CN106840573B - A kind of Flush Airdata Sensing System scaling method - Google Patents
A kind of Flush Airdata Sensing System scaling method Download PDFInfo
- Publication number
- CN106840573B CN106840573B CN201611180132.2A CN201611180132A CN106840573B CN 106840573 B CN106840573 B CN 106840573B CN 201611180132 A CN201611180132 A CN 201611180132A CN 106840573 B CN106840573 B CN 106840573B
- Authority
- CN
- China
- Prior art keywords
- sensing system
- angle
- attack
- error
- dummy vehicle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
Landscapes
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- General Physics & Mathematics (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The present invention relates to a kind of Flush Airdata Sensing System scaling methods, belong to flush airdata sensing system technical field, which demarcates Flush Airdata Sensing System using supersonic wind tunnel.Step wave waveform and the intersection position of demarcated Mach number are measured using stration technique, the shrinkage scale cun of Flush Airdata Sensing System wind tunnel model is designed according to measurement result, installation site and the test angle of attack, ensure test model in wind-tunnel calibration process, locating flow field position can avoid the influence of step wave always.This method alsies specify mould processing precision, and pressure tap installs verticality, various requirements such as model installation level and pressure-measuring pipe road sealing detection.Flush Airdata Sensing System calibration is set to obtain true measurement accuracy.The system calibrating result precision of this method is high, high-efficient and high reliablity.
Description
Technical field
The present invention relates to Flush Airdata Sensing System calibration to design, more particularly to flush airdata sensing system
System wind tunnel test scaling method, belongs to flush airdata sensing system technical field.
Background technique
Flush airdata sensing system (Flush Air Data Sensing, FADS) system, by pressure sensor battle array
The pressure distribution of column measurement aircraft surface, the flight such as static pressure, Mach number, the angle of attack, yaw angle is obtained by special algorithm indirectly
Atmosphere data.FADS system is that supersonic aircraft obtains main path of the flight atmosphere data as flight control input, special
It is not particularly important for the supersonic aircraft of air suction type.
FADS system calibrating includes wind tunnel test and flight test calibration, but since flight test general technical is very multiple
It is miscellaneous, research fund is expensive, target range difficult coordination and test preparatory period are long etc., so carrying out flight test calibration difficulty very
Greatly.Therefore ground wind tunnel test becomes FADS system calibrating technical way.FADS system wind tunnel test calibration is a pole
Its fine special test is influenced by many factors such as flow field quality, modelling processing and installations.
FADS system main operational principle is passed through by the pressure distribution of array of pressure sensors measurement aircraft surface
Special algorithm obtains aircraft incoming flow atmosphere data indirectly.Therefore system calibrating primary key technical problems be stablized,
Uniform high-quality Flow Field in Wind Tunnel guarantees that the pressure value of measuring point on model is not influenced ability accurate calibration by extraneous other factors
The measurement accuracy of FADS system.If failing to exclude extraneous other factors interference, i.e., can not accurate calibration FADS system accurately survey
Accuracy of measurement will lead to the failure of FADS system development.For domestic existing type of production wind-tunnel, Flow Field in Wind Tunnel Quality Design is usually equal
It is to be designed according to national military standard technical standard, mainly meets aircraft dynamometry and survey the routine tests mission requirements such as pressure, wind-tunnel
Most of position flow field quality is to be unable to satisfy the similar special test of FADS system calibrating in test section.Therefore wind-tunnel is completed
Experimental Calibration must carry out in-depth analysis research to Flow Field in Wind Tunnel structure, avoid various interference and find suitable FADS system calibrating
Flow field regions.Flow field quality is influenced in wind tunnel test mainly by two aspect influence factors of step wave and wind tunnel wall interference.Step wave
The gap docked by the jet pipe of wind-tunnel with test section generates expansion wave system and forms, and can not eliminate the effects of the act can only avoid.Hole wall is dry
Disturbing is to get to wall reflection in hole by test model head base body shock wave to cause, can only be by designing a model blockage ratio and Control experiment most
The big angle of attack is to reduce its influence.Both the above factor will affect the pressure value on FADS system pressure tap array layout, cause
There is deviation in algorithm calculation result, seriously affects FADS system calibrating measurement accuracy.Second key technical problems is that model is set
Meter and processing, it has to be warranted that the location error of pressure tap, pressure tap and wall surface verticality, it is ensured that the installation knot at model docking
Structure intensity buffets to generate to loosen to prevent supersonic speed high frequency during test and causes measurement error.Third key technical problems are
The installation accuracy control of test model, for the axisymmetric model of air-breathing missile, model, which changes the outfit, is easiest to occur circumferential deviation,
The angle of attack and sideslip measurement error are significantly affected generation.Therefore it must be carried out according to certain standard after model is installed corresponding
Detection and fine tuning.
Since domestic FADS system research is started late, specific standard is had not yet been formed to system calibrating method, lacks system
The Specifications of system calibration.
Summary of the invention
The present invention proposes a kind of Flush Airdata Sensing System to overcome the shortcomings of existing Wind Tunnel Technique condition
Scaling method, can be with the measurement accuracy of accurate calibration system.
The object of the invention is achieved by following technical solution:
A kind of Flush Airdata Sensing System scaling method is provided, is included the following steps:
(1) parameter of supersonic wind tunnel, jet pipe length L1mm, sectional dimension B1mm × B1mm, test Mach number Ma are obtained
It is 2.0~4.0, aircraft prototype length is L2mm, equivalent diameter Dmm；According to test demand estimate range of angles of attack be-α~
α；
(2) using stration technique measurement Mach number 2~4 when, the position of the angle of up/down steps wave and upper and lower wave intersection point is surveyed
Intersection point is measured apart from nozzle entry sectional position size X1mm, the cone of upper and lower wave and nozzle section composition region, along wind-tunnel incoming flow
The median plane in direction is isosceles triangle, base angle 2 and apex angle α 3；
(3) the ratio value Sc of modulus type contracting ratio is 1；
(4) as step wave intersection point X1 < 1/2L1mm, jet pipe back segment is selected as test area, strut lengths L0It calculates public
Formula are as follows:
As step wave intersection point X1 >=1/2L1mm, leading portion test area, strut lengths L are selected0Calculation formula are as follows:
L0≥(Sc.D/2)/tan((arctan(B1/(L1-L3))×180/π-α)×π/180)+L3-L4-100
It is L3mm that wherein attack angle mechanism, which turns the heart apart from nozzle exit, and attack angle mechanism and strut connector are apart from nozzle exit
L4mm；
(5) if the length L of calculated strut0It is not at effective range, then readjusts step range of angles of attack and contracting
The ratio value of ratio takes Sc=Sc-0.1 or α=α -2, return step (4)；If L0In effective range, it is determined that α, Sc and L0；
(6) according to the dimensioned dummy vehicle after the ratio value Sc contracting ratio of contracting ratio, according to L0Process strut；
(7) Flush Airdata Sensing System is mounted on inside dummy vehicle, dummy vehicle is pacified by strut
On attack angle mechanism；
(8) blasting is carried out in 2~4 range of Mach number, attack angle mechanism is run within the scope of-α to α, embedded big
Destiny measures static pressure, Mach number, the angle of attack and the yaw angle of dummy vehicle according to sensor-based system；By flush airdata sensing system system
The static pressure of the dummy vehicle of static pressure, Mach number, the angle of attack and sideslip angular amount and the wind tunnel system measurement of the dummy vehicle of unified test,
Mach number, the angle of attack compare with yaw angle, judge whether airdata sensing system measurement is accurate.
Preferably, the length L of strut0Effective range be greater than 0.1L1, and be less than 0.6L1.
Preferably, judge whether Flush Airdata Sensing System measurement is accurate method particularly includes: static pressure error≤
500Pa, Mach number error≤0.1, angle of attack error≤0.5 °, sideslip angle error≤0.5 ° then judge that embedded atmosphere data passes
Sensing system measurement is accurate；If wherein thering is parameter error to be more than threshold range, judge that Flush Airdata Sensing System should
Parameter measurement error precision meet design requirements.
Preferably, dummy vehicle is processed according to Sc, each dimension precision requirement in dummy vehicle surface reaches ± 0.02mm,
Dummy vehicle surface pressure tap status requirement: the pressure tap position circumferential direction deviation positioned at dummy vehicle vertex is less than 0.1mm,
The error of perpendicularity≤3 ', the position axial error for playing pressure tap on wall are less than 0.1mm, and circumferential error is less than 0.1mm, processes survey
After pressing hole, to dummy vehicle surface polishing.
Preferably, horizontal gradient error≤3 ' after dummy vehicle is mounted on strut.
Preferably, pressure-measuring pipe road is installed, load cell goes out through pressure-measuring pipe pass is mounted on flight model table in pressure tap
Face after dummy vehicle is installed, is pressurized to 105kPa~130kPa to pressure tap, Flush Airdata Sensing System is anti-
Between seasonable≤0.5s, stablize time >=3s；Negative pressure is to 90kPa~5kPa, system reaction time≤0.5s.
The invention has the following advantages over the prior art:
(1) present invention avoids step wave by adjusting the installation site (length of strut) and size of dummy vehicle
Interference, obtain uniform flow field, guarantee the quality of wind tunnel test.The system calibrating result precision of this method is high, it is high-efficient and
High reliablity.
(2) present invention avoids wind tunnel wall interference, further increases by adjusting the size and range of angles of attack of dummy vehicle
Test accuracy, it is ensured that the confidence level of test.
(3) invention provides for the sealing detection standard of model accuracy and pressure-measuring pipe road, it ensure that the reliability of test.
Detailed description of the invention
Fig. 1 is the original dimension schematic diagram of dummy vehicle in the embodiment of the present invention；
Fig. 2 is FADS system and wind-tunnel peg model scheme of installation in the embodiment of the present invention；
Fig. 3 is test schematic diagram of the invention.
Specific embodiment
Referring to Fig. 1, FADS system is mounted on inside flight model, and P1 to P9 is the pressure-measuring pipe road of FADS system, is surveyed pressure and is passed
Sensor goes out through pressure-measuring pipe pass is mounted on flight model surface, measures surface pressing.Flight model is mounted on wind-tunnel by strut
In test section, referring to fig. 2.Flight model is mounted on attack angle mechanism by strut, around angle of attack machine under the control of attack angle mechanism
Structure turns heart rotation, and the range that when test rotates is that range of angles of attack is-α~α.
Supersonic wind tunnel step wave waveform and intersection position are measured using stration technique, as shown in figure 3, conventional method uses
Dummy vehicle, due to not accounting for the influence of step wave, in the sphere of action in step wave, cause calibration fail, this
Invention considers the influence of step wave, dummy vehicle is placed in except the sphere of action of step wave, it is ensured that wind tunnel test calibration
Process model avoids step wave, is calculated using step wave intersection position and obtains model contracting than ratio value and strut lengths.Specific packet
Include the following steps:
(1) parameter of supersonic wind tunnel, jet pipe length L1mm, sectional dimension B1mm × B1mm, test Mach number Ma are obtained
It is 2.0~4.0；Aircraft prototype length is L2mm, equivalent diameter Dmm；According to test demand estimate range of angles of attack be-α~
α；
(2) using stration technique measurement Mach number 2~4 when, the angle of up/down steps wave and the position of upper and lower wave intersection point are obtained
It sets, measurement intersection point is apart from nozzle entry sectional position size X1mm, the cone of upper and lower wave and nozzle section composition region, along wind-tunnel
The median plane of direction of flow is isosceles triangle, base angle 2 and apex angle α 3；
(3) the ratio value Sc of modulus type contracting ratio is 1；
(4) as step wave intersection point X1 < 1/2L1mm, jet pipe back segment is selected as test area, strut lengths L0It calculates public
Formula are as follows:
As step wave intersection point X1 >=1/2L1mm, leading portion test area, strut lengths L are selected0Calculation formula are as follows:
L0≥(Sc.D/2)/tan((arctan(B1/(L1-L3))×180/π-α)×π/180)+L3-L4-100
It is L3mm that wherein attack angle mechanism, which turns the heart apart from nozzle exit, and attack angle mechanism and strut connector are apart from nozzle exit
L4mm；
(5) if the length L of calculated strut0It is not at effective range, then readjusts step range of angles of attack, takes Sc
=Sc-0.1 or α=α -2, return step (4)；If L0In effective range, it is determined that α, Sc and L0；The length L of strut0's
Effective range be greater than 0.1L1, and be less than 0.6L1, if model can not normal mounting not in effective range；
(6) dummy vehicle is processed according to Sc, according to L0Process strut；
(7) Flush Airdata Sensing System is mounted on inside dummy vehicle, and dummy vehicle is installed by strut
On attack angle mechanism；
(8) blasting is carried out in 2~4 range of Mach number, attack angle mechanism is run within the scope of-α to α, embedded big
Destiny measures static pressure, Mach number, the angle of attack and the yaw angle of dummy vehicle according to sensor-based system；Airdata sensing system is surveyed
Static pressure, the Mach of the dummy vehicle of static pressure, Mach number, the angle of attack and sideslip angular amount and the wind tunnel system measurement of dummy vehicle
Number, the angle of attack compare with yaw angle, judge whether airdata sensing system measurement is accurate.Judge that airdata sensing system is surveyed
It whether accurate measures method particularly includes: static pressure error≤500Pa, Mach number error≤0.1, angle of attack error≤0.5 °, yaw angle
Error≤0.5 ° then judges that airdata sensing system measurement is accurate；If wherein thering is parameter error to be more than threshold range, sentence
Disconnected airdata sensing system parameter measurement meet design requirements.
Dummy vehicle is processed according to Sc, each dimension precision requirement in dummy vehicle surface reaches ± 0.02mm, aircraft
Model surface pressure tap status requirement: the pressure tap position circumferential direction deviation positioned at dummy vehicle vertex is less than 0.1mm, verticality
Error≤3 ', the position axial error for playing pressure tap on wall are less than 0.1mm, and circumferential error is less than 0.1mm, processes pressure tap
Afterwards, to dummy vehicle surface polishing.Horizontal gradient error≤3 ' after dummy vehicle is mounted on strut.To guarantee adding for model
Work and installation quality improve measurement accuracy.
Pressure-measuring pipe road is installed, load cell goes out to be mounted on flight model surface through pressure-measuring pipe pass, flies in pressure tap
After device model is installed, 105kPa~130kPa, Flush Airdata Sensing System reaction time are pressurized to pressure tap
≤ 0.5s stablizes time >=3s；Negative pressure is to 90kPa~5kPa, system reaction time≤0.5s, guarantee pressure-measuring pipe road do not block,
It is air tight.
The present invention is described in further detail combined with specific embodiments below:
Embodiment 1
The present embodiment the present invention will be described in detail by taking the Mach number 2 of rating test as an example specific implementation process:
(1) supersonic wind tunnel jet pipe length is 2400mm, and sectional dimension is 1200mm × 1200mm, and model length is
871mm, model maximum gauge are 268mm.It is L3=800mm, angle of attack connector distance spray that attack angle mechanism, which turns the heart apart from nozzle exit,
The distance of pipe outlet is L4=114mm.Testing the range of angles of attack estimated is α=- 12 °~12 °.
(2) the step wave structure of Mach number 2 and the position of upper and lower wave intersection point in supersonic nozzle are measured using stration technique,
Intersection position size X1=1050mm is measured, the coverage of step wave is determined with this.Since intersection position is apart from attack angle mechanism
Turn the heart farther out, if the leading portion of selected jet pipe is unable to satisfy as trial zone is estimated test range of angles of attack, therefore selected supersonic speed
The wind-tunnel stream that the back segment Delta Region (1050mm~2400mm) of wind tunnel nozzle is demarcated as Flush Airdata Sensing System
Field areas is high 2=60.3 ° of base angle for calculating triangular domain with X1.α=- 12 ° of range of angles of attack~12 ° are estimated according to test, are attacked
Angle mechanism turns heart L3=800mm, step wave intersection point X1=1050mm, 2=60.3 ° of step wave Delta Region base angle and α 3=
59.4°；
(3) the length L0=300mm (wherein Ma=2) of the ratio value Sc=0.8 of model contracting ratio and installation strut.Strut is long
Degree is attack angle mechanism connector size at a distance from model bottom.Ensure during the test, all angle of attack state drags are equal
Step wave can be avoided, and guarantees the point on model surface apart from step wave proximal most position >=100mm.
(4) strut lengths L0=300mm, and in effective range, processing experiment model and strut.Mould processing surface
Each dimension precision requirement reaches within ± 0.02mm, pressure tap status requirement on model: the position circumferential direction deviation on vertex≤
0.1mm, the error of perpendicularity≤3 ' play the position axial error≤0.1mm, circumferential error≤0.1mm of pressure tap on wall.It processes
After pressure tap, to dummy vehicle surface polishing.
(5) after model is mounted on strut, its horizontal vertical degree error≤3 ' is detected, if failing the level for reaching requirement
The error of perpendicularity, it is necessary to the roll angle of model not adjusted, until horizontal vertical degree error reaches requirement.
(6) after model is installed, the sealing of pressure-measuring pipe road is detected.Specific requirement be suppress 105kPa~
130kPa, system read reaction time≤0.5s, stablize time >=3s；For negative pressure to 90kPa~5kPa, system reads the reaction time
≤0.5s.It is detected after having blown wind every time to according to step (6), reaches requirement and then continue to dry.If not up to find out
Existing gas leakage or clogging then need to open model and carry out malfunction elimination, overhaul pipeline.
(7) wind tunnel test is dried.
The comparison of this method and conventional scaling method measurement error is shown in Table 1.From table it can be concluded that using this method
The systematic measurement error of calibration improves a lot, and static pressure measurement error is reduced to 490Pa, Mach number measurement error by 3000Pa
It is reduced to 0.05 by 0.15, the angle of attack is reduced to 0.5 ° by 2.5 °, and yaw angle is reduced to 0.5 ° by 1.2 °.The measurement essence of system calibrating
Degree reaches technical requirement.
The comparison (Ma=2) of 1 this method of table and conventional scaling method measurement error
Embodiment 2
(1) supersonic wind tunnel jet pipe length is 2400mm, and sectional dimension is 1.2m × 1.2m, model length 871mm, mould
Type maximum gauge is 268mm.It is 800mm, distance of the angle of attack connector apart from nozzle exit that attack angle mechanism, which turns the heart apart from nozzle exit,
For 114mm.Testing the range of angles of attack estimated is α=- 12 °~12 °.
(2) the step wave structure of Mach number 3 and the position of upper and lower wave intersection point in supersonic nozzle are measured using stration technique,
Intersection position size X1=2300mm is measured, the coverage of step wave is determined with this.Since intersection position is apart from attack angle mechanism
Turn that the heart is closer, therefore the leading portion Delta Region (0mm~2300mm) of selected supersonic wind tunnel jet pipe is as embedded atmosphere data
The Flow Field in Wind Tunnel region of sensor-based system calibration is high 2=75.4 ° of base angle for calculating triangular domain with X1.The angle of attack is estimated according to test
α=- 12 ° of range~12 °, attack angle mechanism turn heart L3=800mm, step wave intersection point X1=1690mm, step wave Delta Region bottom
Angle α 2=70.5 ° and 3=39 ° of apex angle α；
(3) it is met the requirements when length L0=1300mm (the wherein Ma=3) of model contracting ratio Sc=0.8 and installation strut.
Strut lengths are attack angle mechanism connector size at a distance from model bottom.Ensure during the test, under all angle of attack states
Model can avoid step wave, and guarantee the point on model surface apart from step wave proximal most position >=100mm.
Processing experiment model and strut carry out wind tunnel test after installation.
Calibration result is shown in Table 2, it can be concluded that this scaling method, the measurement error precision of calibration are high from table.
The comparison (Ma=3) of 2 this method of table and conventional scaling method measurement error
The above, optimal specific embodiment only of the invention, but scope of protection of the present invention is not limited thereto,
In the technical scope disclosed by the present invention, any changes or substitutions that can be easily thought of by anyone skilled in the art,
It should be covered by the protection scope of the present invention.
The content that description in the present invention is not described in detail belongs to the well-known technique of professional and technical personnel in the field.
Claims (6)
1. a kind of Flush Airdata Sensing System scaling method, which comprises the steps of:
(1) parameter of supersonic wind tunnel, jet pipe length L are obtained1Mm, sectional dimension B1 mm×B1 Mm, test Mach number Ma are 2.0
~4.0, aircraft prototype length is L2Mm, equivalent diameter Dmm；Estimating range of angles of attack according to test demand is-α~α；
(2) using stration technique measurement Mach number 2~4 when, the position of the angle of up/down steps wave and upper and lower wave intersection point, test desk
Distance X of the rank wave intersection point apart from nozzle entry section1Mm, upper and lower wave and nozzle section constitute the cone in region, along wind-tunnel incoming flow
The median plane in direction is isosceles triangle, base angle2And apex angle α3；
(3) the ratio value Sc of modulus type contracting ratio is 1；
(4) work as X1<1/2L1When mm, jet pipe back segment is selected as test area, strut lengths L0Calculation formula are as follows:
Work as X1≥1/2L1Mm selectes leading portion test area, strut lengths L0Calculation formula are as follows:
L0≥(Sc.D/2)/tan((arctan(B1/(L1-L3))×180/π-α)×π/180)+L3-L4-100
It is L that wherein attack angle mechanism, which turns the heart apart from nozzle exit,3Mm, attack angle mechanism and strut connector are apart from nozzle exit L4mm；
(5) if the length L of calculated strut0It is not at effective range, then readjusts the ratio of step range of angles of attack He contracting ratio
Example value, takes Sc=Sc-0.1 or α=α -2, return step (4)；If L0In effective range, it is determined that α, Sc and L0；
(6) according to the dimensioned dummy vehicle after the ratio value Sc contracting ratio of contracting ratio, according to L0Process strut；
(7) Flush Airdata Sensing System is mounted on inside dummy vehicle, dummy vehicle is mounted on by strut
On attack angle mechanism；
(8) blasting is carried out in 2~4 range of Mach number, attack angle mechanism is run within the scope of-α to α, embedded big destiny
According to static pressure, Mach number, the angle of attack and the yaw angle of sensor-based system measurement dummy vehicle；Flush Airdata Sensing System is surveyed
Dummy vehicle static pressure, Mach number, the angle of attack and yaw angle and wind tunnel system measurement dummy vehicle static pressure, Mach
Number, the angle of attack compare with yaw angle, judge whether airdata sensing system measurement is accurate.
2. a kind of Flush Airdata Sensing System scaling method according to claim 1, which is characterized in that strut is long
Spend L0Effective range be greater than 0.1L1, and it is less than 0.6L1。
3. a kind of Flush Airdata Sensing System scaling method according to claim 1, which is characterized in that judge embedding
Whether accurate enter the measurement of formula airdata sensing system method particularly includes: static pressure error≤500Pa, Mach number error≤0.1,
Angle of attack error≤0.5 °, sideslip angle error≤0.5 ° then judge that Flush Airdata Sensing System measurement is accurate；If wherein
There is parameter error then more than threshold range to judge that the Flush Airdata Sensing System parameter measurement error precision is not up to set
Meter requires.
4. a kind of Flush Airdata Sensing System scaling method according to claim 1, which is characterized in that according to Sc
Dummy vehicle is processed, each dimension precision requirement in dummy vehicle surface reaches ± 0.02mm, dummy vehicle surface pressure tap
Status requirement: the pressure tap position circumferential direction deviation positioned at dummy vehicle vertex is less than 0.1mm, and the error of perpendicularity≤3 ' play wall
The position axial error of upper pressure tap is less than 0.1mm, and circumferential error is less than 0.1mm, after processing pressure tap, to dummy vehicle
Surface polishing.
5. a kind of Flush Airdata Sensing System scaling method according to claim 1, which is characterized in that aircraft
Horizontal gradient error≤3 ' after model is mounted on strut.
6. a kind of Flush Airdata Sensing System scaling method according to claim 4, which is characterized in that pressure tap
Interior installation pressure-measuring pipe road, load cell go out to be mounted on dummy vehicle surface through pressure-measuring pipe pass, and dummy vehicle installs
Cheng Hou is pressurized to 105kPa~130kPa, Flush Airdata Sensing System reaction time≤0.5s, when stablizing to pressure tap
Between >=3s；Negative pressure is to 90kPa~5kPa, system reaction time≤0.5s.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201611180132.2A CN106840573B (en) | 2016-12-19 | 2016-12-19 | A kind of Flush Airdata Sensing System scaling method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201611180132.2A CN106840573B (en) | 2016-12-19 | 2016-12-19 | A kind of Flush Airdata Sensing System scaling method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN106840573A CN106840573A (en) | 2017-06-13 |
| CN106840573B true CN106840573B (en) | 2019-04-30 |
Family
ID=59140596
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201611180132.2A Active CN106840573B (en) | 2016-12-19 | 2016-12-19 | A kind of Flush Airdata Sensing System scaling method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN106840573B (en) |
Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109186934B (en) * | 2018-11-15 | 2023-09-26 | 中国航空工业集团公司沈阳空气动力研究所 | Space near-field acoustic explosion characteristic high-precision measuring device |
| CN111157759B (en) * | 2019-12-24 | 2022-11-04 | 太原航空仪表有限公司 | Fixed differential pressure type attack angle sensor and use method |
| CN111605724B (en) * | 2020-05-22 | 2022-08-12 | 成都飞机工业（集团）有限责任公司 | Simulation test method for atmospheric data system |
| CN112505048B (en) * | 2021-02-02 | 2021-05-18 | 北京理工大学 | Testing device and method for calibrating hydrogen concentration |
| CN114061896B (en) * | 2021-11-17 | 2023-04-25 | 中国空气动力研究与发展中心超高速空气动力研究所 | Method for processing internal resistance measurement test data of lifting body ventilation model |
| CN116380396B (en) * | 2023-05-30 | 2023-08-15 | 中国航空工业集团公司沈阳空气动力研究所 | Continuous wind tunnel aircraft atmospheric data system identification test system and method |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102520726A (en) * | 2011-12-19 | 2012-06-27 | 南京航空航天大学 | Estimation method of atmosphere angle of attack and angle of sideslip in high-angle-of-attack flight status |
| CN102607639A (en) * | 2012-02-24 | 2012-07-25 | 南京航空航天大学 | BP (Back Propagation) neural network-based method for measuring air data in flight state with high angle of attack |
| US8930062B2 (en) * | 2009-04-21 | 2015-01-06 | Indian Space Research Organisation | System and method for detecting and isolating faults in pressure sensing of flush air data system (FADS) |
| CN104318107A (en) * | 2014-10-27 | 2015-01-28 | 中国运载火箭技术研究院 | Method for acquiring high-precision atmosphere data of aircraft flying across atmospheric layer |
| CN105066994A (en) * | 2015-08-21 | 2015-11-18 | 中国运载火箭技术研究院 | Data fusion method for flush air data system and inertial navigation system |
| CN105628325A (en) * | 2014-10-29 | 2016-06-01 | 北京临近空间飞行器***工程研究所 | Real-time high-precision acquiring method for conical surface pneumatic pressure field |
-
2016
- 2016-12-19 CN CN201611180132.2A patent/CN106840573B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8930062B2 (en) * | 2009-04-21 | 2015-01-06 | Indian Space Research Organisation | System and method for detecting and isolating faults in pressure sensing of flush air data system (FADS) |
| US9211961B2 (en) * | 2009-04-21 | 2015-12-15 | Indian Space Research Organisation | System and method for detecting and isolating faults in pressure sensing of flush air data system (FADS) |
| CN102520726A (en) * | 2011-12-19 | 2012-06-27 | 南京航空航天大学 | Estimation method of atmosphere angle of attack and angle of sideslip in high-angle-of-attack flight status |
| CN102607639A (en) * | 2012-02-24 | 2012-07-25 | 南京航空航天大学 | BP (Back Propagation) neural network-based method for measuring air data in flight state with high angle of attack |
| CN104318107A (en) * | 2014-10-27 | 2015-01-28 | 中国运载火箭技术研究院 | Method for acquiring high-precision atmosphere data of aircraft flying across atmospheric layer |
| CN105628325A (en) * | 2014-10-29 | 2016-06-01 | 北京临近空间飞行器***工程研究所 | Real-time high-precision acquiring method for conical surface pneumatic pressure field |
| CN105066994A (en) * | 2015-08-21 | 2015-11-18 | 中国运载火箭技术研究院 | Data fusion method for flush air data system and inertial navigation system |
Non-Patent Citations (2)
| Title |
|---|
| 嵌入式大气数据传感***标定试验研究;秦永明 等;《第九届全国实验流体力学学术会议论文》;20130706;502-508 |
| 嵌入式大气数据传感***风洞标定试验研究;秦永明 等;《空气动力学学报》;20150831;第33卷(第14期);487-492 |
Also Published As
| Publication number | Publication date |
|---|---|
| CN106840573A (en) | 2017-06-13 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106840573B (en) | A kind of Flush Airdata Sensing System scaling method | |
| Schulz et al. | Experimental investigation of the three-dimensional flow in an annular compressor cascade | |
| CN103424225B (en) | A kind of method of testing rotatable parts sound amount of unbalance | |
| US8718971B2 (en) | System for determining the airspeed of an aircraft | |
| CN104792323B (en) | Dynamic Water level and attitude angle computational methods | |
| CN209841336U (en) | Temporary impulse type supersonic wind tunnel sonic boom measurement test device | |
| CN103698081A (en) | L-shaped differential pressure type probe | |
| CN104197955B (en) | The micro- disturbance torque measuring method of full-automatic three-axis air-bearing table | |
| AU2012101869A4 (en) | A convenient connection system for pressure measuring of the control surface | |
| US20040011124A1 (en) | Method for determining aerodynamic parameters and method for detecting failure of a probe used for determining aerodynamic parameters | |
| Chou et al. | Instability and transition research in a Mach-6 quiet tunnel | |
| CN102435156B (en) | Large cylindrical workpiece dimension and geometric error measurement method | |
| CN112556917B (en) | Method for measuring pressure by using pressure measuring device | |
| CN104482838A (en) | Propeller blade measuring technology | |
| Ainsworth et al. | The development of fast response aerodynamic probes for flow measurements in turbomachinery | |
| CN111498141B (en) | Method and device for realizing real-time monitoring of airflow angle based on micro probe | |
| CN113551869A (en) | Probe for measuring two-dimensional dynamic boundary layer of rotating-static end wall of multistage gas compressor | |
| Wu et al. | A robust calibration method for seven-hole pressure probes | |
| Cartieri et al. | Using CFD to calculate support interference effects | |
| CN105588512A (en) | Measurement device and method for measuring dimension of quasi-rectangular tunnel segment component | |
| Anscombe et al. | Some comments on high-lift testing in wind tunnels with particular reference to jet-blowing models | |
| CN105021094A (en) | Small-aspect-ratio missile horizontal measuring method | |
| CN104608876A (en) | Digitized dock building method | |
| CN111157759B (en) | Fixed differential pressure type attack angle sensor and use method | |
| You et al. | Recent activities on flow quality assessment at the European transonic windtunnel |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |