CN116380396A - Continuous wind tunnel aircraft atmospheric data system identification test system and method - Google Patents
Continuous wind tunnel aircraft atmospheric data system identification test system and method Download PDFInfo
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Abstract
The invention provides a continuous wind tunnel aircraft atmospheric data system identification test system and method, and belongs to the technical field of atmospheric data system identification test methods. The system comprises an identification test computer, a wind tunnel measurement and control system, an atmosphere data computer, an atmosphere pressure sensor, a wind tunnel total pressure sensor, a wind tunnel static pressure sensor and an aircraft model; the aircraft model is arranged in the wind tunnel test section and is connected with the atmospheric pressure sensor; the atmospheric pressure sensor is connected with the atmospheric data computer; the atmosphere data computer is connected with the identification computer; the identification computer is connected with the wind tunnel measurement and control system through a communication cable; the wind tunnel measurement and control system is respectively connected with the wind tunnel total pressure sensor and the wind tunnel static pressure sensor through cables; the method solves the technical problems of high dependence degree and high cost on flight tests in the traditional atmospheric data system test in the prior art, long system research and development period and high stability and accuracy verification difficulty on the atmospheric data system in the dangerous envelope range.
Description
Technical Field
The application relates to a system identification test method, in particular to a continuous wind tunnel aircraft atmospheric data system identification test system and method, and belongs to the technical field of atmospheric data system identification test methods.
Background
The air data system is a comprehensive and high-precision air data information system for the aircraft, and the flight control system, the navigation system, the instrument display system and the like of the modern aircraft all need accurate static pressure, dynamic pressure, altitude, indicated airspeed and other information. The atmospheric data system is an important on-board electronic device, and the accuracy of the calculated atmospheric parameters is directly related to the flight safety of the aircraft. The precedent that the output of error signals causes dangerous situations or even accidents of the aircraft is frequent due to the abnormal operation of the atmospheric data system (sensor wetting, icing, program logic problems, etc.).
The existing air data system test mainly uses an air data system component as an independent test piece or uses the appearance of a part of aircraft as a test piece to carry out a wind tunnel test, the test method is used as a conventional pressure measurement wind tunnel test to obtain pressure data in different airspeed, attitude angles and other states, after the test is completed, the air data system correction parameters are obtained through data post-processing, and finally, a flight test is carried out to carry out checking and correction to obtain final air data system parameters.
The existing research method uses a flight test as a checking means to check the stability, the accuracy and the like of an atmospheric data system, and has the defects of higher cost of the flight test, additional research and development period and cost are added after problems are found in the flight test, and the flight test has higher danger to the dangerous envelope checking of the atmospheric data system, influences the safety of an aircraft and is generally difficult to find. With the development of the aircraft technology, the requirements on the accuracy and stability of the air data system are gradually increased, and the requirements on the maneuverability and stall state flight of the aircraft are increased, so that the requirements on the dangerous envelope of the air data system are also increasingly important.
Disclosure of Invention
The following presents a simplified summary of the invention in order to provide a basic understanding of some aspects of the invention. It should be understood that this summary is not an exhaustive overview of the invention. It is not intended to identify key or critical elements of the invention or to delineate the scope of the invention. Its purpose is to present some concepts in a simplified form as a prelude to the more detailed description that is discussed later.
In view of the above, the invention provides a continuous wind tunnel aircraft atmospheric data system identification test system and a continuous wind tunnel aircraft atmospheric data system identification test method, which are used for solving the technical problems of high dependence on flight tests, high cost, long system research and development period and high stability and accuracy verification difficulty of a dangerous envelope range atmospheric data system in the traditional atmospheric data system test in the prior art.
Scheme one: a continuous wind tunnel aircraft atmospheric data system identification test system comprises an identification test computer, a wind tunnel measurement and control system, an atmospheric data computer, an atmospheric pressure sensor, a wind tunnel total pressure sensor, a wind tunnel static pressure sensor and an aircraft model;
the aircraft model is arranged in the wind tunnel test section and is connected with an atmospheric pressure sensor through a gas path hose;
the atmospheric pressure sensor is connected with an atmospheric data computer;
the atmosphere data computer is connected with the identification computer;
the identification computer is connected with the wind tunnel measurement and control system;
the wind tunnel measurement and control system is respectively connected with the wind tunnel total pressure sensor and the wind tunnel static pressure sensor.
Scheme II: a continuous wind tunnel aircraft atmospheric data system identification test method comprises the following steps:
s1, controlling the total pressure of a wind tunnel according to the requirements of an identification test, and controlling Mach number of a test section by adjusting the rotating speed of a wind tunnel compressor to establish the required flow field speed, total pressure and static pressure conditions in an aircraft model installation area of the wind tunnel test section;
s2, transmitting the pressure of the position of the airspeed tube or the pressure measuring hole on the aircraft model to an atmospheric pressure sensor through an air path hose;
s3, transmitting the pressure value acquired in real time to an atmosphere data computer by the atmosphere pressure sensor;
s4, the atmospheric data computer calculates the total pressure, the static pressure, the air pressure height, the indicated airspeed, the attack angle, the sideslip angle and the Mach number parameters of the position of the aircraft model in real time according to the received pressure value, and transmits the parameters to the identification test computer;
s5, the identification test computer carries out an identification test on the air data system.
Preferably, S5 specifically includes the following steps:
s51, calibrating a pressure sensor;
s52, correcting the pressure delay of the air pipeline;
s53, calculating atmospheric parameters of a wind tunnel flow field;
s54, atmosphere parameter identification and comparison.
Preferably, the calibration of the S51 pressure sensor, specifically, taking the height of the position of the test model in the wind tunnel test process as a reference, and maintaining the actual height difference between the total pressure and static pressure sensors for controlling the wind tunnel flow field and the position of the model during the calibration
Performing calibration, and maintaining the actual height difference of the atmospheric pressure sensor and the position of the model during the calibration>
Calibration is carried out to eliminate the difference in height>
And->
The resulting pressure deviation.
Preferably, the pressure delay correction of the air pipeline in the S52 is specifically to test the pressure delay time generated by the length of the air pipeline for test under different static pressure conditions when the different static pressures are obtained.
Preferably, the calculating of the atmospheric parameters of the S53 wind tunnel flow field specifically comprises the following steps:
s531, directly measuring to obtain total pressure of wind tunnel flow field
Total temperature of wind tunnel flow field->
Static pressure of wind tunnel residence chamber>
The method comprises the following steps of (1) a model pitch angle theta, a model roll angle phi, a pitch installation angle delta theta of an angle sensor and a model, a yaw installation angle delta phi of the angle sensor and the model, and a roll installation angle delta phi of the angle sensor and the model;
s532, resolving to obtain parameters of the Mach number, the total pressure, the static pressure, the air pressure height, the indicated airspeed, the attack angle and the sideslip angle of the model of the wind tunnel at the position of the aircraft model.
Preferably, the wind tunnel flow field Mach number, total pressure, static pressure, air pressure height, indicated airspeed, attack angle and sideslip angle parameters of the model are obtained by calculation:
mach number of wind tunnel flow field
Wherein M is the Mach number of the wind tunnel flow field,
for the wind tunnel residence chamber reference point mach number calculated from the residence chamber static pressure,
is->
And->
The correction amount of the test section is a specific calibration relation of each test section of each wind tunnel;
static pressure of wind tunnel flow field is
The air pressure is high
When γ+.0:
when γ=0:
wherein,,
is the standard gravity acceleration>
The method comprises the steps of carrying out a first treatment on the surface of the R is the gas constant of air, r= 287.05287 (++>
/K/>
)；/>
A lower limit value of the air pressure height of the corresponding height layer; />
The lower limit value of the atmospheric static pressure of the corresponding height layer; />
An atmospheric temperature lower limit value for the corresponding height layer; gamma is the vertical temperature of the corresponding height layer;
indicating airspeed as
Wherein,,
represents sea level standard sound velocity,/->
；
The attack angle of the model is
The sideslip angle is +.>
Correcting the installation angle from the angle sensor shafting to the model body shafting meets the following matrix relation:
wherein,,
euler rotation matrix of angle sensor relative to ground axis>
For the rotation matrix of the angle sensor relative to the model axis, < >>
For the model shafting relative to the groundRotation matrix of shafting>
A rotation matrix of the model shafting relative to the angle sensor;
the rotation matrix from the ground axis to the model body axis is as follows:
wherein,,
for Euler matrix rotated around x-axis, < >>
For Euler matrix rotated around y-axis, < >>
For Euler matrix rotated around z-axis, < >>
For the roll angle from the ground axis to the test model body axis, +.>
For yaw angle from the ground axis to the test model body axis, +.>
The pitch angle is from the ground axis system to the test model body axis system;
the mounting angle between the angle sensor and the model reference plane satisfies the following relation:
wherein the angle sensor is arranged at a pitching installation angle with the model
Yaw mounting angle->
Roll mounting angle->
The measuring method of (2) is that after 0 DEG leveling of the model reference plane, the angle sensor measures +.>
And->
The model reference plane is 90 DEG vertical to the ground, and the angle sensor measures +.>
；
The rotation matrix from the ground axis to the angle sensor is as follows:
wherein,,
for the roll angle of the angle sensor output, +.>
Yaw angle output by the angle sensor, < +.>
，/>
Pitch angle output for angle sensor>
from the following components
The method comprises the following steps:
angle of attack and sideslip angle of the model:
preferably, S54. The method comprises performing time stamp alignment on an atmospheric data computer, a wind tunnel measurement and control system and an identification test computer before starting a test, wherein the identification test computer receives parameters transmitted by the wind tunnel measurement and control system and parameters transmitted by the atmospheric data computer in real time during the test, performing delay time delta t correction on the received atmospheric data parameters transmitted by the atmospheric data computer, and comparing the corrected atmospheric data parameters with the atmospheric parameters calculated by the parameters transmitted by the wind tunnel measurement and control system, wherein the difference of the atmospheric data parameters is equal to or greater than the difference of the atmospheric data parameters
Wherein->
Is->
Respectively representing the calculated values of the same air data parameter at the same moment output by the wind tunnel flow field acquisition system and the air data computer, obtaining the difference of each air data parameter, and judging whether the difference meets the design index.
The third scheme is an electronic device, comprising a memory and a processor, wherein the memory stores a computer program, and the processor realizes the step of the continuous wind tunnel aircraft atmospheric data system identification test method in the second scheme when executing the computer program.
A fourth aspect is a computer readable storage medium having a computer program stored thereon, wherein the computer program when executed by a processor implements a continuous wind tunnel aircraft atmospheric data system identification test method as described in the second aspect.
The beneficial effects of the invention are as follows: the invention combines the advantages of variable pressure of a continuous transonic wind tunnel, simulates the atmospheric pressure of different heights, airspeeds and the like of the flying vehicle, and obtains Mach number, airspeeds and barometric altitude of the flying vehicle through measurement and calculation of the total pressure of the wind tunnel, a static pressure sensor and a total temperature sensor; by combining the advantage of higher wind tunnel test safety, the dangerous angle envelope of the aircraft is simulated, the attack angle and sideslip angle of the aircraft are calculated through an angle sensor arranged in the model, and the complete envelope range of the atmosphere data system is checked. According to the invention, the part identified by the flight test of the aircraft is verified in the ground wind tunnel test, so that problems can be found in the ground test stage and correction can be completed in the research and development stage, the flight test cost is reduced, the flight test safety is improved, and the identification test speed and angle boundary are enlarged.
Drawings
The accompanying drawings, which are included to provide a further understanding of the application and are incorporated in and constitute a part of this application, illustrate embodiments of the application and together with the description serve to explain the application and do not constitute an undue limitation to the application. In the drawings:
FIG. 1 is a schematic diagram of a continuous wind tunnel aircraft atmospheric data system qualification test system;
FIG. 2 is a schematic diagram of a continuous wind tunnel aircraft atmospheric data system identification test method;
FIG. 3 is a highly schematic view of sensor calibration according to the present invention;
FIG. 4 is a schematic diagram of pressure delay;
fig. 5 is a schematic diagram of the law of pressure delay with static pressure.
Detailed Description
In order to make the technical solutions and advantages of the embodiments of the present application more apparent, the following detailed description of exemplary embodiments of the present application is given with reference to the accompanying drawings, and it is apparent that the described embodiments are only some of the embodiments of the present application and not exhaustive of all the embodiments. It should be noted that, in the case of no conflict, the embodiments and features in the embodiments may be combined with each other.
Example 1, the description of this embodiment with reference to fig. 1, is a continuous wind tunnel aircraft atmospheric data system identification test system, which is characterized by comprising an identification test computer, a wind tunnel measurement and control system, an atmospheric data computer, an atmospheric pressure sensor, a wind tunnel total pressure sensor, a wind tunnel static pressure sensor and an aircraft model (respectively represented by test models in fig. 1 and 3);
the aircraft model is arranged in the wind tunnel test section and is connected with an atmospheric pressure sensor through a gas path hose;
specifically, the aircraft model comprises an aircraft appearance physical model, an atmospheric data system measuring piece such as a airspeed tube and the like arranged on the aircraft model, a pressure measuring hole for calculating the atmospheric data system applied to the surface of the aircraft model, and an angle sensor arranged in the model;
the atmospheric pressure sensor is connected with the atmospheric data computer through a cable;
the atmosphere data computer is connected with the identification computer through a communication cable;
the identification computer is connected with the wind tunnel measurement and control system through a communication cable;
the wind tunnel measurement and control system is respectively connected with the wind tunnel total pressure sensor and the wind tunnel static pressure sensor through cables;
specifically, the atmospheric pressure sensor and the atmospheric data computer are arranged outside the wind tunnel test section and in the wind tunnel residence chamber.
The operating principle of the continuous wind tunnel aircraft atmospheric data system identification test system is as follows: according to the requirements of the identification test, the required conditions such as flow field speed, total pressure, static pressure and the like are established in the aircraft model installation area of the wind tunnel test section, the airspeed tube or the pressure measuring holes on the surface of the aircraft sense the pressure value of the position in real time, the pressure value is transmitted to the atmospheric pressure sensor in the wind tunnel residence chamber outside the test section through the air path hose, the atmospheric pressure sensor measures the pressure value of each pressure measuring hole in real time and transmits the pressure value to the atmospheric data computer, the atmospheric data computer calculates the airspeed, the pressure and other atmospheric parameters of the position of the aircraft in real time and transmits the atmospheric parameters to the identification test computer through the communication cable, and the identification test computer synchronously, stores and compares the atmospheric data parameters transmitted by the wind tunnel measurement and control system and the atmospheric data computer, so that the identification test of the atmospheric data system is completed.
Example 2, the present embodiment will be described with reference to fig. 2 to 5, which is a continuous wind tunnel aircraft atmospheric data system identification test method, wherein an atmospheric data system and an aircraft model are installed in a wind tunnel, and the wind tunnel flow field is used as a reference to perform an assessment and identification on the atmospheric data system, and the stability and the atmospheric data calculation accuracy of the assessment and identification are examined; the advantages of variable pressure of the continuous transonic wind tunnel are combined, the atmospheric pressure of different heights, airspeeds and the like of the flying of the aircraft can be simulated in the wind tunnel, mach numbers, airspeeds, barometric pressures and the like of the aircraft are obtained through calculation of the pressures measured by the wind tunnel total pressure and the static pressure sensor, meanwhile, the dangerous angle envelope of the aircraft can be simulated by combining the advantages of higher safety of the wind tunnel test, the attack angle and sideslip angle of the aircraft are obtained through calculation by an angle sensor arranged in the model, and the complete envelope range of the atmospheric data system is checked; the method specifically comprises the following steps:
s1, controlling the total pressure of a wind tunnel according to the requirements of an identification test, and controlling Mach number of a test section by adjusting the rotating speed of a wind tunnel compressor to establish the required flow field speed, total pressure and static pressure conditions in an aircraft model installation area of the wind tunnel test section;
s2, transmitting the pressure of the position of the airspeed tube or the pressure measuring hole on the aircraft model to an atmospheric pressure sensor through an air path hose;
s3, transmitting the pressure value acquired in real time to an atmosphere data computer by the atmosphere pressure sensor;
s4, the atmospheric data computer calculates the total pressure, the static pressure, the air pressure height, the indicated airspeed, the attack angle, the sideslip angle and the Mach number parameters of the position of the aircraft model in real time according to the received pressure value, and transmits the parameters to the identification test computer;
s5, an identification test computer carries out an identification test on the air data system, and the method specifically comprises the following steps:
s51, calibrating a pressure sensor, specifically referring to FIG. 3, taking the height of the position of a test model in the wind tunnel test process as a reference, and maintaining the actual height difference between the total pressure and static pressure sensors for controlling the wind tunnel flow field and the position of the model during calibration
Performing calibration, and maintaining the actual height difference of the atmospheric pressure sensor and the position of the model during the calibration>
Calibration is carried out to eliminate the difference in height>
And->
The resulting pressure deviation was about 12Pa/m under normal pressure conditions.
S52, correcting pressure delay of the air pipeline, specifically referring to fig. 4-5, because the volume of the atmospheric pressure sensor is larger, the space of the inner cavity of the test model is limited, the atmospheric pressure sensor cannot be installed, the atmospheric pressure sensor is installed outside a wind tunnel test section and in a residence chamber, a pressure transmission delay delta t is generated by an air path hose between the atmospheric pressure sensor and a pressure measuring hole of the wind tunnel test model, and the pressure delay time generated by the length of the air pipeline for test under different static pressure conditions is required to be tested before the test.
S53, calculating atmospheric parameters of a wind tunnel flow field, and specifically comprising the following steps of:
s531, directly measuring to obtain total pressure of wind tunnel flow field
Total temperature of wind tunnel flow field->
Static pressure of wind tunnel residence chamber>
The method comprises the following steps of (1) a model pitch angle theta, a model roll angle phi, a pitch installation angle delta theta of an angle sensor and a model, a yaw installation angle delta phi of the angle sensor and the model, and a roll installation angle delta phi of the angle sensor and the model;
s532, resolving to obtain parameters of the Mach number, the total pressure, the static pressure, the air pressure height, the indicated airspeed, the attack angle and the sideslip angle of the model of the wind tunnel at the position of the aircraft model.
Mach number of wind tunnel flow field
Wherein M is the Mach number of the wind tunnel flow field,
for the wind tunnel residence chamber reference point mach number calculated from the residence chamber static pressure,
is->
And->
The correction amount of the test section is a specific calibration relation of each test section of each wind tunnel;
static pressure of wind tunnel flow field is
The air pressure is high
When γ+.0:
when γ=0:
wherein,,
is the standard gravity acceleration>
The method comprises the steps of carrying out a first treatment on the surface of the R is the gas constant of air, r= 287.05287 (++>
/K/>
)；/>
A lower limit value of the air pressure height of the corresponding height layer; />
The lower limit value of the atmospheric static pressure of the corresponding height layer; />
An atmospheric temperature lower limit value for the corresponding height layer; gamma is the vertical temperature of the corresponding height layer; details are shown in Table 1->
、/>
A relationship table between the two;
TABLE 1
、/>
Relationship table between
Indicating airspeed as
Wherein,,
represents sea level standard sound velocity,/->
。
The attack angle of the model is
The sideslip angle is +.>
Correcting the installation angle from the angle sensor shafting to the model body shafting meets the following matrix relation:
wherein,,
euler rotation matrix of angle sensor relative to ground axis>
For the rotation matrix of the angle sensor relative to the model axis, < >>
For the rotation matrix of the model axis relative to the ground axis, < >>
A rotation matrix of the model shafting relative to the angle sensor;
the rotation matrix from the ground axis to the model body axis is as follows:
wherein,,
for Euler matrix rotated around x-axis, < >>
For Euler matrix rotated around y-axis, < >>
For Euler matrix rotated around z-axis, < >>
For the roll angle from the ground axis to the test model body axis, +.>
For yaw angle from the ground axis to the test model body axis, +.>
The pitch angle is from the ground axis system to the test model body axis system;
the mounting angle between the angle sensor and the model reference plane satisfies the following relation:
wherein the angle sensor is arranged at a pitching installation angle with the model
Yaw mounting angle->
Roll mounting angle->
The measuring method of (2) is that after 0 DEG leveling of the model reference plane, the angle sensor measures +.>
And->
The model reference plane is 90 DEG vertical to the ground, and the angle sensor measures +.>
；
The rotation matrix from the ground axis to the angle sensor is as follows:
wherein,,
for the roll angle of the angle sensor output, +.>
Yaw angle output by the angle sensor, < +.>
，/>
A pitch angle output by the angle sensor;
from the following components
The method comprises the following steps:
angle of attack and sideslip angle of the model:
s54, identifying and comparing the atmospheric parameters, namely, carrying out timestamp alignment on an atmospheric data computer, a wind tunnel measurement and control system and an identification test computer before the test starts, wherein the identification test computer receives the parameters sent by the wind tunnel measurement and control system and the parameters sent by the atmospheric data computer in real time in the test process, carrying out delay time delta t correction on the received atmospheric data parameters sent by the atmospheric data computer, and then comparing the atmospheric data parameters with the atmospheric parameters calculated by the parameters sent by the wind tunnel measurement and control system, wherein the difference of the atmospheric data parameters is calculated
Wherein->
Is->
Respectively representing the calculated values of the same air data parameter at the same moment output by the wind tunnel flow field acquisition system and the air data computer, obtaining the difference of each air data parameter, and judging whether the difference meets the design index.
Different system design indexes and requirements are different, are not important improvement points of the application, and generally require Mach numbers not exceeding +/-0.005, total pressures not exceeding +/-100 Pa, static pressures not exceeding +/-80 Pa, air pressure heights not exceeding +/-20 m, indicated airspeeds not exceeding +/-6 km/h, attack angles and sideslip angles not exceeding +/-0.8 degrees.
The invention can complete the operation of the identification and correction of the air data system in the flight test in the ground wind tunnel test, thereby reducing the test cost;
according to the invention, dangerous flight attitudes which cannot be realized in a flight test can be realized, and more comprehensive atmospheric data system parameters are obtained.
In embodiment 3, the computer device of the present invention may be a device including a processor and a memory, for example, a single chip microcomputer including a central processing unit. And the processor is used for executing the computer program stored in the memory to realize the steps of the continuous wind tunnel aircraft atmospheric data system identification test method.
The processor may be a central processing unit (Central Processing Unit, CPU), other general purpose processors, digital signal processors (Digital Signal Processor, DSP), application specific integrated circuits (Application Specific Integrated Circuit, ASIC), off-the-shelf programmable gate arrays (Field-Programmable Gate Array, FPGA) or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, or the like. A general purpose processor may be a microprocessor or the processor may be any conventional processor or the like.
The memory may mainly include a storage program area and a storage data area, wherein the storage program area may store an operating system, an application program (such as a sound playing function, an image playing function, etc.) required for at least one function, and the like; the storage data area may store data (such as audio data, phonebook, etc.) created according to the use of the handset, etc. In addition, the memory may include high-speed random access memory, and may also include non-volatile memory, such as a hard disk, memory, plug-in hard disk, smart Media Card (SMC), secure Digital (SD) Card, flash Card (Flash Card), at least one disk storage device, flash memory device, or other volatile solid-state storage device.
Embodiment 4, computer-readable storage Medium embodiment
The computer readable storage medium of the present invention may be any form of storage medium readable by a processor of a computer device, including but not limited to, nonvolatile memory, volatile memory, ferroelectric memory, etc., having a computer program stored thereon, which when read and executed by the processor of the computer device, implements the steps of a continuous wind tunnel aircraft air data system qualification test method as described above.
The computer program comprises computer program code which may be in source code form, object code form, executable file or in some intermediate form, etc. The computer readable medium may include: any entity or device capable of carrying the computer program code, a recording medium, a U disk, a removable hard disk, a magnetic disk, an optical disk, a computer Memory, a Read-Only Memory (ROM), a random access Memory (RAM, random Access Memory), an electrical carrier signal, a telecommunications signal, a software distribution medium, and so forth. It should be noted that the computer readable medium contains content that can be appropriately scaled according to the requirements of jurisdictions in which such content is subject to legislation and patent practice, such as in certain jurisdictions in which such content is subject to legislation and patent practice, the computer readable medium does not include electrical carrier signals and telecommunication signals.
While the invention has been described with respect to a limited number of embodiments, those skilled in the art, having benefit of the above description, will appreciate that other embodiments are contemplated within the scope of the invention as described herein. Furthermore, it should be noted that the language used in the specification has been principally selected for readability and instructional purposes, and may not have been selected to delineate or circumscribe the inventive subject matter. Accordingly, many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the appended claims. The disclosure of the present invention is intended to be illustrative, but not limiting, of the scope of the invention, which is defined by the appended claims.
Claims (10)
1. The continuous wind tunnel aircraft atmospheric data system identification test system is characterized by comprising an identification test computer, a wind tunnel measurement and control system, an atmospheric data computer, an atmospheric pressure sensor, a wind tunnel total pressure sensor, a wind tunnel static pressure sensor and an aircraft model;
the aircraft model is arranged in the wind tunnel test section and is connected with an atmospheric pressure sensor through a gas path hose;
the atmospheric pressure sensor is connected with an atmospheric data computer;
the atmosphere data computer is connected with the identification computer;
the identification computer is connected with the wind tunnel measurement and control system;
the wind tunnel measurement and control system is respectively connected with the wind tunnel total pressure sensor and the wind tunnel static pressure sensor.
2. The continuous wind tunnel aircraft atmospheric data system identification test method is characterized by comprising the following steps of:
s1, controlling the total pressure of a wind tunnel according to the requirements of an identification test, and controlling Mach number of a test section by adjusting the rotating speed of a wind tunnel compressor to establish the required flow field speed, total pressure and static pressure conditions in an aircraft model installation area of the wind tunnel test section;
s2, transmitting the pressure of the position of the airspeed tube or the pressure measuring hole on the aircraft model to an atmospheric pressure sensor through an air path hose;
s3, transmitting the pressure value acquired in real time to an atmosphere data computer by the atmosphere pressure sensor;
s4, the atmospheric data computer calculates the total pressure, the static pressure, the air pressure height, the indicated airspeed, the attack angle, the sideslip angle and the Mach number parameters of the position of the aircraft model in real time according to the received pressure value, and transmits the parameters to the identification test computer;
s5, the identification test computer carries out an identification test on the air data system.
3. The continuous tunnel aircraft atmospheric data system identification test method according to claim 2, wherein S5 specifically comprises the following steps:
s51, calibrating a pressure sensor;
s52, correcting the pressure delay of the air pipeline;
s53, calculating atmospheric parameters of a wind tunnel flow field;
s54, atmosphere parameter identification and comparison.
4. The method for testing the identification of the air data system of the continuous wind tunnel aircraft according to claim 3, wherein the calibration of the S51 pressure sensor is specifically to take the height of the position of the test model in the wind tunnel test process as a reference, and the actual height difference between the total pressure and static pressure sensors for controlling the wind tunnel flow field and the position of the model is maintained during the calibration of the total pressure and static pressure sensors
Performing calibration, and maintaining the actual height difference of the atmospheric pressure sensor and the position of the model during the calibration>
Calibration is carried out to eliminate the difference in height>
And->
The resulting pressure deviation.
5. The method for testing the identification of the air data system of the continuous wind tunnel aircraft according to claim 4, wherein the pressure delay of the S52 air pipeline is corrected, specifically, the pressure delay time generated by the length of the air pipeline for testing under different static pressures is tested when different static pressures are obtained.
6. The continuous wind tunnel aircraft atmospheric data system identification test method according to claim 5, wherein the step of calculating the atmospheric parameters of the flow field of the wind tunnel S53 specifically comprises the following steps:
s531, directly measuring to obtain total pressure of wind tunnel flow field
Total temperature of wind tunnel flow field->
Static pressure of wind tunnel residence chamber>
The method comprises the following steps of (1) a model pitch angle theta, a model roll angle phi, a pitch installation angle delta theta of an angle sensor and a model, a yaw installation angle delta phi of the angle sensor and the model, and a roll installation angle delta phi of the angle sensor and the model;
s532, resolving to obtain parameters of the Mach number, the total pressure, the static pressure, the air pressure height, the indicated airspeed, the attack angle and the sideslip angle of the model of the wind tunnel at the position of the aircraft model.
7. The continuous wind tunnel aircraft atmospheric data system identification test method according to claim 6, wherein wind tunnel flow field Mach number, total pressure, static pressure, air pressure height, indicated airspeed, attack angle and sideslip angle parameters of the model are obtained by calculation:
mach number of wind tunnel flow field
Wherein M is the Mach number of the wind tunnel flow field,
for wind tunnel resident reference Mach number calculated from resident static pressure, +.>
Is->
And->
The correction amount of the test section is a specific calibration relation of each test section of each wind tunnel;
static pressure of wind tunnel flow field is
；
The air pressure is high
；
When γ+.0:
when γ=0:
wherein,,
is the standard gravity acceleration>
The method comprises the steps of carrying out a first treatment on the surface of the R is the gas constant of air, r= 287.05287%
/K/>
)；/>
A lower limit value of the air pressure height of the corresponding height layer; />
The lower limit value of the atmospheric static pressure of the corresponding height layer; />
An atmospheric temperature lower limit value for the corresponding height layer; gamma is the vertical temperature of the corresponding height layer;
indicating airspeed as
；
Wherein,,
represents sea level standard sound velocity,/->
；
The attack angle of the model is
The sideslip angle is +.>
；
Correcting the installation angle from the angle sensor shafting to the model body shafting meets the following matrix relation:
wherein,,
euler rotation matrix of angle sensor relative to ground axis>
For the rotation matrix of the angle sensor relative to the model axis, < >>
For the rotation matrix of the model axis relative to the ground axis, < >>
A rotation matrix of the model shafting relative to the angle sensor;
the rotation matrix from the ground axis to the model body axis is as follows:
wherein,,
for Euler matrix rotated around x-axis, < >>
For Euler matrix rotated around y-axis, < >>
For Euler matrix rotated around z-axis, < >>
For the roll angle from the ground axis to the test model body axis, +.>
For yaw angle from the ground axis to the test model body axis, +.>
The pitch angle is from the ground axis system to the test model body axis system;
the mounting angle between the angle sensor and the model reference plane satisfies the following relation:
wherein the angle sensor is arranged at a pitching installation angle with the model
Yaw mounting angle->
Roll mounting angle->
The measuring method of (2) is that after 0 DEG leveling of the model reference plane, the angle sensor measures +.>
And->
The model reference plane is 90 DEG vertical to the ground, and the angle sensor measures +.>
；
The rotation matrix from the ground axis to the angle sensor is as follows:
wherein,,
for the roll angle of the angle sensor output, +.>
Yaw angle output by the angle sensor, < +.>
，/>
A pitch angle output by the angle sensor;
from the following components
The method comprises the following steps:
angle of attack and sideslip angle of the model:
8. the method for testing the continuous wind tunnel aircraft atmospheric data system according to claim 7, wherein S54 is used for identifying and comparing the atmospheric parameters, specifically, the atmospheric data computer, the wind tunnel measurement and control system and the identification test computer are subjected to timestamp alignment before the test is started, the identification test computer receives the parameters sent by the wind tunnel measurement and control system and the parameters sent by the atmospheric data computer in real time in the test process, the delay time delta t of the atmospheric data parameters sent by the received atmospheric data computer is corrected, and then the atmospheric parameters obtained by calculating the parameters sent by the wind tunnel measurement and control system are compared, and the difference of the atmospheric data parameters is calculated
Wherein->
Is->
Respectively representing the calculated values of the same air data parameter at the same moment output by the wind tunnel flow field acquisition system and the air data computer, obtaining the difference of each air data parameter, and judging whether the difference meets the design index.
9. An electronic device comprising a memory and a processor, the memory storing a computer program, the processor implementing the steps of a continuous wind tunnel aircraft atmospheric data system qualification test method of any one of claims 2-8 when executing the computer program.
10. A computer readable storage medium having stored thereon a computer program, which when executed by a processor implements a continuous wind tunnel aircraft atmospheric data system qualification test method according to any one of claims 2 to 8.
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