WO2024017385A1 - 高压涡轮的动叶片 - Google Patents

高压涡轮的动叶片 Download PDF

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Publication number
WO2024017385A1
WO2024017385A1 PCT/CN2023/108842 CN2023108842W WO2024017385A1 WO 2024017385 A1 WO2024017385 A1 WO 2024017385A1 CN 2023108842 W CN2023108842 W CN 2023108842W WO 2024017385 A1 WO2024017385 A1 WO 2024017385A1
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WO
WIPO (PCT)
Prior art keywords
impact
hole
cooling
wall
pressure turbine
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Application number
PCT/CN2023/108842
Other languages
English (en)
French (fr)
Inventor
尤煜龙
丁亮
谭智勇
侯乃先
Original Assignee
中国航发商用航空发动机有限责任公司
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Application filed by 中国航发商用航空发动机有限责任公司 filed Critical 中国航发商用航空发动机有限责任公司
Publication of WO2024017385A1 publication Critical patent/WO2024017385A1/zh

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to high-pressure turbine blades of aerospace turbine engines, and in particular to a moving blade of a high-pressure turbine.
  • the cooling technologies of modern aeroengine high-pressure turbine blades mainly include impact, turbulence, and film cooling.
  • the outer wall of the blade is provided with The air film holes inhibit the convective heat transfer between the outer wall of the turbine and the high-temperature gas.
  • the double-wall cooling structure is considered to have higher cooling efficiency.
  • the double-wall cooling structure is generally a composite cooling unit composed of the outer wall and the inner wall of the turbine blade, with impact holes, air film holes and turbulence structures.
  • the cold air inside the turbine blades impacts and cools the outer wall through the impact holes on the inner wall, then flows through the spoiler structure to further enhance heat transfer, and finally flows out from the air film holes on the outer wall to cool and protect the outer wall. Therefore, the double-wall cooling structure has gradually attracted the attention of technicians and been implemented in the cooling design of high-pressure turbine blades.
  • turbine blade cooling design is a complex engineering exercise.
  • the leading edge of the turbine blade and its surrounding area are also high heat load areas.
  • the gas temperature in the leading edge and pressure side area is high, which is generally the focus of cooling design.
  • the high-speed flow of gas on the suction side causes this location to have a high exchange rate.
  • the thermal coefficient also worsens the operating environment of the moving blades.
  • the static pressure between the leading edge of the moving blade and the pressure surface is relatively high, which requires the cold air to require a higher total pressure to ensure stable outflow and the necessary counterflow margin; while the static pressure on the suction surface is relatively low, and the static pressure flows from the leading edge toward the suction
  • the distribution differences in face directions are usually larger.
  • the double-wall cooling design of turbine blades is more common in high-pressure turbine guide vanes, and the technical solution is mainly to arrange an independent large-sized cold air chamber to supply air to the double-wall channel.
  • This type of structure is not conducive to the control of cold air flow in different areas inside and outside the blade body.
  • the moving blades of the turbine are rotating parts, the requirements for structural strength are more stringent, and the size is smaller than the guide vanes. Therefore, a double-wall design plan is to arrange an independent large-sized cold air chamber in the guide vanes. Not suitable for turbine blades.
  • the technical problem to be solved by the present invention is to provide a high-pressure turbine moving blade in order to overcome the defects in the prior art that the structural design of the high-pressure turbine moving blade of an aerospace engine is unreasonable, resulting in poor cooling effect of the moving blade.
  • a moving blade of a high-pressure turbine The moving blade has an outer wall, an inner wall and an inner chamber.
  • a cooling unit is provided inside the moving blade.
  • the cooling unit has a cooling chamber.
  • the cooling chamber Communicated with the inner chamber through an impingement hole, the cooling unit includes at least one series double-wall cooling unit, the series double-wall cooling unit includes a cooling unit formed between the outer wall and the inner wall.
  • a plurality of the cooling chambers are connected in sequence, the impact hole includes a first impact hole, and the cooling chamber at the head end is connected to the inner chamber through the first impact hole,
  • the cooling chamber at the rear end is connected to the outside through an air film hole or a trailing edge slit.
  • the moving blades of the high-pressure turbine are equipped with cooling units inside the moving blades.
  • the cooling units are used to cool different internal areas of the moving blades, and the cooling units include at least one series double-wall cooling unit.
  • the cold air flows through multiple cooling chambers in the moving blades in sequence, increasing the flow path of the cold air to carry out multiple continuous impact jets in the moving blades to achieve local enhanced heat transfer inside the moving blades.
  • there is no need to increase the amount of additional cold air so that as much cold air as possible can be supplied to the areas with higher temperature of the moving blades for cooling. This enables the regulation of the amount of cold air and effectively reduces the heat load of the moving blades of the high-pressure turbine.
  • the interior of the rotor blade is divided into at least three inner chambers along the axial direction by diaphragms.
  • cold air enters different inner chambers from the cold air channel of the tenon.
  • At least one inner chamber is used to ensure cooling of the leading edge position of the moving blade, and at least one inner chamber is used to ensure cooling of the trailing edge position of the moving blade.
  • At least An inner chamber is used to ensure cooling of the area between the leading and trailing edges of the rotor blades.
  • the series double-wall cooling unit is arranged close to the suction surface of the moving blade.
  • the high-speed flow of gas on the suction surface of the moving blade results in a high heat transfer coefficient at this location, which deteriorates the operating environment of the moving blade.
  • the static pressure between the blade leading edge and the pressure surface is relatively high, which requires the cold air to require a higher total pressure to ensure stable outflow and the necessary counterflow margin; while the static pressure on the suction surface is relatively low, and the static pressure is from the leading edge to the suction surface.
  • the distribution difference in direction is usually large.
  • a series double-wall cooling unit is installed on the suction surface side. The cold air flows out from the air film holes after multiple impact flows, which still ensures a certain total pressure and counterflow margin. It can not only ensure enhanced heat exchange inside the moving blades, but also be used to regulate the amount of cold air in different areas, thereby effectively reducing the heat load on the high-pressure turbine moving blades.
  • the series double wall cooling unit is arranged around the inner chamber in a circumferential direction.
  • the inner chamber can also cool the gas in the series double-wall cooling units located on the peripheral side, reducing the temperature rise of the cold air in the series double-wall cooling units caused by multiple impact heat exchanges.
  • the ratio of the channel width of the cooling chamber located on the suction surface to the aperture of the first impact hole ranges from 2:1 to 6:1;
  • the ratio of the impact distance of the first impact hole to the aperture diameter of the first impact hole ranges from 0.5:1 to 2:1.
  • the air film hole includes a first air film hole and a second air film hole, and the cooling chamber at the head end is connected to the outside through the first air film hole on the outer wall;
  • the cooling chamber at the rear end is connected to the outside through the second air film hole on the outer wall.
  • the cold air entering the inner chamber enters the cooling chamber through the impact hole, and then part of the cold air flows out through the first air film hole, and the other part of the cold air flows through the cooling chamber in sequence, and then flows out from the second air film hole.
  • the moving blades are impacted and cooled.
  • the number of the cooling chambers is two, and a return channel is provided between the two cooling chambers.
  • the cooling chamber at the head end is connected to the return channel through an axial impact hole, and the cooling chamber at the tail end is connected to the return channel through an axial impact hole.
  • the cooling chamber at the end passes through the fifth impact hole and the return channel.
  • the diameter of the axial impact hole does not exceed the impact distance of the first impact hole.
  • the series double-wall cooling unit is disposed close to the trailing edge of the rotor blade.
  • a series double-wall cooling unit is installed at the trailing edge of the moving blade, which not only ensures enhanced heat transfer of the moving blade, but also saves the amount of cold air at the trailing edge, thereby effectively reducing the heat of the moving blade of the high-pressure turbine. load. Since the static pressure on the surface of the moving blades of the high-pressure turbine decreases from the middle position of the pressure surface toward the trailing edge, the cold air flows through multiple impact jets and flows out from the trailing edge split, ensuring a certain total pressure and counterflow margin.
  • the air film hole further includes a third air film hole, the cooling chamber at the head end is connected to the outside through the third air film hole, and the cooling chamber at the tail end is connected through the trailing edge slit. Connect with the outside;
  • the impact hole includes a second impact hole, and the cooling chamber at the head end communicates with the inner chamber through the second impact hole.
  • the cold air entering the inner chamber enters the cooling chamber through the second impact hole, and then part of the cold air flows out through the third air film hole, and the other part of the cold air flows through the cooling chamber in sequence, and then flows out from the trailing edge split.
  • the cold air impacts and cools the moving blades.
  • the ratio of the impact distance of the second impact hole to the aperture diameter of the second impact hole ranges from 0.5:1 to 2:1.
  • the cooling unit further includes a leading edge double-wall cooling unit, the leading edge double-wall cooling unit is disposed close to the leading edge of the moving blade, the impact hole includes a third impact hole, and the front edge double-wall cooling unit is disposed close to the leading edge of the moving blade.
  • An edge double wall cooling unit communicates with the inner chamber through the third impingement hole.
  • a leading edge double-wall cooling unit is installed at the leading edge to reduce the heat load at the leading edge of the moving blade.
  • the ratio of the impact distance of the third impact hole to the aperture diameter of the third impact hole ranges from 1:1 to 3:1;
  • the impact distance is the distance between the third impact hole and the outer wall.
  • the cooling unit further includes a pressure surface double-wall cooling unit, the pressure surface double-wall cooling unit is disposed close to the pressure surface of the moving blade, and the impact hole further includes a fourth impact hole.
  • a pressure side double wall cooling unit communicates with the inner chamber through the fourth impingement hole. on our side In this case, the gas temperature in the leading edge and pressure side areas is high. Therefore, a pressure side double-wall cooling unit is installed at the pressure side to reduce the heat load at the pressure side of the moving blade.
  • the ratio of the channel width of the cooling chamber located on the pressure surface to the aperture of the fourth impact hole ranges from 2:1 to 4:1;
  • the ratio of the impact distance of the fourth impact hole to the aperture diameter of the fourth impact hole ranges from 0.5:1 to 2:1.
  • the number of the impact holes is determined according to the volume of the inner chamber, the volume of the cooling chamber and the flow rate of cold air filled into the inner chamber.
  • the diameter of the impact hole is 0.8mm-1.5mm.
  • the number and rows of the air film holes are determined according to the channel width of the cooling chamber and the flow rate of cold air flowing into the cooling chamber.
  • the ratio of the pore diameter of the air film hole to the pore diameter of the impact hole ranges from 0.4:1 to 0.6:1.
  • a plurality of spoiler columns are arranged at intervals in the cooling chamber.
  • spoiler columns are set up to increase the flow path of cold air and improve the cooling effect.
  • the ratio of the diameter of the spoiler column to the diameter of the impact hole ranges from 0.8:1 to 1.2:1.
  • ribs are also provided inside the inner chamber.
  • the positive and progressive effect of the present invention is that the moving blade of the high-pressure turbine is provided with a cooling unit inside the moving blade.
  • the cooling unit is used to cool different internal areas of the moving blade, and the cooling unit includes at least one serial double-wall cooling unit.
  • the cold air flows through multiple cooling chambers in the moving blade in sequence, increasing the cold air flow path to carry out multiple continuous impact jets in the moving blade to achieve local strengthening of the interior of the moving blade. It can transfer heat without increasing the amount of cold air, so that as much cold air as possible can be supplied to the areas with higher temperature of the moving blades for cooling. This enables the regulation of the amount of cold air and effectively reduces the heat load of the moving blades of the high-pressure turbine.
  • Figure 1 is a schematic three-dimensional structural diagram of the moving blades of the high-pressure turbine according to a preferred embodiment 1 of the present invention.
  • Figure 2 is a cross-sectional view along line A-A in Figure 1.
  • Figure 3 is a schematic three-dimensional structural diagram of a double-layer wall cooling unit connected in series on the suction surface of the rotor blade of the high-pressure turbine according to a preferred embodiment 1 of the present invention.
  • Figure 4 is a schematic three-dimensional structural diagram of a suction surface series double-layer wall cooling unit equipped with spoiler columns according to a preferred embodiment 1 of the present invention.
  • Figure 5 is a schematic three-dimensional structural diagram of a trailing edge series double-wall cooling unit equipped with a spoiler column according to a preferred embodiment 1 of the present invention.
  • Figure 6 is a schematic cross-sectional structural view of the moving blades of the high-pressure turbine according to a preferred embodiment 2 of the present invention.
  • this embodiment discloses a moving blade of a high-pressure turbine.
  • the moving blade 1 has an edge plate 2, a tenon 3, a leading edge 111, a trailing edge 115, and between the leading edge 111 and the trailing edge 115.
  • the trailing edge 115 is between the pressure surface 112 and the suction surface 113 .
  • the rotor blade 1 also has an outer wall 11 , an inner wall 12 and an inner chamber 14 .
  • a cooling unit is provided inside the rotor blade 1.
  • the cooling unit has a cooling chamber 100, and the cooling chamber 100 communicates with the inner chamber through an impact hole.
  • the cooling unit includes at least one series double-wall cooling unit.
  • the series double-wall cooling unit includes a plurality of cooling chambers 100 formed between the outer wall 11 and the inner wall 12.
  • the plurality of cooling chambers 100 are connected in sequence, and the impact holes are Including a first impact hole 121, the cooling chamber 100 at the head end communicates with the inner chamber through the first impact hole 121, and the cooling chamber 100 at the tail end communicates with the outside through the air film hole or the trailing edge split 114.
  • the cold air inside the turbine blade impacts and cools the outer wall 11 through the impact holes on the inner wall 12, and finally flows out from the air film holes on the outer wall 11 to cool and protect the outer wall surface.
  • the cooling unit is used to cool down different areas inside the rotor blade 1, and includes at least one series-connected double-wall cooling unit.
  • cooling air With a certain amount of cooling air, the cold air flows through multiple areas in the rotor blade 1 in sequence.
  • a cooling chamber 100 increases the flow path of the cold air to carry out multiple continuous impact jets in the moving blade 1 to achieve local enhanced heat transfer inside the moving blade 1.
  • the amount of cold air is controlled and the thermal load of the moving blades 1 of the high-pressure turbine is effectively reduced.
  • the interior of the rotor blade 1 is divided into five inner cavities 14 along the axial direction by the diaphragm 13 , and the five inner cavities 14 are the first inner cavity in order from the leading edge 111 to the trailing edge 115 .
  • the IVs are connected via a rotary channel.
  • the five inner chambers 14 are divided into three air inlet channels. Among them, the first inner chamber I and the fifth inner chamber V are separate air inlet channels respectively, and the second inner chamber II and the third inner chamber V are separate air inlet channels.
  • Chamber III and the fourth inner chamber IV are connected through a rotary channel to form an air inlet channel.
  • the cold air entering from the cold air passage of the tenon 3 is divided into three paths and enters different inner cavities in the moving blade 1 .
  • the interior of the rotor blade is divided into at least three internal chambers in the axial direction by diaphragms.
  • the cold air enters different inner chambers from the cold air channel of the tenon 3.
  • At least one inner chamber is used to ensure cooling of the leading edge position of the moving blade.
  • At least one inner chamber is used to ensure cooling of the trailing edge position of the moving blade.
  • At least one inner chamber is used to ensure cooling of the trailing edge position of the moving blade. Used to ensure cooling of the area between the leading edge and trailing edge of the moving blade.
  • the series double-wall cooling unit is disposed close to the suction surface 113 of the rotor blade 1 , which is also called the suction surface series double-wall cooling unit 152 . Due to the high-speed flow of gas on the suction surface 113 of the moving blade 1, this location has a high heat transfer coefficient, which deteriorates the operating environment of the moving blade 1.
  • the static pressure between the blade leading edge and the pressure surface is relatively high, which requires a higher total pressure of the cold air to ensure a stable outflow and a necessary counterflow margin; while the static pressure on the suction surface 113 is relatively low, and the static pressure flows from the leading edge 111 toward The distribution difference in the direction of suction surface 113 is usually large.
  • a series double-wall cooling unit is installed on the suction surface 113 side.
  • the cold air flows out from the air film holes after multiple impact flows, and a certain total pressure and counterflow margin can still be ensured. . It can not only ensure enhanced heat exchange inside the moving blade 1, but also be used to regulate the amount of cold air in different areas, thereby effectively reducing the heat load of the high-pressure turbine moving blade.
  • the series double-wall cooling unit is arranged around the inner chamber in the circumferential direction, so that the cooling gas in the inner chamber can also cool the gas flowing in the series double-wall cooling unit.
  • the suction surface serial double-wall cooling unit 152 is arranged around the first inner chamber I and the second inner chamber II in the circumferential direction.
  • the cooling chamber 100 at the head end communicates with the return channel 153 through the axial impact hole 126
  • the cooling chamber 100 at the rear end communicates with the return channel 153 through the fifth impact hole 125 .
  • the return channel 153 connects the two adjacent cooling chambers 100, and the second inner chamber II can also cool the gas in the return channel 153 located on the peripheral side, reducing the amount of cold air in the series double-wall cooling unit.
  • the setting parameters of the fifth impact hole are the same as the setting parameters of the first impact hole.
  • the air film holes include a first air film hole 1101 and a second air film hole 1102.
  • the cooling chamber 100 at the head end passes through the third air film hole on the outer wall 11.
  • An air film hole 1101 is connected to the outside.
  • the cooling chamber 100 at the rear end is connected to the outside through the second air film hole 1102 on the outer wall 11 .
  • the cold air entering the first inner chamber 1 enters the cooling chamber 100 at the head end through the first impact hole 121 and cools the chamber. Then a part of the cold air flows out through the first air film hole 1101, and the other part of the cold air passes through the axial impact hole.
  • the ratio of the channel width of the cooling chamber 100 located on the suction surface to the aperture of the first impact hole 121 ranges from 2:1 to 6:1.
  • a series double-wall cooling unit is also provided at the trailing edge 115 of the moving blade 1 , also known as the trailing edge series double-wall cooling unit 155 .
  • the trailing edge series double-wall cooling unit 155 not only ensures enhanced heat exchange at the trailing edge 115 of the moving blade 1, but also saves the amount of cooling air at the trailing edge 115 position, thereby effectively reducing the heat load of the moving blade 1 of the high-pressure turbine. Since the static pressure on the surface of the moving blade 1 of the high-pressure turbine decreases from the middle position of the pressure surface 112 toward the trailing edge 115, the cold air flows out from the trailing edge split 114 through multiple impact jets, ensuring a certain total pressure and Counterflow margin.
  • the ratio of the impact distance of the first impact hole 121 to the hole diameter of the first impact hole 121 ranges from 0.5:1 to 2:1.
  • the diameter of the axial impact hole 126 does not exceed the impact distance of the first impact hole 121 .
  • the impact distance of the first impact hole 121 is the distance between the first impact hole 121 and the outer wall 11 .
  • the air film hole also includes a third air film hole 1103 , and the cooling chamber 100 at the head end is connected to the outside through the third air film hole 1103 .
  • the cooling chamber 100 at the rear end is connected to the outside through the trailing edge split 114 .
  • the impact hole includes a second impact hole 124 , and the cooling chamber 100 at the head end communicates with the inner chamber V through a second impact hole 124 .
  • the cooling chambers 100 at the head end and the tail end are connected through another second impact hole 124 .
  • the cold air entering the fifth inner chamber V enters the cooling chamber 100 at the head end through a second impact hole 124, and then a part of the cold air flows out through the third air film hole 1103, and the other part of the cold air passes through another second impact hole 1103.
  • the impact hole 124 flows into the next cooling chamber 100 and cools the chamber, and then flows out from the trailing edge split slit 114. During the flow, the cold air impacts and cools the rotor blade 1.
  • the ratio of the impact distance of the second impact hole to the aperture diameter of the second impact hole ranges from 0.5:1 to 2:1.
  • the impact distance of the second impact hole is the distance between the second impact hole and the outer wall 11 .
  • the cooling unit also includes a leading edge double-wall cooling unit 151.
  • the leading edge double-wall cooling unit 151 is provided close to the leading edge of the moving blade 1.
  • the impact holes include a third impact hole 122.
  • the layer wall cooling unit 151 communicates with the first inner chamber 1 through the third impact hole 122 . Since the gas temperature in the leading edge 111 and the pressure side area is high, a leading edge double wall cooling unit 151 is provided at the leading edge 111 position to reduce the heat load at the leading edge 111 position of the moving blade 1 .
  • the ratio of the impact distance of the third impact hole 122 to the hole diameter of the third impact hole 122 ranges from 1:1 to 3:1.
  • the impact distance is the distance between the third impact hole 122 and the outer wall 11 .
  • the cooling unit also includes a pressure surface double-wall cooling unit 154.
  • the pressure surface double-wall cooling unit 154 is provided close to the pressure surface of the moving blade 1.
  • the impact hole also includes a fourth impact hole 123.
  • the pressure surface double-wall cooling unit 154 is disposed close to the pressure surface of the moving blade 1. 154 communicates with the inner chamber through the fourth impact hole 123. Since the gas temperature in the leading edge 111 and the pressure side area is high, a pressure surface double-wall cooling unit 154 is provided at the pressure surface 112 to reduce the heat load at the pressure surface 112 of the rotor blade 1 .
  • a corresponding setting reduces the heat load at the pressure surface 112 of the moving blade 1.
  • the ratio of the channel width of the cooling chamber 100 located on the pressure surface to the hole diameter of the fourth impact hole 123 ranges from 2:1 to 4:1.
  • the number of impact holes is determined according to the volume of the inner chamber, the volume of the cooling chamber 100 and the flow rate of cold air filled into the inner chamber.
  • the ratio of the impact distance of the fourth impact hole 123 to the hole diameter of the fourth impact hole 123 ranges from 0.5:1 to 2:1.
  • the impact distance of the fourth impact hole 123 is the distance between the fourth impact hole 123 and the outer wall 11 .
  • the hole diameter of the impact hole is 0.8mm-1.5mm.
  • the suction surface serial double-wall cooling unit 152 communicates with the first inner chamber 1 through a row of first impact holes 121 .
  • the number of first impact holes 121 in a single row is 8-12.
  • 1-3 rows of third impact holes 122 are arranged between the leading edge double-wall cooling unit 151 and the first inner chamber 1.
  • the number of third impact holes 122 in a single row is 8-12.
  • the three pressure surface double-wall cooling units 154 are respectively connected to the second inner chamber II, the third inner chamber III and the fourth inner chamber IV through a row of fourth impact holes 123 .
  • Solo fourth rush The number of punch holes 123 is 8-12.
  • the trailing edge series double-wall cooling unit 155 communicates with the fifth inner chamber V through a row of second impact holes 124 .
  • the number of second impact holes 124 in a single row is 8-12.
  • the number and rows of air film holes are determined according to the channel width of the cooling chamber 100 and the flow rate of cold air flowing into the cooling chamber 100 .
  • the ratio of the aperture of the air film hole to the aperture of the impact hole ranges from 0.4:1 to 0.6:1.
  • the outer wall 11 at the leading edge 111 position is provided with 4-8 exhaust film holes, and the number of single exhaust film holes is 14-20.
  • one exhaust film hole is arranged in the pressure surface double-wall cooling unit 154, and the number of single exhaust film holes is 14-20.
  • 1-2 exhaust film holes are arranged at the pressure surface 112 of the series-connected double-wall cooling unit 155 at the trailing edge, and the number of single exhaust film holes is 14-20.
  • a plurality of spoiler columns 16 are spaced in the cooling chamber 100 to increase the flow path of the cold air and improve the cooling effect.
  • the ratio of the diameter of the spoiler column 16 to the diameter of the impact hole ranges from 0.8:1 to 1.2:1.
  • the spoiler pillars 16 and the fourth impact holes 123 are alternately arranged in a row along the radial direction.
  • the first impact hole 121 is arranged upstream, and the spoiler column 16 is arranged downstream.
  • the number of rows of impact holes and spoiler columns 16 in the double-wall channel can be increased.
  • Ribs are also provided inside the inner chamber to enhance internal turbulence and improve convection heat transfer inside the flow channel of the inner chamber.
  • the ratio of the height of the ribs to the height of the inner chamber channel ranges from 1:5 to 1:10.
  • the gas flow process inside the moving blade 1 is as follows.
  • the cold air flow direction 200 in the moving blade 1 is shown by the arrow in Figure 2.
  • the first cold air enters the first inner chamber I from the cold air channel inlet of the tenon 3, and a part of the cold air enters the leading edge double wall unit through the third impact hole 122. 151, and flows out from the air film hole at the front edge 111, and another part of the cold air enters the suction surface series double-wall cooling unit 152 through the first impact hole 121, and repeats 2 times of impact-air film cooling and 1 axial impact cooling, They flow out from the first air film hole 1101 and the second air film hole 1102 respectively.
  • the second cold air first enters the second inner chamber II.
  • a part of the cold air directly enters the pressure surface double-wall unit 154 through the fourth impact hole 123 and flows out from the air film hole.
  • the other part of the cold air flows through the third inner chamber sequentially through the rotation channel.
  • Chamber III and the fourth inner chamber IV implement impact-air film cooling at the same time.
  • the cold air flowing through the second inner chamber II also plays a cooling role in the cold air in the series double-wall cooling unit 152 on the suction surface.
  • the third cold air enters the blade from the fifth inner chamber V, and enters the trailing edge series double-wall cooling unit 155 through a second impact hole 124.
  • Part of the cold air flows out through the third air film hole 1103, and the other part of the cold air flows downstream.
  • the flow enters the cooling chamber 100 from another second impact hole 124 and impacts the spoiler column 16 , and finally flows out from the trailing edge split slit 114 .
  • this embodiment is basically the same as Embodiment 1.
  • the difference is that three cooling chambers are provided in the suction surface series double-wall cooling unit, thereby achieving three times of impact-air film cooling and 2Twice axial impact cooling to improve the cooling effect of the moving blades.
  • the flow direction of the cold air is shown by the arrow in Figure 6.
  • the number of cooling chambers can be selected and set to achieve the dual effects of saving air and enhancing heat exchange inside the rotor blades of the high-pressure turbine.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
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Abstract

一种高压涡轮的动叶片，其内部设有冷却单元，冷却单元具有冷却腔室，冷却腔室通过冲击孔与内腔室连通，冷却单元包括至少一个串联双层壁冷却单元，串联双层壁冷却单元包括在外层壁和内层壁之间形成的多个冷却腔室，多个冷却腔室依次连通，冲击孔包括第一冲击孔，首端的冷却腔室通过第一冲击孔与内腔室连通，尾端的冷却腔室通过气膜孔或尾缘劈缝与外部连通。该高压涡轮的动叶片在其内部设置有冷却单元，冷却单元用于对动叶片的内部不同区域进行降温，并且在冷却单元中包括至少一个串联双层壁冷却单元，冷气依次流过动叶片内的多个冷却腔室，增加冷气流动路径，以在动叶片内进行多次连续冲击射流，降低高压涡轮的动叶片的热负荷。

Description

高压涡轮的动叶片 技术领域
本发明涉及航空涡轮发动机高压涡轮叶片，特别涉及一种高压涡轮的动叶片。
背景技术
随着现代航空发动机技术的发展，涡轮进口温度不断升高。目前的航空发动机高压涡轮进口温度已接近2000K左右，未来更为先进的航空发动机涡轮进口温度将达到2100K-2200K水平，这对于高压涡轮的冷却设计提出了极大挑战。现代航空发动机高压涡轮叶片的冷却技术主要有冲击、扰流以及气膜冷却等方式。设计人员为了提高高压涡轮叶片冷效，通常采用复合冷却结构设计，尽可能增加涡轮叶片内部结构的换热面积，并布置冲击、扰流等冷却形式，强化内部传热，同时在叶片外壁面设置气膜孔，抑制涡轮外壁面与高温燃气的对流换热。
在各类涡轮叶片冷却结构设计方案中，双层壁冷却结构被认为具有较高的冷却效率。双层壁冷却结构一般是由涡轮叶片的外层壁和内层壁组合构成、具有冲击孔、气膜孔和扰流结构的复合冷却单元。涡轮叶片内部的冷气通过内层壁上的冲击孔对外层壁进行冲击冷却，随后流经扰流结构进一步强化换热，最后从外层壁上的气膜孔流出，对外壁面实施冷却保护。因此，双层壁冷却结构逐渐得到技术人员的关注并被实施于高压涡轮叶片的冷却设计中。
实际上，涡轮的动叶片冷却设计是一项复杂的工程实践。涡轮的动叶片前缘及其周围区域也是高热负荷区，前缘和压力侧区域的燃气温度高，一般都是冷却设计重点关注的位置；吸力侧燃气的高速流动导致该位置具有较高的换热系数，也使得动叶片运行环境恶化。动叶片前缘与压力面的静压较高，这要求冷气需要较高的总压才能保证稳定出流和必要的逆流裕度；而吸力面的静压相对较低，静压从前缘朝吸力面方向的分布差异通常较大。现阶段， 涡轮叶片的双层壁冷却设计在高压涡轮导叶中更为常见，且主要是采用布置独立大尺寸的冷气腔室为双层壁通道供气的技术方案。这类结构形式不利于叶身内外部不同区域的冷气流量调控。此外，由于涡轮的动叶片属于旋转部件，对结构强度的要求更为苛刻，且尺寸相较于导叶更小，因此，在导叶中布置独立大尺寸的冷气腔室的双层壁设计方案不适用于涡轮的动叶片。
发明内容
本发明要解决的技术问题是为了克服现有技术中航空发动机高压涡轮动叶的结构设计不合理，导致动叶的冷却效果较差的缺陷，提供一种高压涡轮的动叶片。
本发明是通过下述技术方案来解决上述技术问题：
一种高压涡轮的动叶片，所述动叶片具有外层壁、内层壁和内腔室，所述动叶片的内部设有冷却单元，所述冷却单元具有冷却腔室，所述冷却腔室通过冲击孔与所述内腔室连通，所述冷却单元包括至少一个串联双层壁冷却单元，所述串联双层壁冷却单元包括在所述外层壁和所述内层壁之间形成的多个所述冷却腔室，多个所述冷却腔室依次连通，所述冲击孔包括第一冲击孔，首端的所述冷却腔室通过所述第一冲击孔与所述内腔室连通，尾端的所述冷却腔室通过气膜孔或尾缘劈缝与外部连通。
在本方案中，该高压涡轮的动叶片在其内部设置有冷却单元，冷却单元用于对动叶片的内部不同区域进行降温，并且在冷却单元中包括至少一个串联双层壁冷却单元，在一定量的冷却气量的情况下，冷气依次流过动叶片内的多个冷却腔室，增加冷气流动的路径，以在动叶片内进行多次连续冲击射流，实现动叶片内部的局部强化传热，同时也无需额外增加冷气用量，以便将冷气尽量多的供给动叶片温度较高的区域进行冷却，由此实现了冷气用量的调控，有效降低高压涡轮的动叶片的热负荷。
较佳地，所述动叶片的内部被横隔板沿轴向方向分隔为至少三个所述内腔室。
在本方案中，冷气从榫头的冷气通道进入不同内腔室，至少一个内腔室用于保障动叶片的前缘位置冷却，至少一个内腔室用于保障动叶片的尾缘位置冷却，至少一个内腔室用于保障动叶片的前缘和尾缘之间的区域冷却。
较佳地，所述串联双层壁冷却单元靠近所述动叶片的吸力面设置。
在本方案中，在动叶片的吸力面由于燃气高速流动，导致该位置具有较高的换热系数，使得动叶片运行环境恶化。叶片前缘与压力面的静压较高，这要求冷气需要较高的总压才能保证稳定出流和必要的逆流裕度；而吸力面的静压相对较低，静压从前缘朝吸力面方向的分布差异通常较大，在吸力面侧设置串联双层壁冷却单元，冷气经过多次冲击流动后从气膜孔出流，仍能保证具有一定的总压和逆流裕度。既能够保证动叶片内部的强化换热，还能够用于不同区域的冷气用量调控，从而有效降低高压涡轮动叶的热负荷。
较佳地，所述串联双层壁冷却单元沿周向方向绕所述内腔室设置。
在本方案中，内腔室还能够对设于周侧的串联双层壁冷却单元内的气体进行冷却，降低串联双层壁冷却单元中的冷气因多次冲击换热而产生的升温。
较佳地，位于吸力面的所述冷却腔室的通道宽度与所述第一冲击孔的孔径的比值范围为2∶1至6∶1之间；
和/或，所述第一冲击孔的冲击距与所述第一冲击孔的孔径的比值范围为0.5∶1至2∶1之间。
较佳地，所述气膜孔包括第一气膜孔和第二气膜孔，首端的所述冷却腔室通过所述外层壁上的第一气膜孔与外部连通；
尾端的所述冷却腔室通过所述外层壁上的第二气膜孔与外部连通。
在本方案中，进入内腔室的冷气通过冲击孔进入冷却腔室，然后一部分冷气通过第一气膜孔流出，另外一部分冷气依次流过冷却腔室，然后从第二气膜孔流出，冷气在流动的过程中，对动叶片进行冲击冷却。
较佳地，所述冷却腔室的数量为两个，两个所述冷却腔室之间还设置有回流通道，首端的所述冷却腔室通过轴向冲击孔与所述回流通道连通，尾端的所述冷却腔室通过第五冲击孔与回流通道。
较佳地，所述轴向冲击孔的孔径不超过所述第一冲击孔的冲击距。
较佳地，所述串联双层壁冷却单元靠近所述动叶片的尾缘设置。
在本方案中，在动叶片的尾缘位置设置串联双层壁冷却单元，既保证了动叶片的强化换热，又能够在尾缘位置节省冷气用量，从而有效降低高压涡轮的动叶片的热负荷。由于高压涡轮的动叶片的表面静压从压力面中间位置朝尾缘方向呈降低趋势，冷气经多次冲击射流并从尾缘劈缝流出，还能保证具有一定的总压和逆流裕度。
较佳地，所述气膜孔还包括第三气膜孔，首端的所述冷却腔室通过所述第三气膜孔与外部连通，尾端的所述冷却腔室通过所述尾缘劈缝与外部连通；
所述冲击孔包括第二冲击孔，首端的所述冷却腔室通过所述第二冲击孔与所述内腔室连通。
在本方案中，进入内腔室的冷气通过第二冲击孔进入冷却腔室，然后一部分冷气通过第三气膜孔流出，另外一部分冷气依次流过冷却腔室，然后从尾缘劈缝流出，冷气在流动的过程中，对动叶片进行冲击冷却。
较佳地，所述第二冲击孔的冲击距与所述第二冲击孔的孔径的比值范围为0.5∶1至2∶1之间。
较佳地，所述冷却单元还包括前缘双层壁冷却单元，所述前缘双层壁冷却单元靠近所述动叶片的前缘设置，所述冲击孔包括第三冲击孔，所述前缘双层壁冷却单元通过所述第三冲击孔与所述内腔室连通。
在本方案中，前缘和压力侧区域的燃气温度高，因此，在前缘位置设置前缘双层壁冷却单元，降低动叶片前缘位置的热负荷。
较佳地，所述第三冲击孔的冲击距与所述第三冲击孔的孔径的比值范围为1∶1至3∶1之间；
其中，所述冲击距为所述第三冲击孔距所述外层壁的距离。
较佳地，所述冷却单元还包括压力面双层壁冷却单元，所述压力面双层壁冷却单元靠近所述动叶片的压力面设置，所述冲击孔还包括第四冲击孔，所述压力面双层壁冷却单元通过所述第四冲击孔与所述内腔室连通。在本方 案中，前缘和压力侧区域的燃气温度高，因此，在压力面位置设置压力面双层壁冷却单元，降低动叶片压力面位置的热负荷。
较佳地，位于压力面的所述冷却腔室的通道宽度与所述第四冲击孔的孔径的比值范围为2∶1至4∶1之间；
和/或，所述第四冲击孔的冲击距与所述第四冲击孔的孔径的比值范围为0.5∶1至2∶1之间。
较佳地，所述冲击孔的数量根据所述内腔室的容积大小、所述冷却腔室的容积大小以及充入所述内腔室的冷气流量来确定。
较佳地，所述冲击孔的孔径为0.8mm-1.5mm。
较佳地，所述气膜孔的数量及排数根据所述冷却腔室的通道宽度和流入所述冷却腔室的冷气流量来确定。
较佳地，所述气膜孔的孔径与所述冲击孔的孔径的比值范围为0.4∶1至0.6∶1之间。
较佳地，所述冷却腔室内还间隔设置有多个扰流柱。
在本方案中，设置扰流柱用于增加冷气的流动路径，提高冷却效果。
较佳地，所述扰流柱的直径与所述冲击孔的孔径的比值范围为0.8∶1至1.2∶1之间。
较佳地，所述内腔室的内部还设置有肋条。
在本方案中，强化内腔室的流道内部的对流换热。
本发明的积极进步效果在于：该高压涡轮的动叶片在其内部设置有冷却单元，冷却单元用于对动叶片的内部不同区域进行降温，并且在冷却单元中包括至少一个串联双层壁冷却单元，在一定量的冷却气量的情况下，冷气依次流过动叶片内的多个冷却腔室，增加冷气流动的路径，以在动叶片内进行多次连续冲击射流，实现动叶片内部的局部强化传热，同时也无需额外增加冷气用量，以便将冷气尽量多的供给动叶片温度较高的区域进行冷却，由此实现了冷气用量的调控，有效降低高压涡轮的动叶片的热负荷。
在符合本领域常识的基础上，上述各优选条件，可任意组合，即得本发 明各较佳实例。
附图说明
图1为本发明一较佳实施例1的高压涡轮的动叶片的立体结构示意图。
图2为图1中沿A-A线的剖面图。
图3为本发明一较佳实施例1的高压涡轮的动叶片的吸力面串联双层壁冷却单元的立体结构示意图。
图4为本发明一较佳实施例1的设有扰流柱的吸力面串联双层壁冷却单元的立体结构示意图。
图5为本发明一较佳实施例1的设有扰流柱的尾缘串联双层壁冷却单元的立体结构示意图。
图6为本发明一较佳实施例2的高压涡轮的动叶片的剖面结构示意图。
附图标记：
动叶片1
外层壁11
前缘111
压力面112
吸力面113
尾缘劈缝114
尾缘115
第一气膜孔1101
第二气膜孔1102
第三气膜孔1103
内层壁12
第一冲击孔121
第三冲击孔122
第四冲击孔123
第二冲击孔124
第五冲击孔125
轴向冲击孔126
缘板2
榫头3
第一内腔室I
第二内腔室II
第三内腔室III
第四内腔室IV
第五内腔室V
冷却腔室100
前缘双层壁冷却单元151
吸力面串联双层壁冷却单元152
回流通道153
压力面双层壁冷却单元154
尾缘串联双层壁冷却单元155
横隔板13
内腔室14
扰流柱16
冷气流向200
具体实施方式
下面通过实施例的方式进一步说明本发明，但并不因此将本发明限制在所述的实施例范围之中。
实施例1
如图1-图5所示，本实施例公开了一种高压涡轮的动叶片。如图1所示，动叶片1具有缘板2、榫头3、前缘111、尾缘115以及在前缘111和 尾缘115之间的压力面112和吸力面113。动叶片1还具有外层壁11、内层壁12和内腔室14。动叶片1的内部设有冷却单元，冷却单元具有冷却腔室100，冷却腔室100通过冲击孔与内腔室连通。冷却单元包括至少一个串联双层壁冷却单元，串联双层壁冷却单元包括在外层壁11和内层壁12之间形成的多个冷却腔室100，多个冷却腔室100依次连通，冲击孔包括第一冲击孔121，首端的冷却腔室100通过第一冲击孔121与内腔室连通，尾端的冷却腔室100通过气膜孔或尾缘劈缝114与外部连通。涡轮叶片内部的冷气通过内层壁12上的冲击孔对外层壁11进行冲击冷却，最后从外层壁11上的气膜孔流出，对外壁面实施冷却保护。冷却单元用于对动叶片1的内部不同区域进行降温，并且在冷却单元中包括至少一个串联双层壁冷却单元，在一定量的冷却气量的情况下，冷气依次流过动叶片1内的多个冷却腔室100，增加冷气流动的路径，以在动叶片1内进行多次连续冲击射流，实现动叶片1内部的局部强化传热，同时也无需额外增加冷气用量，以便将冷气尽量多的供给动叶片1温度较高的区域进行冷却，由此实现了冷气用量的调控，有效降低高压涡轮的动叶片1的热负荷。
在本实施例中，动叶片1的内部被横隔板13沿轴向方向分隔为五个内腔室14，五个内腔室14从前缘111至尾缘115的方向依次为第一内腔室I、第二内腔室II、第三内腔室III、第四内腔室IV和第五内腔室V，第二内腔室II、第三内腔室III和第四内腔室IV通过回转通道连通。五个内腔室14被分为三路进气通道，其中，第一内腔室I和第五内腔室V分别为单独的一路进气通道，第二内腔室II、第三内腔室III和第四内腔室IV通过回转通道连通，形成一路进气通道。相应地，从榫头3的冷气通道进入的冷气分为三路进入动叶片1内不同的内腔室。
在其他的实施例中，动叶片的内部被横隔板沿轴向方向分隔为至少三个内腔室。冷气从榫头3的冷气通道进入不同内腔室，至少一个内腔室用于保障动叶片的前缘位置冷却，至少一个内腔室用于保障动叶片的尾缘位置冷却，至少一个内腔室用于保障动叶片的前缘和尾缘之间的区域冷却。
在本实施例中，串联双层壁冷却单元靠近动叶片1的吸力面113设置，又称吸力面串联双层壁冷却单元152。在动叶片1的吸力面113由于燃气高速流动，导致该位置具有较高的换热系数，使得动叶片1运行环境恶化。叶片前缘与压力面的静压较高，这要求冷气需要较高的总压才能保证稳定出流和必要的逆流裕度；而吸力面113的静压相对较低，静压从前缘111朝吸力面113方向的分布差异通常较大，在吸力面113侧设置串联双层壁冷却单元，冷气经过多次冲击流动后从气膜孔出流，仍能保证具有一定的总压和逆流裕度。既能够保证动叶片1内部的强化换热，还能够用于不同区域的冷气用量调控，从而有效降低高压涡轮动叶的热负荷。
串联双层壁冷却单元沿周向方向绕所述内腔室设置，使内腔室的冷却气体还能够对串联双层壁冷却单元中流动的气体降温。在本实施例中，如图2所示，吸力面串联双层壁冷却单元152沿周向方向绕第一内腔室I和第二内腔室II设置。首端的冷却腔室100通过轴向冲击孔126与回流通道153连通，尾端的冷却腔室100通过第五冲击孔125与回流通道153连通。回流通道153将相邻的两个冷却腔室100连通，第二内腔室II还能够对设于周侧的回流通道153内的气体进行冷却，降低串联双层壁冷却单元中的冷气因多次冲击换热而产生的升温。在本实施例中，第五冲击孔的设置参数与第一冲击孔的设置参数相同。
如图2所示，在吸力面串联双层壁冷却单元152中，气膜孔包括第一气膜孔1101和第二气膜孔1102，首端的冷却腔室100通过外层壁11上的第一气膜孔1101与外部连通。尾端的冷却腔室100通过外层壁11上的第二气膜孔1102与外部连通。进入第一内腔室I的冷气通过第一冲击孔121进入首端的冷却腔室100并对该腔室进行冷却，然后一部分冷气通过第一气膜孔1101流出，另外一部分冷气通过轴向冲击孔126进入回流通道153并对该通道进行冷却，再通过第五冲击孔125进入尾端的冷却腔室100并对该腔室进行冷却，然后从第二气膜孔1102流出，冷气在流动的过程中，对动叶片1进行冲击冷却。
优选地，位于吸力面的冷却腔室100的通道宽度与第一冲击孔121的孔径的比值范围为2∶1至6∶1之间。
在本实施例中，在动叶片1的尾缘115位置还设置有串联双层壁冷却单元，又称尾缘串联双层壁冷却单元155。尾缘串联双层壁冷却单元155既保证了动叶片1尾缘115位置的强化换热，又能够在尾缘115位置节省冷气用量，从而有效降低高压涡轮的动叶片1的热负荷。由于高压涡轮的动叶片1的表面静压从压力面112中间位置朝尾缘115方向呈降低趋势，冷气经多次冲击射流并从尾缘劈缝114流出，还能保证具有一定的总压和逆流裕度。
在本实施例中，第一冲击孔121的冲击距与第一冲击孔121的孔径的比值范围为0.5∶1至2∶1之间。轴向冲击孔126的孔径不超过第一冲击孔121的冲击距。其中，第一冲击孔121的冲击距为第一冲击孔121距外层壁11的距离。
如图2所示，在尾缘串联双层壁冷却单元155中，气膜孔还包括第三气膜孔1103，首端的冷却腔室100通过第三气膜孔1103与外部连通。尾端的冷却腔室100通过尾缘劈缝114与外部连通。冲击孔包括第二冲击孔124，首端的冷却腔室100通过一第二冲击孔124与内腔室V连通。首端和尾端的冷却腔室100通过另一第二冲击孔124连通。
在本实施例中，进入第五内腔室V的冷气通过一第二冲击孔124进入首端的冷却腔室100，然后一部分冷气通过第三气膜孔1103流出，另外一部分冷气经过另一第二冲击孔124流入下一个冷却腔室100并对该腔室进行冷却，然后从尾缘劈缝114流出，冷气在流动的过程中，对动叶片1进行冲击冷却。
在本实施例中，第二冲击孔的冲击距与第二冲击孔的孔径的比值范围为0.5∶1至2∶1之间。其中，第二冲击孔的冲击距为第二冲击孔距外层壁11的距离。
在此基础上，冷却单元还包括前缘双层壁冷却单元151，前缘双层壁冷却单元151靠近动叶片1的前缘设置，冲击孔包括第三冲击孔122，前缘双 层壁冷却单元151通过第三冲击孔122与第一内腔室I连通。由于前缘111和压力侧区域的燃气温度高，因此，在前缘111位置设置前缘双层壁冷却单元151，降低动叶片1前缘111位置的热负荷。
优选地，第三冲击孔122的冲击距与第三冲击孔122的孔径的比值范围为1∶1至3∶1之间。其中，冲击距为第三冲击孔122距外层壁11的距离。
进一步地，冷却单元还包括压力面双层壁冷却单元154，压力面双层壁冷却单元154靠近动叶片1的压力面设置，冲击孔还包括第四冲击孔123，压力面双层壁冷却单元154通过第四冲击孔123与内腔室连通。由于前缘111和压力侧区域的燃气温度高，因此，在压力面112位置设置压力面双层壁冷却单元154，降低动叶片1压力面112位置的热负荷。在本实施例中，压力面双层壁冷却单元154为三个，三个力面双层壁冷却单元分别与第二内腔室II、第三内腔室III和第四内腔室IV一一对应设置，降低动叶片1压力面112位置的热负荷。
优选地，位于压力面的冷却腔室100的通道宽度与第四冲击孔123的孔径的比值范围为2∶1至4∶1之间。
在本实施例中，冲击孔的数量根据内腔室的容积大小、冷却腔室100的容积大小以及充入内腔室的冷气流量来确定。
优选地，第四冲击孔123的冲击距与第四冲击孔123的孔径的比值范围为0.5∶1至2∶1之间。其中，第四冲击孔123的冲击距为第四冲击孔123距外层壁11的距离。
优选地，冲击孔的孔径为0.8mm-1.5mm。
在本实施例中，吸力面串联双层壁冷却单元152通过一排第一冲击孔121与第一内腔室I连通。单排第一冲击孔121数量为8-12个。
在本实施例中，前缘双层壁冷却单元151与第一内腔室I之间布置1-3排第三冲击孔122。单排第三冲击孔122数量为8-12个。
在本实施例中，三个压力面双层壁冷却单元154分别通过一排第四冲击孔123与第二内腔室II、第三内腔室III和第四内腔室IV连通。单排第四冲 击孔123数量为8-12个。
在本实施例中，尾缘串联双层壁冷却单元155通过一排第二冲击孔124与第五内腔室V连通。单排第二冲击孔124数量为8-12个。
在本实施例中，气膜孔的数量及排数根据冷却腔室100的通道宽度和流入冷却腔室100的冷气流量来确定。气膜孔的孔径与冲击孔的孔径的比值范围为0.4∶1至0.6∶1之间。
在本实施例中，在前缘双层壁冷却单元151中，前缘111位置的外层壁11布置有4-8排气膜孔，单排气膜孔数量为14-20个。
在本实施例中，在吸力面串联双层壁冷却单元152中，每排第一冲击孔121的下游位置布置1-2排气膜孔，单排气膜孔数量为14-20个。
在本实施例中，在压力面双层壁冷却单元154中布置一排气膜孔，单排气膜孔数量为14-20个。
在本实施例中，在尾缘串联双层壁冷却单元155在压力面112位置布置1-2排气膜孔，单排气膜孔数量为14-20个。
在本实施例中，冷却腔室100内还间隔设置有多个扰流柱16，用于增加冷气的流动路径，提高冷却效果。
进一步地，扰流柱16的直径与冲击孔的孔径的比值范围为0.8∶1至1.2∶1之间。
在压力面双层壁冷却单元154内，扰流柱16与第四冲击孔123沿径向交替布置为一排。
如图4所示，在串联双层壁冷却单元内，第一冲击孔121布置在上游，扰流柱16布置在下游。
当冷却单元的轴向宽度较大时，可以增加双层壁通道内冲击孔和扰流柱16的排数。
内腔室的内部还设置有肋条，以强化内部扰流，提高内腔室的流道内部的对流换热。优选地，肋条高度与内腔室通道高度的比值范围为1∶5至1∶10之间。
在本实施例中，动叶片1内部的气体流动过程如下。动叶片1内的冷气流向200如图2中箭头所示，第一路冷气从榫头3的冷气通道进口进入第一内腔室I，一部分冷气通过第三冲击孔122进入前缘双层壁单元151，并从前缘111位置的气膜孔流出，另一部分冷气通过第一冲击孔121进入吸力面串联双层壁冷却单元152，并重复2次冲击-气膜冷却和1次轴向冲击冷却，分别从第一气膜孔1101和第二气膜孔1102流出。第二路冷气首先进入第二内腔室II，一部分冷气直接通过第四冲击孔123进入压力面双层壁单元154并从气膜孔流出，另一部分冷气通过回转通道依次流经第三内腔室III和第四内腔室IV，同时实施冲击-气膜冷却。流经第二内腔室II的冷气也对吸力面串联双层壁冷却单元152中的冷气起到冷却作用。第三路冷气从第五内腔室V进入叶身，通过一第二冲击孔124进入尾缘串联双层壁冷却单元155，其中一部分冷气通过第三气膜孔1103流出，另一部分冷气朝下游流动，从另一第二冲击孔124进入冷却腔室100并冲击扰流柱16，最后从尾缘劈缝114流出。
实施例2
如图6所示，本实施例与实施例1基本相同，其不同之处在于：在吸力面串联双层壁冷却单元中，设置有三个冷却腔室，从而实现3次冲击-气膜冷却和2两次轴向冲击冷却，提高动叶片的冷却效果。冷气流向如图6中箭头所示。
在其他的实施例中，冷却腔室的数量可根据选择设定，以实现节约用气和高压涡轮的动叶片内部的强化换热的双重作用。
虽然以上描述了本发明的具体实施方式，但是本领域的技术人员应当理解，这仅是举例说明，本发明的保护范围是由所附权利要求书限定的。本领域的技术人员在不背离本发明的原理和实质的前提下，可以对这些实施方式做出多种变更或修改，但这些变更和修改均落入本发明的保护范围。

Claims (22)

  1. 一种高压涡轮的动叶片，其特征在于，所述动叶片具有外层壁、内层壁和内腔室，所述动叶片的内部设有冷却单元，所述冷却单元具有冷却腔室，所述冷却腔室通过冲击孔与所述内腔室连通，所述冷却单元包括至少一个串联双层壁冷却单元，所述串联双层壁冷却单元包括在所述外层壁和所述内层壁之间形成的多个所述冷却腔室，多个所述冷却腔室依次连通，所述冲击孔包括第一冲击孔，首端的所述冷却腔室通过所述第一冲击孔与所述内腔室连通，尾端的所述冷却腔室通过气膜孔或尾缘劈缝与外部连通。
  2. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述动叶片的内部被横隔板沿轴向方向分隔为至少三个所述内腔室。
  3. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述串联双层壁冷却单元靠近所述动叶片的吸力面设置。
  4. 如权利要求3所述的高压涡轮的动叶片，其特征在于，所述串联双层壁冷却单元沿周向方向绕所述内腔室设置。
  5. 如权利要求3所述的高压涡轮的动叶片，其特征在于，位于吸力面的所述冷却腔室的通道宽度与所述第一冲击孔的孔径的比值范围为2∶1至6∶1之间；
    和/或，所述第一冲击孔的冲击距与所述第一冲击孔的孔径的比值范围为0.5∶1至2∶1之间。
  6. 如权利要求3所述的高压涡轮的动叶片，其特征在于，所述气膜孔包括第一气膜孔和第二气膜孔，首端的所述冷却腔室通过所述外层壁上的第一气膜孔与外部连通；
    尾端的所述冷却腔室通过所述外层壁上的第二气膜孔与外部连通。
  7. 如权利要求3所述的高压涡轮的动叶片，其特征在于，所述冷却腔室的数量为两个，两个所述冷却腔室之间还设置有回流通道，首端的所述冷却腔室通过轴向冲击孔与所述回流通道连通，尾端的所述冷却腔室通过第五冲击孔与回流通道。
  8. 如权利要求7所述的高压涡轮的动叶片，其特征在于，所述轴向冲击孔的孔径不超过所述第一冲击孔的冲击距。
  9. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述串联双层壁冷却单元靠近所述动叶片的尾缘设置。
  10. 如权利要求9所述的高压涡轮的动叶片，其特征在于，所述气膜孔还包括第三气膜孔，首端的所述冷却腔室通过所述第三气膜孔与外部连通，尾端的所述冷却腔室通过所述尾缘劈缝与外部连通；
    所述冲击孔包括第二冲击孔，首端的所述冷却腔室通过所述第二冲击孔与所述内腔室连通。
  11. 如权利要求10所述的高压涡轮的动叶片，其特征在于，所述第二冲击孔的冲击距与所述第二冲击孔的孔径的比值范围为0.5∶1至2∶1之间。
  12. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述冷却单元还包括前缘双层壁冷却单元，所述前缘双层壁冷却单元靠近所述动叶片的前缘设置，所述冲击孔包括第三冲击孔，所述前缘双层壁冷却单元通过所述第三冲击孔与所述内腔室连通。
  13. 如权利要求12所述的高压涡轮的动叶片，其特征在于，所述第三冲击孔的冲击距与所述第三冲击孔的孔径的比值范围为1∶1至3∶1之间；
    其中，所述冲击距为所述第三冲击孔距所述外层壁的距离。
  14. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述冷却单元还包括压力面双层壁冷却单元，所述压力面双层壁冷却单元靠近所述动叶片的压力面设置，所述冲击孔还包括第四冲击孔，所述压力面双层壁冷却单元通过所述第四冲击孔与所述内腔室连通。
  15. 如权利要求14所述的高压涡轮的动叶片，其特征在于，位于压力面的所述冷却腔室的通道宽度与所述第四冲击孔的孔径的比值范围为2∶1至4∶1之间；
    和/或，所述第四冲击孔的冲击距与所述第四冲击孔的孔径的比值范围为0.5∶1至2∶1之间。
  16. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述冲击孔的数量根据所述内腔室的容积大小、所述冷却腔室的容积大小以及充入所述内腔室的冷气流量来确定。
  17. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述冲击孔的孔径为0.8mm-1.5mm。
  18. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述气膜孔的数量及排数根据所述冷却腔室的通道宽度和流入所述冷却腔室的冷气流量来确定。
  19. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述气膜孔的孔径与所述冲击孔的孔径的比值范围为0.4∶1至0.6∶1之间。
  20. 如权利要求1所述的高压涡轮的动叶片，其特征在于，所述冷却腔室内还间隔设置有多个扰流柱。
  21. 如权利要求20所述的高压涡轮的动叶片，其特征在于，所述扰流柱的直径与所述冲击孔的孔径的比值范围为0.8∶1至1.2∶1之间。
  22. 如权利要求1-21中任一项所述的高压涡轮的动叶片，其特征在于，所述内腔室的内部还设置有肋条。
PCT/CN2023/108842 2022-07-22 2023-07-24 高压涡轮的动叶片 WO2024017385A1 (zh)

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Citations (5)

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Publication number Priority date Publication date Assignee Title
CN103277145A (zh) * 2013-06-09 2013-09-04 哈尔滨工业大学 一种燃气涡轮冷却叶片
CN205382958U (zh) * 2016-03-02 2016-07-13 中航商用航空发动机有限责任公司 涡轮叶片以及航空发动机
US20170356295A1 (en) * 2016-06-13 2017-12-14 General Electric Company Turbine component cooling holes
CN108999645A (zh) * 2017-06-07 2018-12-14 安萨尔多能源瑞士股份公司 用于燃气涡轮的叶片和包括所述叶片的电力生成设备
CN109139128A (zh) * 2018-10-22 2019-01-04 中国船舶重工集团公司第七0三研究所 一种船用燃气轮机高压涡轮导叶冷却结构

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103277145A (zh) * 2013-06-09 2013-09-04 哈尔滨工业大学 一种燃气涡轮冷却叶片
CN205382958U (zh) * 2016-03-02 2016-07-13 中航商用航空发动机有限责任公司 涡轮叶片以及航空发动机
US20170356295A1 (en) * 2016-06-13 2017-12-14 General Electric Company Turbine component cooling holes
CN108999645A (zh) * 2017-06-07 2018-12-14 安萨尔多能源瑞士股份公司 用于燃气涡轮的叶片和包括所述叶片的电力生成设备
CN109139128A (zh) * 2018-10-22 2019-01-04 中国船舶重工集团公司第七0三研究所 一种船用燃气轮机高压涡轮导叶冷却结构

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