High-temperature turbine cascade test device based on shock tunnel

Abstract

The invention provides a high-temperature turbine cascade test device based on a shock tunnel, and the device comprises a shock tube, a transition section, a turbine cascade section, and a Laval nozzle section. The Mach number, the Reynolds number and the inlet total temperature of the turbine part of the aero-engine / gas turbine under the ground test working condition and the real machine design working condition can be reproduced, and the novel device has the advantages of being high in heating capacity, simple in structure, easy to operate and low in construction, operation and maintenance cost. Shock waves are generated through shock wave tube structures of the high-pressure driving section, the diaphragm and the low-pressure driven section, and high-temperature main flow is generated under the heating action of the shock waves so as to be matched with the incoming flow total temperature of a design working condition; in addition, airflow is organized through the transition section, the turbine cascade section and the Laval nozzle section so as to be matched with the Mach number and the Reynolds number of the design working condition.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine / gas turbine testing. Specifically, it relates to a high-temperature turbine cascade test device based on a shock tunnel, which is applicable to the test and measurement of high-temperature turbine blades. Background Art

[0002] The turbine blade is the core component of an aero-engine / gas turbine, directly affecting the performance of the aero-engine / gas turbine. To improve the overall thermal efficiency, the turbine inlet temperature of the aero-engine / gas turbine has been continuously increased, far exceeding the melting point of the turbine blade material. Taking the aero-engine products of Rolls-Royce in the UK as an example, the turbine inlet temperature of the RB211-524-G model (flying on a Boeing 747) is about 1700K, while the turbine inlet temperature of the latest Trent XWB model (flying on an Airbus A350) has been increased to more than 2000K. In the design, test, and verification of turbine blades, ground tests need to be carried out in a turbine tester to obtain the performance data of the aerodynamics and heat transfer of the turbine blades. Existing turbine testers often reduce the temperature level of the test air flow based on the similarity principle. For example, in the ground test of the high-pressure turbine of the GE-E3 core engine by NASA (Lewis research center), the inlet temperature of the first-stage rotor blade is about 620K, which is very different from the 1750K in the actual design condition of the engine. Therefore, there is an urgent need for a turbine test scheme with strong heating capacity.

[0003] Currently, the commonly used heating methods for turbine cascade test and measurement are: fuel combustion heating, electric heating, heat exchanger, gas compression heating, etc.

[0004] Fuel combustion heating means releasing heat by fuel combustion to generate a high-temperature gas source. The fuel is generally fossil energy or its products, such as natural gas, diesel, gasoline, etc. This method generally requires a high-power fan to drive to form a stable high-temperature air flow, and at the same time requires a complex auxiliary system. Representative patents include: the invention patent with the publication number CN105806874A and the application number CN201610153029.2, which discloses a full-temperature and equal-expansion-ratio cold effect test device for gas turbine blades, including a main flow system for turbine blade tests, a blade cooling air system, and a cooling water system. This device preheats air with an electric heater and introduces fuel into the burner for combustion, and can achieve a turbine inlet temperature of about 1700K. Shanghai Power Equipment Research Institute has also disclosed a cooling and equal-expansion-ratio cold effect test device for gas turbine blades (publication number CN105806873A, application number CN201610152930.8), which can achieve a turbine inlet temperature of about 800K. Similar solutions can also be seen in the invention patents with the publication numbers CN113916542A (a comprehensive test system and method applicable to testing the characteristics of turbine blades under high operating conditions) and CN112903276A (an open-type turbine blade test device). The advantages of this device are strong heating capacity and continuous generation of high-temperature gas; the disadvantages are complex heating system, high construction, operation, and maintenance costs, and high required thermal protection safety level, etc.

[0005] Electric heating generally heats porous metal devices, such as wire meshes, etc., through an external power source, converts electrical energy into Joule heat, and heats the passing air flow. Currently, this method has been widely used in ground tests of turbine cascades. Representative wind tunnel test benches include: the turbine cascade test bench of Shanghai Jiao Tong University, the turbine cascade test bench of the University of Oxford, etc. Representative patents include: the invention patent with the publication number CN112414739A and the application number CN202011315738.9, which discloses a gas turbine test bench that can conduct transient and steady-state measurement tests, including a bypass air path, an experimental air path, and an exhaust section. By simultaneously and quickly adjusting the direction of the valve ball core, the bypass air path and the experimental air path are switched to complete the transient and steady-state measurement tests of the turbine cascade channel. The advantages of this method are simple heating system, low cost, and easy control of heating power; the disadvantages are poor heating capacity and the mainstream temperature is usually less than 400K.

[0006] Heat exchanger means transferring the heat of a high-temperature heat source to the air flow in the form of convective heat transfer through a shell-and-tube heat exchanger or preheating copper tubes, etc. Currently, this method also has application cases in ground tests of turbine cascades. Representative wind tunnel test benches include: the Virginia Tech blowdown wind tunnel test bench. The advantages of this method are simple heating system and low cost; the disadvantages are low heating capacity, the mainstream temperature is usually less than 400K, and the control response of heating power is slow.

[0007] Gas compression heating converts mechanical energy into internal energy through the compression process to increase the temperature of the gas. In the turbine cascade test, this process is mainly achieved through an isentropic light piston compression tube device. Representative units using such devices include the University of Oxford, the von Karman Institute for Fluid Dynamics, etc. The advantages of this method are a simple heating system and low cost; the disadvantages are low heating capacity and a maximum mainstream temperature of less than 500K that can be achieved.

[0008] In summary, among the commonly used heating methods in current turbine cascade tests, the only solution that can match the total inlet temperature of the real engine design conditions of an aeroengine / gas turbine (~1700K) is fuel combustion heating. This solution has strong heating capacity, but also has disadvantages such as a complex heating system, high construction, operation, and maintenance costs, and a high required thermal protection safety level. On the contrary, solutions such as electric heating, heat exchangers, and gas compression heating have the disadvantage of low heating capacity, and the achievable mainstream temperature is generally less than 500K, which is even difficult to match the temperature level of ground tests of turbine testers. For example, the inlet temperature of the first-stage rotor blade of the GE-E3 high-pressure turbine ground test at NASA (Lewis Research Center) is approximately 620K. Therefore, there is an urgent need to develop a tester with strong heating capacity, simple structure, easy operation, and low construction, operation, and maintenance costs to match the ground test conditions of aeroengine / gas turbine turbine components and the temperature level of real engine design conditions.

[0009] The shock tunnel has the advantages of simple structure, low operation cost, and strong heating capacity, and generates a high-temperature gas source through the heating effect of the shock wave. By controlling the intensity of the shock wave, the generated mainstream temperature can reach above 10,000K, which is sufficient to match the ground test conditions of aeroengine / gas turbine turbine components and the temperature level of real engine design conditions. However, currently, this wind tunnel is mainly applied in the aerospace field and is rarely used in the field of aeroengine / gas turbine testing, and relevant test technologies have not been reported. Summary of the Invention

[0010] Aiming at the defects in the prior art, the purpose of the present invention is to provide a high-temperature turbine cascade test device based on a shock tunnel.

[0011] The high-temperature turbine cascade test device based on a shock tunnel provided by the present invention includes the following components connected in sequence: A shock tube for generating a shock wave and heating the test gas through the shock compression effect; A transition section for introducing the high-temperature test gas in the shock tube into the turbine cascade section; A turbine cascade section where the cascade test piece is placed inside; And a Laval nozzle section.

[0012] Preferably, the shock tube includes a high-pressure driving section, a diaphragm, and a low-pressure driven section; The high-pressure driving section is arranged upstream of the low-pressure driven section. The high-pressure driving section is a pipeline for storing high-pressure gas and serves as a driving source for generating high-temperature gas; The diaphragm is arranged between the high-pressure driving section and the low-pressure driven section and is used to isolate the gas in the high-pressure driving section from the test gas in the low-pressure driven section; When the pressure difference between the gas in the high-pressure driving section and the test gas in the low-pressure driven section reaches a preset value, the diaphragm bursts, causing the gas in the high-pressure driving section to expand and compress the test gas in the low-pressure driven section, and generating a shock wave in the test gas, thereby heating the test gas; The low-pressure driven section is a pipeline for storing test gas.

[0013] Preferably, a heavy piston is arranged in the high-pressure driving section to form a free-piston shock tunnel mode; A light piston is arranged in the low-pressure driven section to form a gun tunnel operation mode.

[0014] Preferably, the cross-sectional shape of one end of the transition section connected to the shock tube is the same as the cross-sectional shape of the shock tube pipeline, and the transition section is filled with gas.

[0015] Preferably, the turbine cascade section includes an inlet section, a cascade test piece, and an outlet section arranged in sequence, and the turbine cascade section is filled with gas; The inlet section is used to introduce the gas in the transition section and generate a uniformly distributed incoming flow at the inlet of the turbine cascade; The cascade test piece is a collection of blade test pieces, including a planar cascade test piece or an annular cascade test piece; The outlet section is used to control the outlet flow angle of the cascade test piece.

[0016] Preferably, the angle between the cascade frontal line of the cascade test piece and the upstream incoming flow direction is (90° - α), where α is the inlet flow angle of the cascade test piece.

[0017] Preferably, the blade test piece includes a turbine stator blade or a turbine rotor blade.

[0018] Preferably, the cross-sectional shapes of the inlet section and the outlet section are both matched with the structure of the cascade test piece; When the cascade test piece is a planar cascade test piece, the cross-sectional shapes of the inlet section and the outlet section are both square; When the cascade test piece is an annular cascade test piece, the cross-sectional shapes of the inlet section and the outlet section are both fan-shaped.

[0019] Preferably, the cross-sectional shape of one end of the transition section connecting the turbine cascade section matches the structure of the cascade test piece; When the cascade test piece is a planar cascade test piece, the cross-sectional shape of one end of the transition section connecting the turbine cascade section is square; When the cascade test piece is an annular cascade test piece, the cross-sectional shape of one end of the transition section connecting the turbine cascade section is fan-shaped.

[0020] Preferably, the Laval nozzle section has a structure that first contracts and then expands. The Laval nozzle section is filled with gas, and the size of the throat of the Laval nozzle section needs to satisfy that the mass flow rate at the throat of the Laval nozzle section is equal to the mass flow rate at the cascade outlet.

[0021] Compared with the prior art, the present invention has the following beneficial effects: 1. The present invention creatively applies a shock tunnel to the test of high-temperature turbine cascades, forming a new type of tester with strong heating ability, simple structure, easy operation, and low construction, operation, and maintenance costs, so as to match the temperature levels of the ground test conditions and the real engine design conditions of the turbine components of aeroengines / gas turbines.

[0022] 2. The high-temperature turbine cascade test device based on a shock tunnel provided by the present invention can simply and low-costly generate a mainstream above 600K, matching the cascade outlet Mach number, cascade inlet Reynolds number, and total inlet temperature of the ground test conditions of the turbine tester, and can more simply and low-costly generate a mainstream above 1700K, matching the cascade outlet Mach number, cascade inlet Reynolds number, and total inlet temperature of the real engine design conditions.

[0023] 3. The high-temperature turbine cascade test device based on a shock tunnel provided by the present invention has a simple structure, can be adapted to any blade and any arrangement method, and the Mach number of the cascade flow field remains constant under different shock tube design parameters (i.e., changing cascade inlet conditions). BRIEF DESCRIPTION OF THE DRAWINGS

[0024] By reading the detailed description of the non-limiting embodiments with reference to the following drawings, other features, objects, and advantages of the present invention will become more apparent: Figure 1 is a schematic diagram of the overall structure of the present invention; Figure 2 is a schematic diagram of the numerical simulation calculation domain and boundary conditions of an embodiment of the present invention; Figure 3 is the total inlet pressure of the cascade in Embodiment 1 of the present invention ( P t,i ), the static pressure at the cascade outlet ( P s,e ), and their pressure ratio ( P t,i / P s,e ) Schematic diagram of the variation curve with calculation time; Figure 4 Schematic diagram of the variation curve of the total inlet temperature of the cascade in Embodiment 1 and Embodiment 2 of the present invention with calculation time; Figure 5 In the present invention, for Embodiment 1 and Embodiment 2 within the effective test time, the distribution schematic diagram of the cascade outlet pressure loss coefficient ( P t,i -P t,e ) / ( P t,i -P s,e ) along the circumferential direction; Figure 6 In the present invention, for Embodiment 1 and Embodiment 2 within the effective test time, the distribution schematic diagram of the ratio of the static pressure on the surface of the middle blade to the total inlet pressure of the cascade along the axial direction; Figure 7 For the total inlet pressure (P t,i ) of the cascade, the static outlet pressure (P s,e ), and their pressure ratio (P t,i / P s,e ) of Embodiment 2 of the present invention, the schematic diagram of the variation curve with calculation time.

[0025] As shown in the figure: Specific embodiments The present invention will be described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but do not limit the present invention in any form. It should be noted that for those of ordinary skill in the art, without departing from the concept of the present invention, several changes and improvements can still be made. These all belong to the protection scope of the present invention.

[0026] Aiming at the problems in the commonly used heating methods in current turbine cascade tests, the present invention discloses a high-temperature turbine cascade test device based on a shock tunnel. This device can achieve a mainstream temperature above 600K, so as to match the temperature level of the ground test conditions of turbine components, and can also achieve a mainstream temperature above 1700K, so as to match the temperature level of the real machine design conditions of turbine components. Moreover, it has a simple structure, is easy to operate, and has low construction, operation and maintenance costs. The design objectives of this test device are as follows: First, the total inlet flow temperature is greater than 600K; second, the deviation of the cascade outlet Mach number and the cascade inlet Reynolds number from the design values is within ±5%; third, the effective test time is more than 6ms (meeting the measurement requirements of aerodynamic forces), and the aerodynamic parameters of the flow field are stable during this period; fourth, the cascade flow field is periodic; fifth, the Mach number of the flow field of this device remains constant under different shock tube design parameters (generating varying cascade inlet flow conditions).

[0027] To achieve the above objectives, the high-temperature turbine cascade test device based on a shock tunnel disclosed in the present invention includes a shock tube 1, a transition section 2, a turbine cascade section 3, and a Laval nozzle section 4 connected in sequence.

[0028] The shock tube 1 includes a high-pressure driver section 1-1, a diaphragm 1-2, and a low-pressure driven section 1-3. The function of the shock tube 1 is to generate a shock wave and heat the test gas through the shock compression effect. The design parameters of the shock tube 1 are: the length, initial pressure, initial temperature, gas type, pipe diameter, pipe cross-sectional shape, and whether there is a heavy piston, etc. in the high-pressure driver section 1-1, and the length, initial pressure, initial temperature, gas type, pipe diameter, pipe cross-sectional shape, and whether there is a light piston, etc. in the low-pressure driven section 1-3. These design parameters determine the flow conditions of the high-temperature gas generated by the shock tube 1, that is, the inlet flow conditions of the turbine cascade section 3.

[0029] The high-pressure driver section 1-1 is a pipe for storing high-pressure gas. Its cross-sectional shape can be arbitrary, such as circular, square, etc. Any gas can be filled therein, including: hydrogen, helium, nitrogen, carbon dioxide, air, etc. Its function is to store energy and serve as a driving source for generating high-temperature gas. In the high-pressure driver section 1-1, a heavy piston can be added to generate the free-piston shock tunnel mode, thereby increasing the pressure of the high-pressure driver section 1-1 and ultimately increasing the total fluid temperature at the outlet of the shock tube 1 (i.e., the inlet of the turbine cascade).

[0030] The function of the diaphragm 1-2 is to isolate the gas in the high-pressure driver section 1-1 from the test gas in the low-pressure driven section 1-3. Under a certain pressure difference, the diaphragm 1-2 bursts, so that the gas in the high-pressure driver section 1-1 expands, compresses the test gas in the low-pressure driven section 1-3, and generates a shock wave in the test gas to heat the test gas. The diaphragm 1-2 can be made of any material, and different set bursting pressure differences can be achieved by adjusting the thickness of the diaphragm 1-2, presetting scratches on the surface of the diaphragm 1-2, etc. Controlled bursting of the diaphragm 1-2 can also be achieved by means such as acupuncture, electric heating, and detonation devices.

[0031] The low-pressure driven section 1-3 is a pipeline for storing the test gas, which can be gas at any pressure and temperature, preferably air taken from the environment (pressure 1 bar, temperature 300K). Its function is to provide sufficient test gas. A lightweight piston can be added inside the low-pressure driven section 1-3 to generate the operation mode of the shock tunnel, thereby extending the effective test time.

[0032] The function of the transition section 2 is to connect the shock tube 1 and the turbine cascade section 3, and introduce the high-temperature gas in the driven section of the shock tube 1 into the turbine cascade section 3. Its cross-sectional shape can be arbitrary: if the cross-section of the shock tube 1 is square and the turbine cascade section 3 uses a planar cascade test piece 3-2, then the transition section 2 is a square tube; if the cross-section of the shock tube 1 is circular and the turbine cascade section 3 uses a planar cascade test piece 3-2, then the transition section 2 is circular to square; if the cross-section of the shock tube 1 is circular and the turbine cascade section 3 uses an annular cascade test piece 3-2, then the transition section 2 is circular to fan-shaped. Its initial state can be filled with gas at any pressure and temperature, preferably air taken from the environment (pressure 1 bar, temperature 300K). The turbine cascade section 3 consists of an inlet section 3-1, a cascade test piece 3-2, and an outlet section 3-3, and is used to place the test piece of the turbine blade to carry out experimental tests on aerodynamics, heat transfer, and flow. Its initial state can be filled with gas at any pressure and temperature, preferably air taken from the environment (pressure 1 bar, temperature 300K).

[0033] The function of the inlet section 3-1 is to introduce the gas in the transition section 2 and generate a uniformly distributed inlet flow for the turbine cascade. Its arrangement direction is the same as that of the shock tube 1 and the transition section 2. When using a planar cascade, its cross-sectional shape is square; when using an annular cascade, its cross-sectional shape is fan-shaped.

[0034] The cascade test piece 3-2 is a collection of blade test pieces, and can adopt the arrangement of a planar cascade or an annular cascade. The blade type can be a turbine stator blade or a turbine rotor blade. The profile lines of the end walls on both sides of the cascade can be selected as the boundary streamlines of the cascade design condition, or design optimization can be carried out. Its goal is to ensure the periodicity of the intermediate test blades. The included angle between the frontal line of the cascade and the oncoming flow direction upstream of it is (90° - α ) α which is the inlet flow angle of the selected cascade.

[0035] The function of the outlet section 3-3 is to control the outlet flow angle of the cascade to the designed value to ensure that the pressure drop ratio of the cascade (the ratio of the total pressure at the cascade inlet to the static pressure at the cascade outlet) matches the designed value, so as to match the Mach number at the cascade outlet. When using a planar cascade, its cross-sectional shape is square; when using an annular cascade, its cross-sectional shape is fan-shaped.

[0036] The nozzle section 4 has a structure that first contracts and then expands. The flow velocity at the throat is the speed of sound (Mach number is 1), and the flow downstream of the throat is supersonic. For a given turbine cascade section, the area of the throat of the nozzle can be determined by the principle of mass conservation (i.e., the mass flow rate through the cascade is equal to the mass flow rate at the throat of the nozzle). According to the area-Mach number relationship of quasi-one-dimensional flow in compressible flow theory, the Mach number at the exit of the cascade (i.e., the inlet of the nozzle) is only related to its area. Therefore, for a given geometric size of the turbine cascade section, the Mach number at the exit of the cascade remains constant and does not change with the incoming flow conditions of the cascade, that is, it is not affected by the design parameters of the shock tube. When a planar cascade is used, its cross-sectional shape is square; when an annular cascade is used, its cross-sectional shape is fan-shaped. Its initial state can be filled with gas at any pressure and temperature, and preferably air taken from the environment (pressure 1 bar, temperature 300K).

[0037] The present invention generates a shock wave through the shock tube 1 to heat the test gas to achieve the first above-mentioned objective. The second, fourth, and fifth objectives are achieved by designing the profiles of the turbine cascade section 3 and the nozzle section 4. The third objective is achieved by optimizing the design parameters of the shock tube 1 (such as extending the lengths of the high-pressure driving section 1-1 and the low-pressure driven section 1-3).

[0038] The operating principle of the present invention: High-pressure gas is filled into the high-pressure driving section 1-1 until the pressure reaches the bursting pressure of the diaphragm 1-2. The diaphragm 1-2 ruptures instantaneously, and the high-pressure gas in the high-pressure driving section 1-1 pushes the low-pressure gas in the low-pressure driven section 1-3, and a shock wave is formed in the low-pressure driven section 1-3 to heat the test gas. After the shock wave reaches the cascade test piece 3-2 and forms a reflection, the gas temperature rises again. After the unsteady startup process, the internal flow field parameters in the cascade become stable, and the effective test time begins. When the next shock wave or expansion wave reaches the cascade test piece 3-2, the effective test time terminates. The flow velocity at the throat of the nozzle section 4 is the speed of sound (Mach number is 1), and the flow downstream of the throat is supersonic; after the areas of the inlet and the throat of the nozzle section 4 are given, according to the area-Mach number relationship of quasi-one-dimensional flow in compressible flow theory, the Mach number at the inlet of the nozzle (i.e., the outlet section 3-3) remains constant and does not change with the incoming flow conditions of the cascade, that is, it is not affected by the design parameters of the shock tube. The optimization operation method of the present invention is as follows. Methods for increasing the total temperature of the incoming flow include: adding a heavy piston to the high-pressure driving section 1-1 of the shock tube to produce a free piston shock tunnel mode; increasing the pressure of the high-pressure driving section 1-1; filling the high-pressure driving section 1-1 with low molecular weight "light" gas, such as hydrogen, helium, etc.; heating the pipeline of the low-pressure driven section 1-3, etc. Methods for extending the effective test time include: adding a lightweight piston to the low-pressure driven section 1-3 of the shock tube to produce a gun wind tunnel operation mode; extending the pipeline length of the low-pressure driven section 1-3 and the high-pressure driving section 1-1; filling the end of the high-pressure driving section 1-1 with a high-molecular weight gas with a low sound velocity; inserting a driving plug-in into the high-pressure driving section 1-1; adjusting the pressure, temperature and filling gas type of the high-pressure driving section 1-1 and the low-pressure driven section 1-3 based on the suture conditions, etc. Methods for optimizing the periodicity of the cascade flow field include: optimizing the profile of the end walls on both sides of the cascade test piece 3-2; suctioning the boundary layers of the end walls on both sides of the cascade test piece 3-2; increasing the number of blades in the cascade test piece 3-2, etc.

[0039] The computational fluid dynamics simulation results show that the present invention can easily and cheaply generate a mainstream with a total temperature of more than 600K, matching the total temperature of the incoming flow, the Mach number at the cascade outlet, and the Reynolds number at the cascade inlet of the turbine tester ground test conditions. The effective test time is about 8 ms, and the flow field of the turbine cascade section 3 is periodic, and its Mach number remains constant under different shock tube design parameters. In addition, by optimizing the shock tube design parameters (such as filling the high-pressure drive section 1-1 with low molecular weight gas), the present invention can also generate a mainstream with a total temperature of 1700K, matching the total temperature of the incoming flow, the Mach number at the cascade outlet, and the Reynolds number at the cascade inlet of the real design conditions of the aircraft engine / gas turbine turbine.

[0040] Example 1 Example 1 is a test scheme for reproducing the mainstream temperature of the ground test condition of the turbine tester. The shock tube 1 and the transition section 2 are both square pipes, and the blade test piece 3-2 is a flat blade, and its profile is selected as GE-E 3 The profile of the first stage moving blade tip. The layout direction of the inlet section 3-1 is the same as that of the shock tube 1 and the transition section 2. Figure 1 shown x The blade profile of Example 1 has an inlet airflow angle α =29.7°, therefore, the blade arrangement direction and the vertical direction y The angle is α The airflow angle of the blade outlet of this embodiment is β =64.4°, therefore, the angle between the inlet section 3-1 and the outlet section 3-3 is 85.9°=(180°- α - β ).

[0041] The test effect of Example 1 is evaluated by means of computational fluid dynamics simulation below. The numerical calculation domain and boundary conditions of Example 1 are as Figure 2 shown. The length from the leading edge of the middle blade to the outlet of the transition section 2 is 0.1 m, and the length from the trailing edge of the middle blade to the inlet of the Laval nozzle section 4 is 0.15 m. The length of the high-pressure driving section 1-1 of Example 1 L 1 is 5 m, and the lengths of the low-pressure driven section 1-3 and the transition section 2 L 4 are 15 m. The outlet of the Laval nozzle section 4 is a pressure outlet boundary with a static pressure of 1 bar, and the rest are adiabatic no-slip wall surfaces. In terms of the initial conditions, the high-pressure driving section 1-1 is set to air with a pressure of 30 bar and a temperature of 300 K, and the rest are set to air with a pressure of 1 bar and a temperature of 300 K. The test conditions corresponding to this computational fluid dynamics simulation are easy to achieve, and only need to fill the high-pressure driving section 1-1 pipeline with high-pressure air, and the rest are placed in an environment of normal temperature and pressure. After the verification and validation of the computational fluid dynamics simulation, the calculation results of Example 1 are independent of the grid number, time step, spatial difference format, and turbulence model. The specific calculation results are shown below.

[0042] Figure 3 Shows the variation curves of the total pressure at the inlet of the cascade ( P t,i ), the static pressure at the outlet of the cascade ( P s,e ), and its pressure ratio ( P t,i / P s,e ) with the calculation time. The origin of time corresponds to the moment when the diaphragm 1-2 bursts. It can be found that for Example 1, in the interval of 27 - 35 ms, the total pressure at the inlet of the cascade, the static pressure at the outlet, and its pressure ratio are stable. Therefore, the effective test time of Example 1 is about 8 ms. The total pressure at the inlet of the cascade within the effective test time is 1000 ± 15 kPa, and the static pressure at the outlet of the cascade is 630 ± 3 kPa. Therefore, the pressure ratio between the inlet and outlet of the cascade is 1.58 ± 0.02, and the corresponding isentropic Mach number at the outlet of the cascade is 0.82 ± 0.02. In addition, according to Figure 4 the calculation results of the total temperature at the inlet of the cascade (inlet total temperature of the incoming flow) shown, the incoming flow total temperature within the effective test time is 635 ± 10 K, and the Reynolds number is calculated as 1.59×10 5 . Finally, the differences between the incoming flow total temperature, the Mach number at the outlet of the cascade, and the Reynolds number at the inlet of the cascade of Example 1 and the design conditions are within 5%, as shown in Table 1, meeting the requirements of the flow and heat transfer test of the high-temperature turbine cascade.

[0043] Regarding the effect of the periodicity of the cascade flow field, Figure 5Shows the distribution of the cascade exit pressure loss coefficient during the effective test time, that is, (total pressure at the cascade inlet P t,i - total pressure at the cascade exit P t,e ) / (total pressure at the cascade inlet P t,i - static pressure at the cascade exit P s,e ). It can be found that Example 1 has a uniform cascade exit pressure loss coefficient. Especially in the corresponding area of the middle 3 blades, the pressure loss coefficient is 0.09 ± 0.01, indicating that the periodicity of the cascade flow field is at an acceptable level. Further methods to optimize the periodicity of the cascade flow field include: optimizing the profile of the end walls on both sides of the cascade test piece 3-2; sucking the boundary layer of the end walls on both sides of the cascade test piece 3-2; increasing the number of blades in the cascade test piece 3-2, etc.

[0044] Regarding the effectiveness of blade loading, Figure 6 shows the surface pressure ratio curve of the blade during the effective test time, which is defined as the ratio of the static pressure on the blade surface to the total pressure at the cascade inlet. It can be found that the deviation between the surface pressure ratio of the blade in Example 1 and the design value is within ±5%, indicating that the blade loading in Example 1 can effectively reflect the working state of the blade under the ground test conditions of the turbine tester. The Mach number at the throat of the nozzle section 4 of the Laffal nozzle is constantly 1, and the throat area is determined by the mass flow through. For this embodiment, the area of the cascade exit section 3-3 (i.e., the nozzle inlet) is given. According to the area-Mach number relationship of quasi-one-dimensional flow in compressible flow theory, the Mach number of the cascade exit section 3-3 (i.e., the nozzle inlet) is only determined by its area and does not change with the incoming flow conditions of the cascade. Therefore, it is not affected by the change of the shock tube design parameters either.

[0045] Combining all the above results, Example 1 can meet the experimental objectives of the high-temperature turbine cascade, can generate a mainstream with a temperature above 600K, can match the inlet Reynolds number of the cascade, the exit Mach number of the cascade and the total incoming flow temperature under the ground test conditions of the turbine cascade tester, and has the advantages of simple structure, easy operation, low construction, operation and maintenance costs.

[0046] Example 2 Example 2 is a test scheme for reproducing the turbine inlet temperature under the design conditions of an aeroengine / gas turbine real machine. The numerical calculation domain and boundary conditions of the example are as Figure 2 shown. The differences from Example 1 are as follows: First, in terms of the initial conditions, the high-pressure drive section 1-1 in Example 2 is helium with a pressure of 90 bar and a temperature of 300K. Second, the length of the high-pressure drive section 1-1 in Example 2 L 4 is extended to 10 m, and the lengths of the low-pressure driven section 1-3 and the transition section 2 L1 Extended to 40 m, with the goal of extending the effective test time. Methods for further extending the effective test time include: adding a lightweight piston in the low-pressure driven section 1-3 to generate the operating mode of the gun shock tunnel; extending the pipe lengths of the low-pressure driven section 1-3 and the high-pressure driving section 1-1; filling a high molecular weight gas with a low sonic speed at the end of the high-pressure driving section 1-1; inserting a driving plug in the high-pressure driving section 1-1; adjusting the pressure, temperature, and filling gas type of the high-pressure driving section 1-1 and the low-pressure driven section 1-3 based on the stitching conditions.

[0047] Figure 7 Shows the variation curves of the total pressure at the inlet of the cascade ( P t,i ), the static pressure at the outlet of the cascade ( P s,e ), and their pressure ratio ( P t,i / P s,e ) with the calculation time. For Example 2, the effective test time is about 8 ms. The total pressure at the inlet of the cascade within the effective test time ( P t,i ) is 5750 ± 250 kPa, and the static pressure at the outlet of the cascade ( P s,e ) is 3600 ± 100 kPa. Therefore, the pressure ratio at the inlet and outlet of the cascade is 1.58 ± 0.025, and the corresponding isentropic Mach number at the outlet of the cascade is 0.82 ± 0.02. In addition, according to Figure 4 , the total incoming flow temperature within the effective test time of Example 2 is 1700 ± 50 K, which matches the turbine inlet temperature of the aero-engine / gas turbine real engine condition. Methods for further increasing the total incoming flow temperature of the test rig include: adding a heavy piston in the high-pressure driving section 1-1 of the shock tube to generate the free piston shock tunnel mode; increasing the pressure of the high-pressure driving section 1-1; filling the high-pressure driving section 1-1 with "light" gases of low molecular weight such as hydrogen, helium, etc.; heating the pipe of the low-pressure driven section 1-3, etc.

[0048] Other flow characteristics of Example 2 are similar to those of Example 1, such as the periodicity of the cascade flow field, the effectiveness of the blade load, etc., as shown in Figure 5 and Figure 6 . Finally, the differences between the Mach number at the outlet of the cascade, the Reynolds number at the inlet of the cascade, and the total incoming flow temperature of Example 2 and the real engine design conditions are within 5%, as shown in Table 1, meeting the requirements for the flow and heat transfer test of the high-temperature turbine cascade.

[0049] In summary, Example 2 meets the high-temperature turbine cascade test objectives. By using helium to improve the driving ability of the gas in the high-pressure driving section 1-1, when the pressure in the high-pressure driving section 1-1 reaches 90 bar, the total inlet temperature of the incoming flow reaches 1700K, thus matching the total inlet temperature of the turbine under the actual working conditions of an aeroengine / gas turbine.

[0050] Ground test condition Actual machine design condition Example 1 Example 2 Cascade inlet Mach number 0.22 0.22 0.22 0.22 Cascade outlet Mach number 0.82 0.82 0.82 0.82 Cascade inlet Reynolds number 1.59×105 3.76×105 1.59×105 3.76×105 Total inlet temperature (K) 620 1750 635 1700 Total inlet pressure (MPa) 0.216 1.697 1.0 5.750 Chord length (cm) 3.745 3.745 0.834 1.079 Table 1 In the description of the present application, it should be understood that the orientation or positional relationship indicated by the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings, and is only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation to the present application.

[0051] The specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Without conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A high-temperature turbine cascade test device based on a shock tunnel, characterized in that, it includes components connected in sequence: a shock tube (1) for generating a shock wave to heat the test gas through the shock compression effect; a transition section (2) for introducing the high-temperature test gas in the shock tube (1) into the turbine cascade section (3); a turbine cascade section (3) with a cascade test piece (3-2) placed inside; and a Laval nozzle section (4).

2. The high-temperature turbine cascade test device based on a shock tunnel according to claim 1, characterized in that, the shock tube (1) includes a high-pressure drive section (1-1), a diaphragm (1-2), and a low-pressure driven section (1-3); the high-pressure drive section (1-1) is arranged upstream of the low-pressure driven section (1-3), and the high-pressure drive section (1-1) is a pipe for storing high-pressure gas, serving as the drive source for generating high-temperature gas; the diaphragm (1-2) is arranged between the high-pressure drive section (1-1) and the low-pressure driven section (1-3) to isolate the gas in the high-pressure drive section (1-1) from the test gas in the low-pressure driven section (1-3); when the pressure difference between the gas in the high-pressure drive section (1-1) and the test gas in the low-pressure driven section (1-3) reaches a preset value, the diaphragm (1-2) bursts, causing the gas in the high-pressure drive section (1-1) to expand and compress the test gas in the low-pressure driven section (1-3), and generating a shock wave in the test gas to heat the test gas; the low-pressure driven section (1-3) is a pipe for storing test gas.

3. The high-temperature turbine cascade test device based on a shock tunnel according to claim 2, characterized in that, a heavy piston is arranged inside the high-pressure drive section (1-1) to form a free-piston shock tunnel mode; a light piston is arranged inside the low-pressure driven section (1-3) to form a gun tunnel operation mode.

4. The high-temperature turbine cascade test device based on a shock tunnel according to claim 2, characterized in that, the cross-sectional shape of one end of the transition section (2) connected to the shock tube (1) is the same as the cross-sectional shape of the shock tube (1), and the inside of the transition section (2) is filled with gas.

5. The high-temperature turbine cascade test device based on a shock tunnel according to claim 1, characterized in that, the turbine cascade section (3) includes an inlet section (3-1), a cascade test piece (3-2), and an outlet section (3-3) arranged in sequence, and the inside of the turbine cascade section (3) is filled with gas; the inlet section (3-1) is used to introduce the gas in the transition section (2) to generate a uniformly distributed inlet flow for the turbine cascade; the cascade test piece (3-2) is a collection of blade test pieces, including a planar cascade test piece (3-2) or an annular cascade test piece (3-2); the outlet section (3-3) is used to control the outlet flow angle of the cascade test piece (3-2).

6. The high-temperature turbine cascade test device based on a shock tunnel according to claim 5, characterized in that, The included angle between the leading edge line of the cascade test piece (3-2) and the oncoming flow direction upstream thereof is (90° - α), where α is the inlet air flow angle of the cascade test piece (3-2).

7. The high-temperature turbine cascade test device based on a shock tunnel according to claim 5, characterized in that the blade test piece includes a turbine stator blade or a turbine rotor blade.

8. The high-temperature turbine cascade test device based on a shock tunnel according to claim 5, characterized in that the cross-sectional shapes of the inlet section (3-1) and the outlet section (3-3) are both matched with the structure of the cascade test piece (3-2); when the cascade test piece (3-2) is a planar cascade test piece (3-2), the cross-sectional shapes of the inlet section (3-1) and the outlet section (3-3) are both square; when the cascade test piece (3-2) is an annular cascade test piece (3-2), the cross-sectional shapes of the inlet section (3-1) and the outlet section (3-3) are both sector-shaped.

9. The high-temperature turbine cascade test device based on a shock tunnel according to claim 5, characterized in that the cross-sectional shape of the end of the transition section (2) connecting one end of the turbine cascade section (3) is matched with the structure of the cascade test piece (3-2); when the cascade test piece (3-2) is a planar cascade test piece (3-2), the cross-sectional shape of the end of the transition section (2) connecting one end of the turbine cascade section (3) is square; when the cascade test piece (3-2) is an annular cascade test piece (3-2), the cross-sectional shape of the end of the transition section (2) connecting one end of the turbine cascade section (3) is sector-shaped.

10. The high-temperature turbine cascade test device based on a shock tunnel according to claim 1, characterized in that the Laval nozzle section (4) has a structure of first contracting and then expanding. The Laval nozzle section (4) is filled with gas, and the size of the throat of the Laval nozzle section (4) needs to satisfy that the mass flow rate at the throat of the Laval nozzle section (4) is equal to the mass flow rate at the cascade outlet.

Images