CN114659745B - Transition state performance test system and experimental method for turbine component of aeroengine - Google Patents
Transition state performance test system and experimental method for turbine component of aeroengine Download PDFInfo
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- CN114659745B CN114659745B CN202210269668.0A CN202210269668A CN114659745B CN 114659745 B CN114659745 B CN 114659745B CN 202210269668 A CN202210269668 A CN 202210269668A CN 114659745 B CN114659745 B CN 114659745B
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- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
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Abstract
A transition state performance test system of an aeroengine turbine component comprises a heater, wherein an inlet of the heater is connected with an air source, and an outlet of the heater is divided into two paths which are a test air path and an adjusting bypass respectively; along the air flow direction, a quick response pneumatic valve and a test measurement model are arranged on the test air path; the test measurement model is an aeroengine turbine component or a simulated aeroengine turbine component; the power of the heater is regulated to enable the gas in the test gas circuit to show a temperature rise change process or a temperature reduction change process, so that the time-varying temperature transition state working condition simulation is realized; the pressure transition state working condition simulation is realized by controlling the change rate of the quick response pneumatic valve in the process from on to off or from off to on, further adjusting the air inlet pressure change process at the upstream of the test measurement model, simulating the pressure condition of the inlet of the turbine component under the actual working condition of starting and stopping, accelerating or decelerating the aeroengine. The invention also provides a corresponding test method, which provides references for component improvement and performance improvement under real conditions.
Description
Technical Field
The invention belongs to the technical field of impeller machinery, and particularly relates to a transition state performance test system and an experimental method for a turbine component of an aero-engine.
Background
Various components within an aircraft engine tend to operate in a severe environment at high temperatures and pressures. In the running process, besides steady-state running, the aeroengine can also experience various unstable working conditions such as start-stop, acceleration, deceleration and the like, under the unstable working conditions, the running states of related components (blade grids, disc cavities, sealing, exhaust systems and the like) can obviously fluctuate, on one hand, the unstable fluctuation can reduce the running efficiency and increase the fuel consumption, and on the other hand, the potential safety hazard of the aeroengine running can also be increased due to the complex structure. Therefore, the flow field change rule and the aerodynamic characteristics of the relevant parts under various operation conditions are deeply known, the relevant parts are improved based on the flow field change rule and the aerodynamic characteristics, and the stability and the operation efficiency of the aeroengine under the variable operation conditions are improved.
In the field of aeroengine research, an experimental method and a numerical simulation method are two important research methods, wherein the experimental method does not need to simplify boundary conditions, and the method has the characteristics of strong intuitiveness and high reliability. At present, most of aeroengine turbine component performance test platforms for mechanical research at home and abroad are designed and built based on steady-state working conditions, and the aerodynamic performance change of real components under variable working conditions is often ignored.
Disclosure of Invention
In order to overcome the defects of the prior art, the invention provides a transition state performance test system and an experimental method for an aeroengine turbine component under a transition state, and the operation characteristics of the corresponding turbine component under an unstable working condition are further and truly explored through transition state test research, so that references are provided for optimizing and improving the aeroengine, and the operation efficiency of the aeroengine is further improved.
In order to achieve the above purpose, the technical scheme adopted by the invention is as follows:
a transition state performance test system of an aeroengine turbine component comprises a heater, wherein an inlet of the heater is connected with an air source, and an outlet of the heater is divided into two paths which are a test air path and an adjusting bypass respectively;
along the air flow direction, the test air circuit is provided with a quick response pneumatic valve and a test measurement model; the test measurement model is an aeroengine turbine component or a simulated aeroengine turbine component;
the power of the heater is regulated to enable the gas in the test gas circuit to show a temperature rise change process or a temperature reduction change process, so that the time-varying temperature transition state working condition simulation is realized;
the pressure transition state working condition simulation is realized by controlling the change rate of the quick response pneumatic valve in the process from on to off or from off to on, further adjusting the air inlet pressure change process at the upstream of the test measurement model, simulating the pressure condition of the inlet of the turbine component under the actual working condition of starting and stopping, accelerating or decelerating the aeroengine.
In one embodiment, the first back pressure valve is arranged on the regulating bypass to regulate the initial pressure of the test gas circuit, and meanwhile, the gas collecting effect when the quick response pneumatic valve is closed is avoided; and a back pressure valve II is arranged on the test gas path and positioned at the downstream of the test measurement model so as to adjust the outlet back pressure of the test measurement model.
In one embodiment, the fast response pneumatic valve time range is 1-10 seconds.
In one embodiment, the aeroengine turbine component is a cascade, disk cavity, seal, or exhaust system.
In one embodiment, a dynamic pressure sensor and a temperature sensor are arranged on the test gas path between the quick response pneumatic valve and the test measurement model to monitor whether the pressure change curve and the gas temperature meet target requirements.
In one embodiment, the test measurement model is provided with an NI acquisition system, and the NI acquisition system is used for acquiring and recording the changes of pressure, speed and temperature of the test measurement model along with time in a transitional working condition and corresponds to an upstream pressure fluctuation curve of the test measurement model caused by the action of the quick response pneumatic valve.
In one embodiment, the valve action signal of the quick response pneumatic valve is triggered synchronously with the acquisition of the physical quantity signal, and the pulse signal is sent out while the valve action of the quick response pneumatic valve is triggered, so that the NI acquisition system is triggered to start to acquire and record data.
In one embodiment, the inlet pressure upstream of the pilot measurement model varies linearly or non-linearly.
In one embodiment, when the turbine component is a rotating component, the rotating speed of the motor is gradually increased or gradually decreased by controlling a power transformation cabinet of the motor, so that the variable rotating speed working condition simulation is realized; or when a vibration experiment is carried out, exciting force is applied to the rotating part from the outside by using the vibration exciter, and the vibration exciting force is gradually increased or gradually reduced by adjusting the vibration exciting force, so that the vibration exciting force changing working condition simulation is realized.
The invention also provides an experimental method of the transition state performance test system based on the turbine component of the aeroengine, which comprises the following steps:
the power of the heater is regulated to enable the gas in the test gas circuit to show a temperature rise change process or a temperature reduction change process, so that an experiment is carried out, and the time-varying temperature transition state working condition simulation is realized;
the initial pressure of the test gas circuit is regulated to the initial target pressure of the variable working condition through regulating the regulating bypass;
the pressure condition of the turbine component inlet under the actual working conditions of starting, stopping, accelerating or decelerating of the aeroengine is simulated by controlling the change rate of the quick response pneumatic valve in the process from on to off or from off to on, so that the pressure transition state working condition simulation is realized.
Compared with the prior art, the invention has the beneficial effects that:
in the current related test research about turbine components of aeroengines, a plurality of test researches under steady-state working conditions are carried out, but in a real running environment, related equipment can undergo transition state running processes such as deceleration, acceleration, start-stop and the like. During these processes, the aerodynamic parameters inside the relevant components change drastically and the operating conditions fluctuate. This affects the efficiency of operation of the devices, and on the other hand, may increase the safety hazards of operation of the devices due to the operation of the devices in high temperature and high pressure environments. The transition state test system can heat the gas in the pipeline to different temperatures by adjusting the power of the heater, or maintain the gas in the pipeline in a heating or cooling process in a period of time; the programmable control quick response pneumatic valve is arranged at the upstream of the test measurement model, and the action rule of the quick response pneumatic valve is controlled through the related program, so that pressure fields with different change rates can be formed at the downstream of the quick response pneumatic valve, and the variable working condition air inlet condition of related equipment under the real running condition can be simulated. Test data obtained by the multi-physical-quantity synchronous measurement system can be used for deeply knowing the operation characteristics of related equipment under the transitional working condition. The method is beneficial to deepening the understanding of the operation state of the turbine component of the aeroengine under the transition state working condition, provides reference for the improvement scheme of related equipment, and further is beneficial to improving the operation efficiency of the related equipment and enhancing the operation stability of the related equipment.
Drawings
FIG. 1 is a schematic diagram of a transitional performance testing system according to the present invention.
FIG. 2 is a schematic diagram of the pulse signal of the present invention in response to the actuation of the pneumatic valve.
FIG. 3 is a schematic diagram of a transition state operating mode pressure simulation of the present invention.
FIG. 4 is a schematic diagram of an operation interface of the synchronous measurement system for multiple physical quantities according to the present invention.
FIG. 5 is a schematic diagram of the test measurement results of the transient state operation mode of the present invention.
Detailed Description
Embodiments of the present invention will be described in detail below with reference to the accompanying drawings and examples.
Developing a transition state test is a necessary means to explore the true operating state of an aeroengine turbine component under variable operating conditions. Based on the system, the invention provides a transition state performance test system of an aeroengine turbine component and a test method based on the system.
As shown in fig. 1, the transition state performance test system mainly comprises an air source 1, a heater 3, a test air path 4, a regulating bypass 8, a quick response pneumatic valve 5, a test measurement model 6 and the like.
The air source 1 is connected with the heater 3 to heat the gas for test, and the outlet of the heater 3 is divided into two paths, namely a test air path 4 and an adjusting bypass 8. The gas can be provided by a compressor and stabilized by a large gas storage tank, and a dryer 2 can be arranged between the gas source 1 and the heater 3 to dry the gas for test in order to optimize the test effect. The heater 3 may heat the gas to different temperatures.
In the invention, the power of the heater 3 is regulated to enable the gas in the test gas circuit 4 to present a temperature rise change process or a temperature reduction change process, so that the time-varying temperature transition state working condition simulation can be realized.
The quick response pneumatic valve 5 and the test measurement model 6 are disposed on the test gas path 4, and the quick response pneumatic valve 5 is located upstream of the test measurement model 6 in the gas flow direction.
The pressure transition state working condition simulation can be realized by controlling the change rate of the quick response pneumatic valve 5 in the process from on to off or from off to on, and further adjusting the air inlet pressure change process at the upstream of the test measurement model 6, so as to simulate the pressure condition (namely the air inlet condition) of the inlet of the turbine component under the actual working condition of starting, stopping, accelerating or decelerating the aeroengine. And carrying out a corresponding test under the condition of the target air inlet pressure, and then carrying out test exploration of the required transition state working condition.
The present invention provides for the purpose of regulating the initial pressure of the test gas circuit 4 while avoiding the gas collection effect when the fast response pneumatic valve 5 is closed. Specifically, the first back pressure valve 9 may be provided on the regulation bypass 8, and the above object can be achieved by regulating the first back pressure valve 9.
The invention can also be provided with the second back pressure valve 7 on the test gas path 4, and the second back pressure valve 7 is arranged at the downstream of the test measurement model 6 and is used for adjusting the back pressure of the outlet of the test measurement model 6.
The heater 3, the quick response pneumatic valve 5 and the controllable bypass 8 can also be simultaneously adjusted, so that the combination of transition state working conditions such as time-varying temperature, pressure and the like is realized.
The valve change time range of the quick response pneumatic valve 5 is adjustable within 1-10 seconds.
The test and measurement model 6 of the invention is an aero-engine turbine component or a simulated aero-engine turbine component, such as a seal component, a cascade component, a disk cavity, a seal or exhaust system, etc. The test measurement model 6 can be replaced according to the study requirements.
In the embodiment of the invention, the quick response pneumatic valve 5 is controlled by a pneumatic valve control box 13 connected with the quick response pneumatic valve 5, the pneumatic valve control box 13 adopts a programmable control mode, and the change rate of the valve from on to off or from off to on can be controlled through a corresponding program, so that a specific target pressure curve can be formed downstream of the quick response pneumatic valve 5. Wherein the control program can be outputted from the computer terminal 14. By inputting different control programs, the intake air pressure upstream of the test measurement model 6 can be varied in different target pressure curves, including linear (e.g., linear function curves) and non-linear (e.g., exponential function curves) processes.
On the test gas circuit 4, a dynamic pressure sensor 10 and a temperature sensor 11 with high response frequency are arranged between a quick response pneumatic valve 5 and a test measurement model 6, the dynamic pressure sensor 10 and the temperature sensor 11 have the characteristic of quick response, the dynamic pressure sensor 10 is used for monitoring a pressure change curve at the upstream of the test measurement model 6 and verifying whether the pressure curve meets the transition state test requirement, and the temperature sensor 11 is used for monitoring whether the gas temperature meets the test requirement.
Corresponding measuring probes such as pressure, temperature and the like can be arranged at corresponding positions on the test measuring model 6 according to requirements, signals of the various probes are collected by a set of NI collecting system 12, and data processing is carried out at a computer terminal 14 according to corresponding programs. The NI acquisition system 12 acquires and records the changes of pressure, speed, temperature and the like of the test measurement model 6 along with time in the transient state working condition, and corresponds to an upstream pressure fluctuation curve of the test measurement model 6 caused by the action of the quick response pneumatic valve 5.
In the present invention, NI in the NI acquisition system 12, referred to as national instruments corporation (National Instruments), is a data acquisition device commonly used in the art.
For example, the NI acquisition system 12 and the pneumatic valve control box 13 may be integrated using a LABVIEW program to achieve multiple physical quantity simultaneous acquisition. Thus, a multi-physical quantity synchronous measurement system is formed based on the NI acquisition system 12. At this time, the valve action signal of the fast response pneumatic valve 5 and the collection of the physical quantity signal (such as pressure) can be synchronously triggered, the fast response pneumatic valve 5 sends out a pulse signal while the valve action of the fast response pneumatic valve 5, the NI collection system 12 is triggered to start to collect and record data, and the frequency of the pulse signal can be 50HZ.
Under the above system structure, it is easy to understand that the transient state working condition is not limited to the change of physical quantities such as inlet pressure, temperature, etc., and when the turbine component is a rotating component, the transient state working condition process such as the variable rotation speed, the variable excitation force, etc. of the test measurement model 6 can be also included. Specifically, if the motor is a rotating part, the rotating speed of the motor can be gradually increased or gradually decreased by controlling the power transformation cabinet of the motor, so that the variable-rotating-speed working condition simulation is realized. When a vibration experiment (or a rotor dynamics experiment) is performed, a set of vibration exciter is installed, and an exciting force is applied to the rotating part from the outside by using the vibration exciter, so that the vibration experiment is performed. The exciting force can be regulated to be gradually increased or gradually reduced through software, so that the variable exciting force working condition simulation is realized.
The test procedure of the invention comprises:
the gas of the gas source 1 enters the heater 3 after passing through the dryer 2, and part of the gas heated by the heater 3 enters the test measurement model 6 through the test gas circuit 4, and the other part of the gas enters the regulation bypass 8.
The power of the heater 3 is regulated to enable the gas in the test gas circuit 4 to show a temperature rise change process or a temperature reduction change process, so that an experiment is carried out, and the time-varying temperature transition state working condition simulation is realized;
the initial pressure of the test gas circuit 4 is regulated to the initial target pressure of the variable working condition by regulating the regulating bypass 8;
the pressure condition of the turbine component inlet under the real working condition of starting and stopping, accelerating or decelerating of the aeroengine is simulated by controlling the change rate of the quick response pneumatic valve 5 from on to off or from off to on, further adjusting the air inlet pressure change process at the upstream of the test measurement model 6, and performing an experiment to realize the pressure transition state working condition simulation.
That is, the invention can heat the gas in the test gas path 4 to different temperatures by adjusting the power of the heater 3, or subject the gas in the test gas path 4 to a heating or cooling process in a period of time; by adjusting the valve action rule of the fast response pneumatic valve 5, the pressure change process of the inlet of the turbine component (blade grid, disc cavity, sealing, exhaust system and the like) can be simulated when the real engine runs in a variable working condition transition state (take-off, acceleration, landing, tactical action and the like) at the downstream of the fast response pneumatic valve, namely at the upstream of the test measurement model 6.
In one embodiment of the invention, when a transition state test is performed, the air source 1 can be provided by a large screw compressor, air flows out of the screw compressor and then enters the air storage tank, the air storage tank has the function of stabilizing pressure, then flows into the heater 3 through the pipe after passing through the dryer 2, and the air in the pipe can be heated to different initial temperatures by adjusting the power of the heater 3 and then flows into the programmable control quick response pneumatic valve 5, and the valve of the programmable control quick response pneumatic valve 5 is kept in a full-open state. In the piping downstream of the fast response pneumatic valve 5, a steady flow honeycomb is installed for providing a uniform and stable experimental air flow downstream.
In one embodiment of the present invention, as shown in fig. 2, the programmable fast response pneumatic valve 5 is independently supplied by the air pump through the pipeline, the actuation time of the fast response pneumatic valve 5 (the time from on to off or from off to on) can be adjusted within the range of 1-10s, and when the fast response pneumatic valve 5 is in the moment of starting the command actuation, the valve will send a pulse voltage with the frequency of 50Hz, and the pulse voltage can trigger the NI acquisition system 12 to acquire multiple physical quantity signals.
In one embodiment of the invention, the target static pressure curve must be reduced from 3Kpa to 0.5Kpa in a conic within 1.5 seconds, as shown in fig. 3. And adjusting the valve opening of the back pressure regulating valve 9 of the bypass 8 according to the initial target pressure value required by the transition state working condition, so that the initial pressure of the test measurement model 6 is similar to the target static pressure. And then inputting a corresponding control program of the quick response pneumatic valve according to the target pressure curve, wherein the program can adjust the action rate of the valve from on to off, and finely adjust and control relevant parameters of the quick response pneumatic valve 5 program until the pressure change process of the high-frequency dynamic pressure sensor 10 is basically consistent with the target pressure curve change process in a transitional working condition. As shown in FIG. 3, the relative time "0" corresponds to the valve action initial time of the quick response pneumatic valve 5, and the pressure curve within the time range of 0.5-2 seconds of the valve action time is taken as the actual change pressure of the transitional working condition. In the time period, the static pressure of the measuring point is changed from 2.862KPa to 0.503KPa, the target static pressure curve is changed from 3KPa to 0.5KPa, the deviation between the actual static pressure change rate and the target curve change rate is 4.9%, the maximum relative deviation in the static pressure change process is 4.6%, and the overall static pressure simulation error is less than 5%, so that the transient state test requirement can be met.
In one embodiment of the invention, after the transient state pressure change working condition is debugged, the quick response pneumatic valve 5 is kept normally open, the air flow of the pipeline 4 to be tested is stable, and the arrangement of the measuring instrument related to the test measuring model 6 is confirmed. The computer terminal 14 sends a valve action command, the pneumatic valve 5 is quickly responded to start to act, and meanwhile, a pulse signal is sent, and the pulse signal triggers the NI acquisition system 12 to stop acting until the valve stops acting. The LABVIEW system of the computer terminal 14 records and stores the relevant data after the rapid response pneumatic valve 5 starts to act, and the data acquisition under the transient state working condition can be obtained after processing according to the relevant program.
In one embodiment of the present invention, as shown in fig. 4, an operation interface of the multi-physical-quantity synchronous acquisition system is written for LABVIEW during the test.
In one embodiment of the present invention, as shown in fig. 5, a schematic diagram of the data recorded and stored by the multi-physical-quantity synchronous acquisition system after processing is shown, where the diagram shows the characteristics of total, static pressure and speed change of a certain position of the test measurement model 6 under the transient state working condition.
Therefore, the invention is helpful to understand the real aerodynamic characteristics of the test model under the variable working condition by exploring the multi-physical-quantity change characteristics of the test measurement model under the transition state working condition, and provides references for the improvement and performance improvement of related equipment under the real condition.
Claims (7)
1. The transition state performance test system of the turbine component of the aeroengine is characterized by comprising a heater (3), wherein an inlet of the heater (3) is connected with an air source (1), and an outlet of the heater is divided into two paths, namely a test air path (4) and an adjusting bypass (8);
along the air flow direction, the test air path (4) is provided with a quick response pneumatic valve (5) and a test measurement model (6); the test measurement model (6) is an aeroengine turbine component or a simulated aeroengine turbine component; the first back pressure valve (9) is arranged on the regulating bypass (8) to regulate the initial pressure of the test gas circuit (4) and avoid the gas collecting effect when the quick response pneumatic valve (5) is closed; a back pressure valve II (7) is arranged on the test gas circuit (4) and positioned at the downstream of the test measurement model (6) so as to adjust the outlet back pressure of the test measurement model (6);
the power of the heater (3) is regulated to enable the gas in the test gas circuit (4) to present a temperature rise change process or a temperature reduction change process, so that the time-varying temperature transition state working condition simulation is realized;
the change rate of the quick response pneumatic valve (5) in the process from on to off or from off to on is controlled, so that the change process of the air inlet pressure at the upstream of the test measurement model (6) is regulated, the pressure condition of the inlet of the turbine component under the actual working conditions of starting, stopping, accelerating or decelerating of the aeroengine is simulated, and the pressure transition state working condition simulation is realized;
the test measurement model (6) is provided with an NI acquisition system (12), and the NI acquisition system (12) is used for acquiring and recording the changes of pressure, speed and temperature of the test measurement model (6) along with time in a transitional working condition and corresponds to an upstream pressure fluctuation curve of the test measurement model (6) caused by the action of the quick response pneumatic valve (5); the valve action signal of the quick response pneumatic valve (5) is synchronously triggered with the physical quantity signal acquisition, the quick response pneumatic valve (5) emits a pulse signal when the valve action is carried out, and the NI acquisition system (12) is triggered to start acquiring and recording data;
the rapid response pneumatic valve (5) is controlled by a pneumatic valve control box (13) connected with the rapid response pneumatic valve, the pneumatic valve control box (13) adopts a programmable control mode, and the change rate of the valve in the process from on to off or from off to on is controlled by a corresponding program, so that a specific target pressure curve is formed at the downstream of the rapid response pneumatic valve (5);
the NI acquisition system (12) and the pneumatic valve control box (13) are integrated by utilizing a LABVIEW program to realize synchronous acquisition of multiple physical quantities, so that a synchronous measurement system of the multiple physical quantities is formed based on the NI acquisition system (12); at the moment, the valve action signal of the quick response pneumatic valve (5) is synchronously triggered with the physical quantity signal acquisition, the pulse signal is sent out while the valve action of the quick response pneumatic valve (5) is simultaneously triggered, and the NI acquisition system (12) is triggered to start to acquire and record data.
2. The transition state performance test system of an aircraft engine turbine component of claim 1, wherein the fast response pneumatic valve (5) valve change time ranges from 1 to 10 seconds.
3. The transition state performance test system of an aircraft engine turbine component of claim 1, wherein the aircraft engine turbine component is a cascade, disk cavity, seal, or exhaust system.
4. The transition state performance test system of an aeroengine turbine component according to claim 1, wherein a dynamic pressure sensor (10) and a temperature sensor (11) are arranged between the fast response pneumatic valve (5) and the test measurement model (6) on the test gas path (4) to monitor whether the pressure change curve and the gas temperature meet target requirements.
5. The transition state performance test system of an aircraft engine turbine component according to claim 1, characterized in that the inlet air pressure upstream of the test measurement model (6) varies linearly or non-linearly.
6. The transition state performance test system of the turbine component of the aeroengine according to claim 1, wherein when the turbine component is a rotating component, the rotating speed of the motor is gradually increased or gradually decreased by controlling a power transformation cabinet of the motor, so as to realize the simulation of the working condition of changing the rotating speed; or when a vibration experiment is carried out, exciting force is applied to the rotating part from the outside by using the vibration exciter, and the vibration exciting force is gradually increased or gradually reduced by adjusting the vibration exciting force, so that the vibration exciting force changing working condition simulation is realized.
7. An experimental method for a transition state performance testing system based on a turbine component of an aircraft engine according to claim 1, comprising:
the power of the heater (3) is regulated to enable the gas in the test gas circuit (4) to present a temperature rise change process or a temperature reduction change process, so that an experiment is carried out, and the time-varying temperature transition state working condition simulation is realized;
the initial pressure of the test gas circuit (4) is adjusted to the initial target pressure of the variable working condition through adjusting the adjusting bypass (8);
the change rate of the quick response pneumatic valve (5) in the process from on to off or from off to on is controlled, so that the change process of the air inlet pressure at the upstream of the test measurement model (6) is regulated, the pressure condition of the inlet of the turbine component under the actual working conditions of starting, stopping, accelerating or decelerating the aeroengine is simulated, an experiment is carried out, and the pressure transition state working condition simulation is realized.
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