises an upper edge plate (21), a guide vane blade (22) and a lower edge plate (23), the upper edge plate (21) is provided with a through hole (211) for introducing cold gas, the guide vane blade (22) is provided with a front edge cooling flow channel (221) in communication with the through hole (211), and the lower edge plate (23) is provided with a through hole (231) in communication with the front edge cooling flow channel (221); the axial center line of the through hole (231) is arranged obliquely to the tangent of the turbine guide vane (2) at the through hole (231), so that the airflow flowing out of the through hole (231) has a velocity component in the same direction as the rotation direction of the first stage turbine (1); the axial center line of the through hole (231) and the tangent of the turbine guide vane (2) at the through hole (231) form a pre-rotation angle α; The lower edge plate (23) is provided with a sealing ring (6), the first flow guide disc (4), the sealing ring (6) and the lower edge plate (23) jointly enclose a front stage inter-stage cooling sealing cavity (1a), the front stage inter-stage cooling sealing cavity (1a) is in communication with a gas flow channel (2a), the second flow guide disc (5), the sealing ring (6) and the lower edge plate (23) jointly enclose a rear stage inter-stage cooling sealing cavity (3a), and the rear stage inter-stage cooling sealing cavity (3a) is in communication with the gas flow channel (2a); The design method of the gas supply cooling system comprises the following steps: S1, determining the rotor speed Ng, the number of guide vane blades N, the cold gas flow limit value m0_max, the material temperature limit T1_max of the first flow guide disc and the material temperature limit T2_max of the second flow guide disc; S2, determining the minimum sealing cold gas flow m1_min of the front stage inter-stage and the minimum sealing cold gas flow m2_min of the rear stage inter-stage according to the guide vane front stage inter-stage rotating structure, the guide vane rear stage inter-stage rotating structure and the inter-stage gas parameters; S3, determining the flow m0 passing through the through hole according to the flow continuity principle; S4, determining the hole diameter D of the through hole; S5, determining the pre-rotation angle α of the through hole; S6, calculating the actual sealing cold gas flow m1 of the front stage inter-stage and the actual sealing cold gas flow m2 of the rear stage inter-stage, and determining whether m1≧m1_min and m2≧m2_min are satisfied; S7, if the actual sealing cold gas flow does not satisfy the minimum sealing flow requirement, then S6 and S7 are repeated until m1≧m1_min and m2≧m2_min are satisfied; S8, determining the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc, and determining whether T1≦T1_max and T2≦T2_max are satisfied; S9, if the temperature T1 of the first flow guide disc and the temperature T2 of the second flow guide disc do not satisfy the design requirement, then the flow m0 passing through the through hole is increased, and S4-S8 are repeated until T1≦T1_max and T2≦T2_max are satisfied. S10, output the design result, i.e. the aperture D of the through hole, the pre-swirl angle α of the through hole, the flow m0 through the through hole, the actual sealing flow m1 between the front stages, the actual sealing flow m2 between the rear stages, the temperature T1 of the first flow guiding disc and the temperature T2 of the second flow guiding disc.

2. The method of designing a gas supply cooling system according to claim 1, wherein, The pre-swirl angle α is 30-60°.

3. The method of designing a gas supply cooling system according to claim 2, wherein, The first flow guiding disc (4) comprises a first shaft ring part (402) extending axially and arranged opposite the inner ring surface of the sealing ring (6), and the second flow guiding disc (5) comprises a second shaft ring part (502) extending axially and arranged opposite the inner ring surface of the sealing ring (6), and the first shaft ring part (402) and the second shaft ring part (502) are overlapped; The first flow guiding disc (4), the first shaft ring part (402), the second shaft ring part (502) and the second flow guiding disc (5) form an internal flow guiding disc cavity (4b), and a plurality of grid teeth (403) are arranged on the first shaft ring part (402) and the second shaft ring part (502) respectively and gap-fit with the sealing ring (6).

4. The method of designing a gas supply cooling system according to claim 1, wherein, Step S4 specifically comprises calculating the aperture D of the through hole according to the total flow m0 through the through hole, the number N of the guide vane blades, the number n of the through holes arranged at the bottom of each guide vane blade and the cold air parameters.

5. The method of designing a gas supply cooling system according to claim 4, wherein, Step S5 specifically comprises obtaining the outlet radius Rout of the through hole, and determining the pre-swirl angle α of the through hole according to the outlet radius Rout of the through hole and the rotor speed Ng.

6. The method of designing a gas supply cooling system according to claim 5, wherein, Step S6 specifically comprises carrying out fluid dynamics calculation according to the sealing structure parameters of the guide vane front stage inter-stage structure, the sealing structure parameters of the guide vane rear stage inter-stage structure, the radial gap parameters of the grid teeth and the sealing ring, the rotor speed Ng, the aperture D of the through hole, the number n of the through holes at the bottom of each guide vane blade, the pre-swirl angle α, the inter-stage gas static pressure, the inter-stage gas total temperature, the cold air inlet total pressure, the cold air inlet total temperature, determining the actual sealing cold air flow m1 between the front stages and the actual sealing cold air flow m2 between the rear stages, and judging whether m1≧m1_min and m2≧m2_min are satisfied.

7. The method of designing a gas supply cooling system according to claim 6, wherein, In step S7, adjusting the radial gap of the grid teeth and the sealing ring specifically comprises: when m1 8. The method of designing a gas supply cooling system according to claim 7, wherein, Step S8 specifically comprises calculating the heat exchange boundary of the first flow guiding disc and the second flow guiding disc according to m1, m2, the temperature corresponding to m1, the pressure corresponding to m1, the temperature corresponding to m2 and the pressure corresponding to m2 determined in S7, carrying out temperature analysis of the first flow guiding disc and the second flow guiding disc, and obtaining the temperature T1 of the first flow guiding disc and the temperature T2 of the second flow guiding disc.

9. The method of designing a gas supply cooling system according to claim 1, wherein, Step S9 further comprises: when the flow m0 through the through hole is increased to the initial cold air flow limit value m0_max, and the temperature T1 of the first flow guiding disc and the temperature T2 of the second flow guiding disc still do not satisfy the design requirements, increasing the value of the cold air flow limit value m0_max, and repeating S4-S8 until T1≦T1_max and T2≦T2_max are satisfied.

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