CN114279714B - Aeroengine turbine test bed under high altitude and low Reynolds number, simulation method and application - Google Patents
Aeroengine turbine test bed under high altitude and low Reynolds number, simulation method and application Download PDFInfo
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Abstract
The invention provides a simulation method and application of an aircraft engine turbine test system under high altitude and low Reynolds number, wherein the method comprises the steps of obtaining input conditions such as a performance design point and a limit boundary of an air inlet system of an aircraft engine by sorting, and obtaining a design point and a limit boundary of an exhaust system under a working environment of high altitude and low Reynolds number by analyzing; acquiring the injection coefficients of active flow and passive flow of the ejector according to the design principle of the ejector; obtaining preliminary design parameters of the exhaust system through numerical simulation; repeatedly optimizing the initial design parameters according to the limiting conditions, and finally determining the exhaust profile; since the pressure and flow ranges required by the air intake system and the active flow and the passive flow of the ejector are different, the pressure reducing valve needs to be selected based on the pressure reducing valve design standard for JISB3372-1982 compressed air, and finally the opening degree of the pressure reducing valve and the longest test time under the working condition are determined. The turbine test system simulation method provided by the invention can realize rapid design and shaping of the test system, so that the test process is simpler, more convenient and more accurate.
Description
Technical Field
The invention relates to the field of test of aero-engine tests, and discloses a turbine test bed of an aero-engine under high altitude and low Reynolds number, a simulation method and application.
Background
The majority of the working time of an aircraft engine is high altitude, the pressure P of the high altitude is about 28 percent of the atmospheric pressure, and the pressure of ten thousand meters of the high altitude is about 28 percent of the atmospheric pressure, so that the air density rho is reduced, and the formula according to the Reynolds number
(in the formula, U represents the movement speed, mu represents the dynamic viscosity, and l represents the characteristic dimension), the high-altitude flight is under the condition of low Reynolds number relative to the ground, so when the test is carried out on the ground, in order to meet the real high-altitude environment and various extreme conditions, the pressure drop ratio similar to the real condition needs to be ensured, and the wider pressure drop ratio range of low altitude is considered, which also puts higher requirements on the exhaust system of the test bed.
Application No.: CN 2020214967694, publication No.: CN212657059U discloses a multi-nozzle supersonic ejector for high-altitude simulation test bed, which comprises an air inlet pipe, an intake chamber, a mixing pipe, a mixing chamber, a primary pressure expansion chamber, a secondary pressure expansion chamber, an outlet pipe, a plurality of nozzles and a flow deflector. This ejector sets up the nozzle in the passageway department that suction chamber and intake pipe formed, can avoid forming blind chamber at the front end of mixing chamber, compensaties the extra loss of pressure that the multiinjector on the mixing tube arouses, has reduced the start-up degree of difficulty of ejector to a certain extent, has promoted the performance of ejector. Application No.: CN 2020113675994, publication no: CN112611567A, discloses an automatic model selection method for an aircraft engine high altitude simulation test bed regulating valve, which comprises the following steps: analyzing to obtain limit boundaries of the air working environment simulation air inlet system and the air exhaust system; obtaining a working boundary of each regulating valve on the pneumatic branch; determining the specification of an optimal regulating valve according to the working boundary of each regulating valve on the pneumatic branch; according to the working boundary of each regulating valve on the pneumatic branch, based on the strength design EN12516 of the industrial valve shell in the EN standard series, the suitable material and temperature and pressure grade of the regulating valve are obtained through inquiry; designing EN12516 based on the strength of the industrial valve shell according to the temperature and pressure grade and the specification of the optimal regulating valve, and quantitatively determining the calculated pressure Pc, the stress factor S and the corrosion resistance constant c; calculating the minimum shell wall thickness e of the regulating valve meeting the minimum strength reliability requirement of the regulating valve according to a minimum wall thickness calculation formula min (ii) a And (4) selecting the type of the flange of the regulating valve body and the mating flange. Application No.: CN 2019112844259, no: CN110865664A, disclosureThe utility model provides a turbofan engine test bed high altitude cabin front deck pressure quick adjustment device, relates to aerospace ground equipment field, including programmable controller, sensor and hydraulic pressure governing valve, the sensor sets up in test bed high altitude cabin front deck to signal transfer to programmable controller who will monitor, programmable controller control regulation hydraulic pressure governing valve aperture. Further, application No.: CN2019100304264, publication/announcement No.: CN 109668739A/CN 109668739B, disclosing a test platform for integrated research of a multi-duct turbine nozzle, comprising an air inlet section, a turbine section, a culvert, a third duct air inlet system and a nozzle section, and carrying out integrated research on the mixing problem of the turbine, the nozzle and the culvert; an electric heater is arranged in front of the air inlet volute, so that the total temperature difference between the main flow and the bypass fluid is less than or equal to 300K; in order to better simulate the mixing between the inner culvert airflow and the outer culvert airflow, a lobe ejector is arranged at the outlet of the turbine section; the outer culvert inlet is positioned above the ejector and at the spray pipe inlet, and the spray pipe inlet is provided with a mixing section which is approximately equal to straight; the third duct is positioned between the throat and the outlet of the spray pipe, the inlet areas of the external duct and the third duct are adjustable, and the duct ratio is changed within the range of 0-1.0. The above patent only aims at a certain component in the test system, or only proposes a test platform, but how to adjust the corresponding component according to the experiment parameters is not solved.
In order to simulate the working state of the aircraft engine under different high-altitude low Reynolds number environments, the type of each component of the test system needs to be selected to meet the requirement. The type selection reliability directly determines the operation safety of the test bed, the type selection convenience also influences the test progress of the aeroengine, and the important function of the system in the whole test bed is self-evident. As described above, the same aircraft engine often needs to perform complex test subjects under different conditions, and different test subjects need to simulate working environments with wide variation ranges in air flow, pressure drop ratio, total temperature and other indexes. If the model selection is carried out only manually, the workload is large and the efficiency is low, so that the design of the turbine test system simulation method can quickly determine the parameters of each part according to various input conditions, and the matching of the drop-pressure ratio required by the test is an urgent technical problem to be solved.
Disclosure of Invention
The invention aims to provide a simulation method and application of a turbine test system of an aeroengine at high altitude and low Reynolds number, and aims to solve the problem that parameters of all components of the turbine test system cannot be rapidly and accurately selected in the prior art.
In order to solve the technical problem, the invention provides a turbine test bed of an aeroengine under high altitude and low Reynolds number, which has the following technical scheme:
a high-altitude low-Reynolds-number aero-engine turbine test bed comprises an air storage tank, an air inlet system, an active flow input end and a passive flow input end of an ejector, a pressure reducing valve, an ejector body and an exhaust volute; it is characterized in that: the three outlets of the gas storage tank are respectively connected with the gas inlet system and the active flow input end and the passive flow input end of the ejector, and a pressure reducing valve is arranged on each path; the gas in the gas storage tank enters the aero-engine to burn and do work after passing through the gas inlet system, and then is discharged into the exhaust volute; the active flow and the passive flow input by the ejector body are mixed in the ejector body and discharged into the exhaust volute after reaching a preset exhaust pressure, so that different exhaust pressures of the aircraft engine in a high-altitude low Reynolds number environment are simulated.
The invention also discloses a method for testing the turbine of the aircraft engine at high altitude and low Reynolds number, which is characterized by comprising the following steps: the method comprises the following steps:
step 1: determining a design point and a limit boundary of an exhaust system in a high-altitude low Reynolds number working environment;
and 2, step: obtaining the injection coefficients of the active flow and the passive flow of the ejector at each design point;
and step 3: carrying out numerical simulation according to the exhaust and injection coefficient parameters of each design point to obtain the initial design parameters of an exhaust system, and carrying out corresponding modification and optimization on the size of the ejector body according to the numerical simulation result, wherein the inner diameters of an inlet and an outlet of a spray pipe are mainly used for finally determining the most matched ejector body profile;
and 4, step 4: according to different pressure and flow ranges required by an air inlet system and active flow and passive flow of an ejector, a pressure reducing valve with a proper specification is selected based on the design standard of the pressure reducing valve for JISB3372-1982 compressed air;
and 5: and determining the opening degree of the pressure reducing valve and the longest test time under the working condition according to the specific working condition.
In addition, the invention also discloses a simulation experiment method for the turbine test bed of the aero-engine under high altitude and low Reynolds number, which is applied to the design of the aero-engine.
Advantageous effects
The invention realizes rapid and accurate model selection of the turbine test system; compared with the complex model selection analysis work which is completely dependent on testers from beginning to end and is associated with each other, the model selection tool for the turbine test system based on the Labview environment changes the working mode of the turbine test system of the current simulation aeroengine high-altitude low Reynolds number test bed, greatly improves the working efficiency, and has a certain popularization effect in relevant pressure and flow regulation application scenes.
Drawings
FIG. 1 is a schematic view of a pneumatic connection of a test stand;
FIG. 2 is a characteristic curve of an axial compressor;
FIG. 3 is a flow chart of injector injection coefficient calculation;
FIG. 4 is a turbine test simulation methodology logic diagram;
in fig. 1: the cross section (1) of the throat part of the ejector and the cross section (2) of the outlet of the ejector.
Detailed Description
The following detailed description of the present invention is provided in connection with the accompanying drawings and examples. The following examples are intended to illustrate the invention, but are not intended to limit the scope of the invention.
Referring to fig. 1, the test bed for the turbine of the aircraft engine at high altitude and low reynolds number provided by the embodiment of the invention comprises: the ejector mainly comprises an air inlet system, an ejector active flow input port, an ejector passive flow input port, a pressure reducing valve and an ejector body; the test gas is provided by high-pressure gas in the gas storage tank, and because the required pressure of the gas inlet system, the active flow and the passive flow of the ejector is different, different pressure reducing valves are required, and different test working conditions require different flow rates, so the opening degree of the pressure reducing valves needs to be correspondingly adjusted. The air storage tank is divided into three paths of outlets, the three paths of outlets are respectively connected with an air inlet system, an active flow input port and a passive flow input port of an ejector, a pressure reducing valve is arranged on each path, air enters the aircraft engine to be combusted and do work after passing through the air inlet system, then the air is discharged into an exhaust volute, the active flow and the passive flow are mixed in the ejector to reach preset exhaust pressure and are also discharged into the exhaust volute, and different exhaust pressures of the aircraft engine in a high-altitude low Reynolds number environment are simulated.
On the basis of the test bed, the simulation method of the aeroengine turbine test system under the conditions of high altitude and low Reynolds number is provided, the logic of the simulation method is shown in figure 4, and the simulation method comprises the following steps:
step 1: determining the working point (design point) and limit boundary of the exhaust system under the high-altitude low Reynolds number working environment: the method comprises the following steps of sorting and determining a typical high-altitude low Reynolds number working environment of a certain type of aircraft engine, which is a main test object of a test bed, and a limit boundary and a typical working point (design point) of the performance of an air inlet system of the aircraft engine, as a basis of subsequent model selection analysis, and finally obtaining the design point and the limit boundary of an exhaust system through calculation, wherein the method specifically comprises the following steps: inquiring an international standard atmosphere meter according to the maximum rising limit of a certain type of aircraft engine to obtain the limit boundaries of density rho, inlet static pressure P and static temperature T, and selecting the takeoff speed, the maximum/minimum horizontal flight speed, the cruise speed, the force application speed, the rotating speed at the moment and the inlet total pressure P at the moment of the design point of the engine 0 And total temperature T 0 According to the formula
Calculating out a total pressure P therein 0 And total temperature T 0 Is the limit boundary; inlet flow m 1 From m 1 Calculation of = ρ vA based on axial flow compressor characteristic curve (as shown in FIG. 2), since the rotation speed and flow rate are known, the pressure ratio can be obtained by looking up the characteristic curve
P 0 Is an inlet assemblyPressure, as already determined above, P 2 Is the total outlet pressure, thereby obtaining the total outlet pressure P 2 And furthermore according to the continuous equation m 1 ＝m 2 +m f Flow rate of outlet m 2 Is equal to the inlet flow m 1 Plus fuel consumption m f To obtain the outlet flow m 2 The outlet flow velocity can be according to v = m 2 The/rho A is calculated, the outlet static temperature is equal to the inlet static temperature, and the total outlet temperature can be calculated according to
Where v is the flight velocity, c p And k is the specific heat capacity at constant pressure, ρ is the density, and A is the area, so that the design points and limit boundaries of the total temperature, the total pressure and the flow of the corresponding exhaust system are obtained.
Step 2: acquiring the injection coefficients of the active flow and the passive flow of the ejector at each working point (design point);
a summary analysis of the operating environment and intake system performance was performed, and as shown in table 1 (table 1 is only a small example), the typical operating point of the intake system of the test subject determined the extreme limits of the exhaust system. For example, the limit range of the total temperature and the total pressure of the exhaust system is determined according to the pressure drop ratio, the total pressure of the inlet air and the total temperature, and the flow rate of the inlet air determines the size of the pipe diameter of the exhaust system and the opening degree of the pressure regulating valve, including the limit rotating speed range, the limit pressure regulating range and the limit flow regulating range. Knowing the exhaust pressure required to be provided by the ejector outlet, the design objective is to obtain the injection coefficients of the active flow and the passive flow of the ejector by the flow chart shown in fig. 3 according to the design principle of the ejector. The concrete steps are as shown in figure 3, firstly inputting the inlet parameters of the ejector, namely pressure and flow, then giving an initial value of 0 to the injection coefficient u, and then passing through the following formulas (1) and (2)
Table 1 is as follows:
| parameter(s) | Unit of | Design Point 1 | Design Point 2 | Design Point 3 | Design Point 4 | Design Point 5 |
| Physical speed of rotation | rpm | 17000 | 6000 | 15205.3 | 12000 | 10820 |
| Inlet total pressure | kPa | 400 | 400 | 250 | 280 | 350 |
| Total temperature of inlet | K | 1000 | 500 | 800 | 900 | 700 |
| Inlet flow rate | kg/s | 7.12 | 6.31 | 4.975 | 5.815 | 5.201 |
| Power output | MW | 1.6108 | 0.925 | 1.023 | 1.254 | 1.420 |
| Total pressure of exhaust | kPa | 144.95 | 354.61 | 90.58 | 80.12 | 251.52 |
| Total temperature of exhaust gas | K | 806.2 | 486.3 | 590.2 | 540.5 | 650.8 |
| Exhaust flow rate | kg/s | 7.82 | 6.63 | 5.38 | 5.98 | 5.65 |
And step 3: carrying out numerical simulation according to the exhaust and injection coefficient parameters of each design point to obtain the initial design parameters of an exhaust system, and carrying out corresponding modification and optimization on the size of the ejector body according to the numerical simulation result, wherein the inner diameters of an inlet and an outlet of a spray pipe are mainly used for finally determining the most matched ejector body profile; because the theoretical calculation has a certain deviation from the actual situation, numerical simulation is carried out by adopting a realizable k-epsilon turbulence model according to input design points, namely conditions such as total temperature and pressure of inlet air, injection coefficient and the like as boundary conditions,
in the above equation: g k Showing turbulent kinetic energy generation due to mean velocity gradient, gb is used to show turbulent kinetic energy generation due to buoyancy effect, β is the coefficient of thermal expansion, Y M Representing the effect of compressible turbulent pulsating expansion on the overall dissipation ratio, C 1 、C 2 And C 1ε Is a coefficient; sigma k And σ ε Turbulence prandtl number which is the turbulence energy and its dissipation ratio; s is the vortex viscosity coefficient.
Obtaining the pressure of the exhaust system, taking the pressure of a design point of the exhaust system as a convergence condition, wherein the convergence condition is that the difference between the pressure of the exhaust system obtained by simulation and the pressure of the design point does not exceed +/-2% of the pressure value of the design point, namely
If not, convergence is considered; if the contraction degree is less than 2%, the area of the throat section (1) is increased (the radius is increased by 1 mm), and then numerical simulation is carried out until the throat section is converged; if the value is more than 2 percent, the contraction degree is too small, the area of the outlet section (2) is increased (the radius is increased by 1 mm), and then numerical simulation is carried out until convergence; and finally determining the most matched ejector body profile through the adjustment of the steps.
And 4, step 4: since the gas source is high-pressure gas in the gas tank, the pressure and flow ranges required by the gas inlet system and the ejector active flow and passive flow are different, a proper type of the pressure reducing valve needs to be selected based on the design standard (shown in table 2) of the pressure reducing valve for the JISB3372-1982 compressed air. As shown in Table 2, the nominal pressure of the pressure reducing valve model is required to be higher than the gas pressure of each path, and the nominal path D is required to flow the required flow, i.e. the required flow
ρ is the density and v is the flow velocity, and the model with the lightest weight is selected on the basis of the density and the flow velocity. TABLE 2 such asThe following:
and 5: according to the corresponding relation between the flow and the opening of the pressure reducing valve (which is the product performance given by a pressure reducing valve manufacturer) and a compressed gas differential form continuous equation
Wherein p is * Is the total outlet pressure, T * The total outlet temperature, A the outlet area, K, R the gas constant and v the gas specific volume, and the opening degree of each pressure reducing valve under the working condition and the longest test time t capable of keeping the working condition can be calculated according to the formula.
In conclusion, the method for simulating the turbine test system of the aircraft engine under the high altitude and low Reynolds number and the application thereof provided by the invention have the advantages that the turbine test system is comprehensively subjected to model selection analysis based on the high altitude and low Reynolds number working environment and the test boundary of the aircraft engine, and the high altitude simulation test progress of the engine with the important model under the research is ensured. In Labview environment, the turbine test system full-flow automatic model selection software is designed based on the design standard of the pressure reducing valve for JISB3372-1982 compressed air, the efficiency and the accuracy of the model selection work of the turbine test system are greatly improved, designers are liberated from the complicated model selection work, and the construction progress of other similar test benches in China is facilitated.
The foregoing shows and describes the general principles, principal features, and advantages of the invention. It will be understood by those skilled in the art that the present invention is not limited to the embodiments described above, which are merely illustrative of the principles of the invention, but that various changes and modifications may be made without departing from the spirit and scope of the invention, which fall within the scope of the invention as claimed. The scope of the invention is defined by the appended claims and equivalents thereof.
Claims (4)
1. The simulation method of the aircraft engine turbine test system under the high altitude and low Reynolds number comprises an aircraft engine turbine test bed under the high altitude and low Reynolds number, wherein the test bed comprises an air storage tank, an air inlet system, an active flow input end and a passive flow input end of an ejector, a pressure reducing valve, an ejector body and an exhaust volute; it is characterized in that: the air storage tank is divided into three paths of outlets which are respectively connected with an air inlet system, and an active flow input end and a passive flow input end of the ejector, and a pressure reducing valve is arranged on each path; the gas in the gas storage tank enters the aircraft engine to burn and apply work after passing through the gas inlet system, and then is discharged into the exhaust volute; after the active flow and the passive flow input by the ejector body are mixed in the ejector body and reach a preset exhaust pressure, the mixed flow is also discharged into an exhaust volute, and different exhaust pressures of the aircraft engine in a high-altitude low Reynolds number environment are simulated; it is characterized in that: the method comprises the following steps:
step 1: determining a design point and a limit boundary of an exhaust system in a high-altitude low Reynolds number working environment;
step 2: obtaining the injection coefficients of the active flow and the passive flow of the ejector at each design point;
and step 3: carrying out numerical simulation according to the exhaust and injection coefficient parameters of each design point to obtain initial design parameters of an exhaust system, carrying out corresponding modification and optimization on the size of the ejector body according to the numerical simulation result, and finally determining the most matched ejector body outline; repeatedly modifying and optimizing the initial design parameters according to the initial design parameters and the limiting conditions in the step 1-2, and finally determining the exhaust profile, wherein the method specifically comprises the following steps: because theoretical calculation has certain deviation from actual conditions, numerical simulation is carried out according to input design points, namely air inlet conditions and injection coefficient conditions, the input design points are used as known conditions to obtain the pressure of an exhaust system, the pressure of the design point of the exhaust system is used as a convergence condition, iterative calculation analysis is carried out by adjusting the areas of the throat section (1) and the outlet section (2) of the ejector, and the most matched ejector body profile is determined;
and 4, step 4: according to different pressure and flow ranges required by an air inlet system and active flow and passive flow of an ejector respectively, selecting a pressure reducing valve with a proper specification based on the design standard of the pressure reducing valve for JISB3372-1982 compressed air;
and 5: determining the opening degree of the pressure reducing valve and the longest test time under the working condition according to the specific working condition;
the step 1 further comprises the following steps: inquiring an international standard atmosphere meter according to the maximum rising limit of a certain type of aircraft engine to obtain the limit boundaries of density rho, inlet static pressure P and static temperature T, and selecting the takeoff speed, the maximum/minimum horizontal flight speed, the cruise speed, the force application speed, the rotating speed at the moment and the inlet total pressure P at the moment of the design point of the engine 0 And total temperature T 0 According to the formula
Calculating out a total pressure P therein 0 And total temperature T 0 Is the limit boundary; inlet flow m 1 From m 1 Calculation of = ρ vA, based on axial-flow compressor characteristic curve, since the rotation speed and flow are known, the pressure ratio can be obtained by looking up the characteristic curve
P 0 For the total inlet pressure, which has already been determined above, P 2 Is the total outlet pressure, thereby obtaining the total outlet pressure P 2 And furthermore according to the continuous equation m 1 ＝m 2 +m f Flow rate of outlet m 2 Is equal to the inlet flow m 1 Plus fuel consumption m f To obtain the outlet flow m 2 Outlet flow velocity according to v = m 2 The/rho A is calculated, the outlet static temperature is equal to the inlet static temperature, and the total outlet temperature can be calculated according to
Where v is the flying speed, c p And k is the specific heat capacity at constant pressure, ρ is the density, and A is the area, so that the design points and limit boundaries of the total temperature, total pressure and flow of the corresponding exhaust system are obtained.
2. The simulation method of the high-altitude low-Reynolds-number aircraft engine turbine test system according to claim 1, wherein the simulation method comprises the following steps: the step 2 further comprises the following steps: and (3) obtaining exhaust pressure through the step (1), and determining the injection coefficients of the active flow and the passive flow of the ejector by combining the design principle of the ejector.
3. The simulation method for the high-altitude low-Reynolds-number aircraft engine turbine test system according to claim 1, wherein: the step 4 further includes the following steps of selecting a pressure reducing valve with a proper specification, specifically including: the type of the pressure reducing valve is determined based on the design standard of the pressure reducing valve for compressed air of JISB3372-1982 according to different pressure and flow ranges required by the air inlet system and the active flow and the passive flow of the ejector respectively.
4. The simulation method of the high-altitude low-Reynolds-number aircraft engine turbine test system according to claim 1, wherein the simulation method comprises the following steps: the step 5 further includes determining the opening of the pressure reducing valve and the longest test time under the working condition, and specifically includes: according to the corresponding relation between the flow and the opening of the reducing valve and a differential form continuous equation of the compressed gas; the opening degree of each pressure reducing valve and the longest test time for maintaining the working condition are determined.
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| CN115494892B (en) * | 2022-08-31 | 2023-06-30 | 中国航发四川燃气涡轮研究院 | Decoupling control method for air inlet environment simulation system of high-altitude simulation test bed |
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| CN116150915A (en) * | 2023-04-03 | 2023-05-23 | 中国航发四川燃气涡轮研究院 | Setting method of adjusting device of high-altitude simulation test bed of large and medium-sized engine |
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