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CN-122016221-A - Cooling turbine pneumatic performance test device

CN122016221ACN 122016221 ACN122016221 ACN 122016221ACN-122016221-A

Abstract

The invention relates to the technical field of aeroengine tests and discloses a cooling turbine pneumatic performance test device which comprises a main flow high-temperature fuel gas flow channel, a turbine rotor, a rotating shaft, a supporting structure and a sliding cavity, wherein blades of the turbine rotor are arranged on the main flow high-temperature fuel gas flow channel, one end of the rotating shaft is connected with the turbine rotor, the other end of the rotating shaft is suitable for being connected with a dynamometer of a test stand, the supporting structure surrounds the outer side of the rotating shaft, a bearing is arranged between the supporting structure and the rotating shaft and is positioned at the downstream of the turbine rotor, and the sliding cavity is formed between the supporting structure and the rotating shaft. The whole lubricating oil cavity is positioned at the downstream of the turbine rotor, the temperature of the air inlet of the main flow high-temperature fuel gas flow channel is extremely high, the temperature of the air flow can be obviously reduced after the air flow is expanded and acted by the blades of the turbine rotor, and the bearing and the lubricating oil cavity are arranged at the downstream of the turbine rotor, so that the bearing and the lubricating oil cavity are far away from a high-temperature core area, the direct heat radiation and the structural heat conduction of high-temperature fuel gas are reduced from the space layout, and the heat input of the lubricating oil cavity is reduced from the source.

Inventors

  • FANG HUA
  • CHEN BING
  • JIANG KANGHE
  • TAO JIANJUN
  • CHEN RUIYUAN
  • WU ZHIFAN
  • HU HANHUI

Assignees

  • 中国航发湖南动力机械研究所

Dates

Publication Date
20260512
Application Date
20260310

Claims (10)

  1. 1. A cooling turbine aerodynamic performance test device, comprising: a main flow high temperature gas flow path (1); a turbine rotor (2), wherein blades of the turbine rotor (2) are arranged in the main flow high-temperature gas flow channel (1); one end of the rotating shaft (3) is connected with the turbine rotor (2), and the other end of the rotating shaft is suitable for being connected with a dynamometer of the test bed; The support structure (4) surrounds the outside of the rotating shaft (3), a bearing (5) is arranged between the support structure (4) and the rotating shaft (3), the bearing (5) is positioned at the downstream of the turbine rotor (2), and a lubricating oil cavity (6) is formed between the support structure (4) and the rotating shaft (3).
  2. 2. The cooling turbine aerodynamic performance test device according to claim 1, characterized in that a heat insulation chamber (10) is arranged between the main flow high temperature gas flow channel (1) and the lubricating oil chamber (6).
  3. 3. The cooling turbine aerodynamic performance test device according to claim 2, characterized in that the insulation chamber (10) is provided with a cold air inlet (1001) and a cold air outlet (1002), the cold air inlet (1001) being in communication with a directional flow guiding system.
  4. 4. A cooling turbine aerodynamic performance test device according to claim 3, characterized in that a heat insulation plate (11) is arranged in the heat insulation cavity (10), the heat insulation plate (11) is parallel to the axis of the rotating shaft (3), the heat insulation plate (11) divides the heat insulation cavity (10) into a plurality of cooling channels distributed along the radial direction, two adjacent cooling channels are communicated through a communication port (1101), the cold air inlet (1001) is communicated with the cooling channel located at the innermost side, and the cold air outlet (1002) is communicated with the cooling channel located at the outermost side.
  5. 5. The cooling turbine aerodynamic performance test device according to claim 4, characterized in that the communication ports (1101) on two adjacent heat shields (11) are distributed in axial offset.
  6. 6. The cooling turbine aerodynamic performance test device according to claim 4, characterized in that the cold air inlet (1001) communicates with the center of the innermost cooling flow channel.
  7. 7. The cooling turbine aerodynamic performance test device according to any of claims 3-6, characterized in that a temperature sensor is arranged in the lubricating oil cavity (6), the directional diversion system comprises a flow regulating valve and a controller, and the controller is in communication connection with the temperature sensor and the flow regulating valve.
  8. 8. The cooling turbine aerodynamic performance test device according to any of claims 1-6, characterized in that the main flow high temperature gas flow channel (1) has an inner annular peripheral wall (101) and an outer annular peripheral wall (102), a first sealing assembly being provided between the turbine rotor (2) and the inner annular peripheral wall (101); a second sealing component is arranged between the joint of the turbine rotor (2) and the rotating shaft (3) and the lubricating oil cavity (6); And a third sealing assembly is arranged between one end, far away from the turbine rotor (2), of the supporting structure (4) and the rotating shaft (3).
  9. 9. The cooling turbine aerodynamic performance test device of claim 8, wherein the inner annular peripheral wall (101) comprises a first peripheral wall section (1011) and a second peripheral wall section (1012), the blades of the turbine rotor (2) passing through a gap between the first peripheral wall section (1011) and the second peripheral wall section (1012), the first seal assembly comprising: The first connecting piece (12) is connected with the first peripheral wall section (1011), a first sealing cavity (13) is formed between the first connecting piece (12) and the turbine in a surrounding mode, and the first connecting piece (12) is provided with a first groove (1201) and a first connecting part (1202); a first flange (201) provided on a side of the turbine rotor (2) facing the first peripheral wall section (1011), the first flange (201) being inserted into the first groove (1201); A second flange (202) provided on a side of the turbine rotor (2) facing the second peripheral wall section (1012), the second flange (202) extending to an outer peripheral side of the second peripheral wall section (1012); And a third flange (203) arranged on one side of the turbine rotor (2) facing the first peripheral wall section (1011), wherein a grating tooth sealing structure is formed between the third flange (203) and the first connecting part (1202).
  10. 10. The cooling turbine aerodynamic performance test device of claim 9, wherein the second seal assembly comprises: The second connecting piece (14) is fixedly connected with the supporting structure (4), a second sealing cavity (15) is formed between the second connecting piece (14), the supporting structure (4), the second peripheral wall section (1012) and the turbine rotor (2), a third sealing cavity (16) is formed between the second connecting piece (14) and the rotating shaft (3), and the second connecting piece (14) is provided with a second connecting part (1401) and a third connecting part (1402); The fourth convex edge (204) is arranged on one side, facing the supporting structure (4), of the turbine rotor (2), a grate sealing structure is formed between the fourth convex edge (204) and the second connecting part (1401), and one end of the rotating shaft (3) is fixedly connected with the fourth convex edge (204); And a grate tooth sealing structure is formed between the third connecting part (1402) and the rotating shaft (3).

Description

Cooling turbine pneumatic performance test device Technical Field The invention relates to the technical field of aeroengine tests, in particular to a cooling turbine pneumatic performance test device. Background The turbine is a core power component of an aeroengine, and the aerodynamic performance of the turbine directly determines the thrust, fuel efficiency and reliability of the engine. In order to develop the high-performance aero-engine, the real working condition is simulated through a cooling turbine aerodynamic performance test, and the influence of parameters such as a cold air injection mode, flow and the like on turbine efficiency and aerodynamic stability is analyzed. In aeroengine cooling turbine aerodynamic performance test, high temperature fuel gas (the temperature can reach 800 ℃ to 1200 ℃ or even higher) in a turbine main runner can transfer heat to a bearing oil cavity through heat conduction, heat radiation and other modes, so that the temperature of the lubricating oil exceeds the allowable working temperature (usually less than or equal to 150 ℃), carbonization of the lubricating oil is caused, viscosity is reduced or even the lubricating oil fails, and bearing abrasion and clamping stagnation are caused, so that test safety is seriously threatened. Meanwhile, the stable operation of the turbine rotor can be damaged due to the failure of lubricating oil, the real working condition of the engine cannot be accurately simulated, the research data of the influence of cold air on the turbine performance is distorted, and the real performance of the turbine is difficult to accurately verify. In the related art, the test piece structure is arranged in a symmetrical layout structure, two supporting points of the bearing are positioned at the front end and the rear end of the turbine rotor blade, the temperature of the airflow at the front end of the turbine rotor blade is highest, and the temperature can be reduced after the high-temperature airflow pushes the turbine rotor blade to do work, so that the heat conduction and heat radiation of the bearing and the lubricating oil positioned at the front end of the turbine rotor blade are high due to the high-temperature airflow, and the failure risk is easy to occur. Disclosure of Invention In view of the above, the invention provides a cooling turbine aerodynamic performance test device to solve the problem that lubricating oil is easy to fail in an aeroengine cooling turbine aerodynamic performance test. The invention provides a cooling turbine aerodynamic performance test device, which comprises: A main flow high temperature gas flow path; The blades of the turbine rotor are arranged on the main flow high-temperature fuel gas flow passage; One end of the rotating shaft is connected with the turbine rotor, and the other end of the rotating shaft is suitable for being connected with a dynamometer of the test bed; The support structure surrounds the outside of pivot, support structure with be equipped with the bearing between the pivot, the bearing is located turbine rotor's low reaches, support structure with be formed with the oil slick chamber between the pivot. The high-temperature gas flow channel has the beneficial effects that the blades of the turbine rotor are arranged in the main flow high-temperature gas flow channel, when high-temperature gas passes through the main flow high-temperature gas flow channel, the turbine rotor is pushed to rotate, one end of the rotating shaft is connected with the turbine rotor, and the other end of the rotating shaft is connected with the dynamometer of the test bed, so that the rotating speed of the turbine rotor can be measured. Because the one end and the turbine rotor of pivot are connected, be equipped with the bearing between bearing structure and the pivot, the bearing is located turbine rotor's low reaches, be formed with the oil slick chamber between bearing structure and the pivot, consequently whole oil slick chamber is located turbine rotor's low reaches, and mainstream high temperature gas flow path air inlet temperature is high, after turbine rotor's blade inflation acting, the air current temperature can show the reduction, place bearing and oil slick chamber in turbine rotor's low reaches, can make bearing and oil slick chamber keep away from high temperature core area, reduce the direct heat radiation and the structure heat conduction of high temperature gas from spatial layout, reduce the heat input of oil slick chamber from the source, thereby avoid lubricating oil and bearing inefficacy. In addition, the bearing and the lubricating oil cavity are arranged at the downstream of the turbine rotor, so that the cantilever length of the rotating shaft can be shortened, the stability of the system is improved, and structural guarantee is provided for test accuracy. In an alternative embodiment, an insulating chamber is provided between the main flow high temperature gas flow passage and th