CN114166460B - Aircraft air inlet passage test device and system and hot gas anti-icing test stability judgment method - Google Patents
Aircraft air inlet passage test device and system and hot gas anti-icing test stability judgment method Download PDFInfo
- Publication number
- CN114166460B CN114166460B CN202210129678.4A CN202210129678A CN114166460B CN 114166460 B CN114166460 B CN 114166460B CN 202210129678 A CN202210129678 A CN 202210129678A CN 114166460 B CN114166460 B CN 114166460B
- Authority
- CN
- China
- Prior art keywords
- icing
- temperature
- lip
- temperature measuring
- hot gas
- 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
Images
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
- G01M9/02—Wind tunnels
- G01M9/04—Details
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/60—Testing or inspecting aircraft components or systems
-
- 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
- G01M9/06—Measuring arrangements specially adapted for aerodynamic testing
Landscapes
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- General Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- Manufacturing & Machinery (AREA)
- Transportation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention is suitable for the technical field of wind tunnel tests and provides an aircraft air inlet channel test device, an aircraft air inlet channel test system and a hot air anti-icing test stability judging method. The aircraft air inlet passage test device, the aircraft air inlet passage test system and the hot gas anti-icing test stability judging method provided by the invention have the effects of high anti-icing and deicing efficiency and hot gas resource saving.
Description
Technical Field
The invention relates to the technical field of wind tunnel tests, in particular to an aircraft air inlet channel test device, an aircraft air inlet channel test system and a hot gas anti-icing test stability judgment method.
Background
In the prior art, the icing on the outer shape surface of the aircraft changes the aerodynamic shape of the aircraft, so that the flight performance is deteriorated, and the icing is one of main dangerous sources of safety accidents of the aircraft. Therefore, the easy-to-freeze part of the aircraft must be protected by an anti-icing and deicing system, different parts are relatively independent, icing conditions are different, for example, an aircraft engine can supply a heat source by itself, the adopted icing protection mode is hot air anti-icing and deicing, and other parts mainly comprise electric heating anti-icing and deicing, mechanical anti-icing and thermal coupling anti-icing and deicing and the like. Considering the icing environment and self conditions of different parts, applying an appropriate anti-icing and anti-icing means to the different parts is the important research point of the airplane icing protection on the premise of limited airborne energy.
The hot gas deicing and preventing technology is characterized in that hot gas resources generated by an engine are utilized to protect the key positions of the wings and the engine of an airplane, but the amount of the extracted hot gas is related to the working state of the engine. At present, the hot gas anti-icing technology is difficult to realize that the protection component is not interfered by icing while the stable work of an engine is ensured, and the reasonable utilization and design of hot gas air-entraining resources are difficult to realize, so that the anti-icing efficiency is low, or the waste of the hot gas air-entraining resources is caused.
In summary, the technical problems of the prior art are as follows:
1. the application of an ice prevention and control section suitable for different components in the prior art is the key point of research on the icing protection of an airplane;
2. the hot air deicing technology in the prior art can not realize reasonable utilization and design of hot air entraining resources, so that deicing efficiency is low, or the hot air entraining resources are wasted.
Disclosure of Invention
The invention aims to provide an aircraft air inlet channel test device and system with high deicing efficiency and hot gas resource saving and a hot gas deicing test stability judgment method.
The invention provides an aircraft air inlet channel test device, wherein a lip is arranged at the windward end of an air inlet channel, a hot air cavity is arranged in the lip, a plurality of monitoring structures are arranged on the side wall of the lip along the extension direction of a front edge, each monitoring structure comprises a temperature measuring component and an icing measuring component, the temperature measuring components and the icing measuring components are positioned on the same cross section of the front edge, the temperature measuring components are used for acquiring the temperature of the outer surface of the lip, and the icing measuring components are used for acquiring the icing information of the outer surface of the lip.
Further, the temperature measurement component comprises a middle temperature measurement point, an upper temperature measurement point and a lower temperature measurement point, the middle temperature measurement point is arranged on the front edge of the front edge, the upper temperature measurement point is arranged on the upper side wall of the lip, and the lower temperature measurement point is arranged on the lower side wall of the lip.
Further, the number of the upper temperature measuring points is equal to that of the lower temperature measuring points.
Furthermore, the temperature measuring component is embedded in a skin, and the skin is arranged on the outer surface of the lip.
Further, the icing measuring assembly comprises a receiving end and a transmitting end, the receiving end is arranged between the middle temperature measuring point and the upper temperature measuring point, and the transmitting end is arranged between the middle temperature measuring point and the lower temperature measuring point; or the receiving end is arranged between the middle temperature measuring point and the lower temperature measuring point, and the transmitting end is arranged between the middle temperature measuring point and the upper temperature measuring point.
The invention provides an aircraft air inlet channel test system in a second aspect, which comprises a control module and a data monitoring module which are connected with each other, wherein the data monitoring module comprises the aircraft air inlet channel test device, and the data monitoring module is connected with a monitoring structure.
Furthermore, the data monitoring module is connected with the temperature measurement component and the icing measurement component.
The third aspect of the invention provides a stability judging method for a hot gas anti-icing test, which comprises the following steps:
step S10: starting the aircraft air inlet channel test system;
step S20: controlling hot gas parameters until temperature difference values of preset temperature measuring points in a preset time period are smaller than a preset temperature difference value and icing information is not monitored by an icing measuring assembly in the preset time period, and judging that the test condition is stable; the preset temperature measuring points comprise a middle temperature measuring point, at least part of upper temperature measuring points and at least part of lower temperature measuring points.
Further, the preset temperature measuring points are all the temperature measuring points.
Further, the hot gas parameters include hot gas temperature and/or hot gas flow.
In summary, the present invention has at least the following technical effects:
1. in order to efficiently and reasonably use hot air energy, the critical anti-icing hot air quantity is searched through an icing wind tunnel test aiming at different icing environments, and the critical anti-icing condition obtains data through a plurality of monitoring structures arranged on the lip, so that the reasonable utilization and design of hot air entraining resources are realized, the anti-icing efficiency is improved, and the hot air entraining resources are saved;
2. the critical anti-icing condition can be comprehensively embodied by the existence of icing and the surface temperature, the monitoring structure is a temperature measuring component and an icing measuring component, the temperature measuring component is used for acquiring the temperature of the outer surface of the lip, the icing measuring component is used for acquiring the icing information of the outer surface of the lip, and the temperature and the icing information of the outer surface of the lip are further fed back to further optimize hot gas entraining resources as a reference basis;
3. according to the invention, the middle temperature measuring point is arranged at the front edge to obtain the temperature data of the foremost end of the windward side of the lip, the upper temperature measuring point is arranged on the upper side wall of the lip to obtain the temperature data of the upper surface of the windward side of the lip, the lower temperature measuring point is arranged on the lower side wall of the lip to obtain the temperature data of the lower surface of the windward side of the lip, and the temperature data of the foremost end, the upper surface and the lower surface of the windward side of the lip are comprehensively reflected to accurately and comprehensively reflect the temperature of the surface of the lip, so that an important data basis is provided for realizing a critical anti-icing condition.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings needed to be used in the embodiments of the present invention or in the description of the prior art will be briefly described below, and it is obvious that the drawings described below are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is a schematic view of an aircraft air scoop test apparatus of the present invention;
FIG. 2 is a schematic view of a monitoring structure on a lip section in the present invention;
FIG. 3 is a schematic view of the distribution of the lip circumference monitoring structure according to the present invention;
FIG. 4 is a schematic diagram of an aircraft inlet test system according to the present invention;
FIG. 5 is a schematic diagram of a stability judging method in the hot gas anti-icing test of the present invention.
The device comprises 100-an aircraft air inlet test device, 110-a lip, 111-a leading edge, 112-a hot air cavity, 113-a skin, 120-an icing measurement assembly, 121-a receiving end, 122-a transmitting end and 130-a temperature measurement assembly.
Detailed Description
The following description provides many different embodiments, or examples, for implementing different features of the invention. The particular examples set forth below are illustrative only and are not intended to be limiting.
In the description of the present invention, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like, indicate orientations or positional relationships based on those shown in the drawings, are merely for convenience of description and simplicity of description, and do not indicate or imply that the referenced devices or elements must have a particular orientation, be constructed in a particular orientation, and be operated, and thus, are not to be construed as limiting the present invention. Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, features defined as "first", "second", may explicitly or implicitly include one or more of the described features. In the description of the present invention, "a plurality" means a plurality or more unless specifically limited otherwise.
In the description of the present invention, it should be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection: may be mechanically connected, may be electrically connected or may be in communication with each other; the connection may be direct or indirect through an intermediate medium, and may be a connection between a plurality of elements or an interaction relationship between a plurality of elements. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In the present invention, unless otherwise expressly stated or limited, "above" or "below" a first feature means that the first and second features are in direct contact, or that the first and second features are not in direct contact but are in contact with each other via another feature therebetween. Also, the first feature being "on," "above" and "over" the second feature includes the first feature being directly on and obliquely above the second feature, or merely indicating that the first feature is at a higher level than the second feature. A first feature being "under," "below," and "beneath" a second feature includes the first feature being directly under and obliquely below the second feature, or simply meaning that the first feature is at a lesser elevation than the second feature.
The present invention may repeat reference numerals and/or letters in the various examples, such repetition is for the purpose of simplicity and clarity and does not in itself dictate a relationship between the various embodiments and/or configurations discussed. In addition, the present invention provides examples of various specific processes and materials, but one of ordinary skill in the art may recognize applications of other processes and/or use of other materials.
The first embodiment is as follows:
as shown in fig. 1, a first embodiment of the present invention provides an aircraft inlet testing apparatus 100, a lip 110 is disposed at an windward end of the inlet, a hot air chamber 112 is disposed inside the lip 110, a plurality of monitoring structures are disposed on a side wall of the lip 110 along an extending direction of a leading edge 111, each monitoring structure includes a temperature measuring component 130 and an icing measuring component 120, the temperature measuring component 130 and the icing measuring component 120 are located on a same cross section of the leading edge 111, the temperature measuring component 130 is used for obtaining a temperature of an outer surface of the lip 110, and the icing measuring component 120 is used for obtaining icing information of the outer surface of the lip 110.
The applicant finds out through long-term practical experience that: the reasonably designed bleed air flow is the key point of hot air anti-icing design, namely, the stable operation of an engine is guaranteed, and meanwhile, components are protected from being interfered by icing. The hot-air deicing design verification of the airplane component mainly depends on an icing wind tunnel at present, the deicing effect of the hot-air deicing design verification is embodied in the surface temperature change of the component and the existence of visible ice, and the surface temperature change is usually obtained through the feedback of a temperature sensor; the existence of visible ice is difficult to capture, the difficulty in distinguishing the ice coating on the surface of a human eye real-time observation model is high, the test design condition is close to a critical point, the conditions of micro icing and rapid ice falling exist, and the test result is influenced, so that the acquisition of the visible ice is very necessary for searching the critical condition for preventing and removing the ice.
Therefore, the applicant is provided with the icing measurement component 120 to obtain icing information of the surface of the observation model and feed back the icing condition of the surface of the observation model; as shown in fig. 1 and 2, an aircraft air inlet duct generally has a circular structure, and the temperatures and icing information at different circular cross-sectional positions in an icing wind tunnel are different, in this embodiment, the temperature measurement component 130 and the icing measurement component 120 are arranged on the same cross section of the leading edge 111, so that the temperature information obtained by the temperature measurement component 130 and the icing information obtained by the icing measurement component 120 can be matched with each other, that is, the temperature information and the icing information on the same interface can be integrated to reflect whether the cross section reaches a critical anti-icing condition.
The present embodiment utilizes the surface temperature of the thermometric assembly 130, and the icing measuring assembly 120 measures whether visible ice is present. The critical anti-icing condition is that under a certain environment, no icing is just on the surface of the object, and the temperature of the surface of the object just reaches a critical temperature value. The critical temperature value at normal temperature and normal pressure is 0 ℃.
In this embodiment, an aircraft inlet is used as a test model, and a lip 110 having a hot air chamber 112 therein is disposed at the front end of the test model. However, in practical applications, any test model similar to the air intake duct and the structure of which the air intake end is provided with the lip 110 similar to the above lip including the hot air chamber 112 can be taken as a modified object of the present embodiment.
In order to efficiently and reasonably use hot air energy, the hot air quantity of critical anti-icing is searched through an icing wind tunnel test aiming at different icing environments, and data is obtained by a plurality of monitoring structures arranged on the lip 110 under the critical anti-icing condition, so that the reasonable utilization and design of hot air entraining resources are realized, the anti-icing efficiency is improved, and the hot air entraining resources are saved.
As shown in fig. 3, a common aircraft inlet model is in a regular circular structure, and in order to obtain a critical anti-icing condition of the inlet lip 110, during a test, it is required to measure whether visible ice exists in each direction of the inlet lip 110, and obtain a temperature distribution condition of the inlet lip 110 in each direction. The circumferential direction of the lip 110 can be equally divided into 4 sections by equally dividing the circular structure into a plurality of parts, such as 4 parts (dotted line part) in fig. 3, and the temperature measuring unit 130 and the icing measuring unit 120 are disposed on each interface. The more the number of the divided parts is, the more accurate the obtained data is, the more the experimental results are, but the more the temperature measuring component 130 and the icing measuring component 120 are required, the cost is increased, the probability of generating dead spots is also high, and the accuracy of measurement is further influenced.
At this time, when the aircraft inlet model is a regular circular structure, the icing protection conditions on two sides symmetrical in the vertical direction during the test are relatively consistent, so that only half of the lip 110 in the circumferential direction can be divided in the vertical direction, and a plurality of equally divided monitoring structures are arranged only in the half of the lip 110 in the circumferential direction, as shown by the dotted line and the dashed-dotted line on the left side in fig. 3. Thus, not only can cost be reduced, but also accurate test data can be obtained. The number of divided parts is not particularly limited.
The critical anti-icing condition can be comprehensively embodied by the existence of icing and the surface temperature, the monitoring structure is a temperature measuring component 130 and an icing measuring component 120, the temperature measuring component 130 is used for obtaining the temperature of the outer surface of the lip 110, the icing measuring component 120 is used for obtaining the icing information of the outer surface of the lip 110, and the temperature and the icing information of the outer surface of the lip 110 are further fed back to further serve as a reference basis for optimizing hot gas entraining resources.
Further, as shown in FIG. 2, the thermometric assembly 130 includes a middle thermometric pointT 0 Upper temperature measuring pointT f And lower temperature measuring pointT b The middle temperature measuring pointT 0 Is arranged at the front edge 111, the upper temperature measuring pointT f Is arranged on the upper side wall of the lip 110, and has a lower temperature measuring pointT b Is disposed on the lower sidewall of the lip 110.
A plurality of temperature measuring points are arranged according to the section contour of the lip 110, each temperature measuring point is provided with a temperature sensor, the front edge 111 is used as a boundary, the upper side of the front edge 111 is an upper temperature measuring point, and the lower side of the front edge 111 is a lower temperature measuring point.
The middle temperature measuring point is arranged on the front edge 111 to obtain the temperature data of the foremost end of the windward side of the lip 110, the upper temperature measuring point is arranged on the upper side wall of the lip 110 to obtain the temperature data of the upper surface of the windward side of the lip 110, the lower temperature measuring point is arranged on the lower side wall of the lip 110 to obtain the temperature data of the lower surface of the windward side of the lip 110, and the temperature data of the foremost end, the upper surface and the lower surface of the windward side are comprehensively reflected to accurately and comprehensively reflect the temperature of the surface of the lip 110, so that an important data basis is provided for realizing a critical anti-icing condition.
Further, the number of the upper temperature measuring points is equal to that of the lower temperature measuring points.
Further, the temperature measuring component 130 is embedded in the skin 113, and the skin 113 is disposed on the outer surface of the lip 110.
Further, as shown in fig. 2 and 3, the icing measuring assembly 120 includes a receiving end 121 and an emitting end 122, the receiving end 121 is disposed between the middle temperature measuring point and the upper temperature measuring point, and the emitting end 122 is disposed between the middle temperature measuring point and the lower temperature measuring point; or the receiving end 121 is arranged between the middle temperature measuring point and the lower temperature measuring point, and the transmitting end 122 is arranged between the middle temperature measuring point and the upper temperature measuring point.
As shown in fig. 3, the ice condensation measuring assembly 120 is embedded on the outer surface of the skin 113 in the present embodiment. The transmitting end 122 and the receiving end 121 of the icing measuring assembly 120 are respectively located at two sides of the leading edge 111, and if icing covers the transmitting end 122 and the receiving end 121, the transmitting end 122 generates an optical signal and transmits the optical signal to the receiving end 121 through the ice, so that the existence of visible ice is monitored. While the icing measurement assembly 120 is typically positioned near the leading edge 111 because of the susceptibility to icing near the leading edge 111. The ice measurement component 120 may be a fiber optic ice sensor.
Example two:
as shown in fig. 4, a second embodiment of the present invention provides an aircraft air inlet testing system, which includes a control module and a data monitoring module that are connected to each other, where the data monitoring module includes the aircraft air inlet testing apparatus 100, and the data monitoring module is connected to a monitoring structure.
The control module is used for adjusting and controlling hot gas parameters provided for the hot gas cavity 112 in the lip 110 according to the data provided by the data monitoring module; the data monitoring module is used for acquiring monitored data in the monitoring structure, and the data at least comprises temperature data and icing information.
Further, the monitoring structure includes a temperature measurement component 130 and an icing measurement component 120, and the data monitoring module is connected with the temperature measurement component 130 and the icing measurement component 120.
Example three:
as shown in fig. 5, a third embodiment of the present invention provides a stability judging method for a hot gas anti-icing test, including the following steps:
step S10: starting the aircraft air inlet channel test system;
step S20: controlling hot gas parameters until temperature difference values of preset temperature measuring points in a preset time period are smaller than a preset temperature difference value and the icing measuring component 120 does not monitor icing information in the preset time period, and judging that the test condition is stable; the preset temperature measuring points comprise a middle temperature measuring point, at least part of upper temperature measuring points and at least part of lower temperature measuring points.
As shown in FIG. 2, the thermometric assembly 130 includes a middle thermometric point according to the cross-sectional position of the lip 110T 0 Upper temperature measuring pointT f And lower temperature measuring pointT b And numbering the upper temperature measuring point or the lower temperature measuring point according to the distance from the upper temperature measuring point or the lower temperature measuring point to the front edge 111 from small to large, such as 1, 2, 3, the
、
、
、...、
The lower temperature measuring point is
、
、
、...、
。
At least part of the upper temperature measuring points and at least part of the lower temperature measuring points are used as preset temperature measuring points, so that the burden of data processing is reduced, and the calculation rate is improved. In general, the upper temperature measurement point and the lower temperature measurement point are selected to preferentially select the temperature measurement point having a relatively large distance from the leading edge 111 as the preset temperature measurement point.
When the preset time period is 12s, the preset temperature difference value is 5 ℃, and the preset temperature measuring point is a middle temperature measuring pointT 0 The nth upper temperature measuring point
And the nth lower temperature measuring point
Time, middle temperature measuring pointT 0 The highest temperature value of the middle temperature measuring point occurs at the 8 th time (t =1 s) to the 12 th time (t =12 s), that is, at the 1 st time (t =1 s)
(ii) a The lowest temperature value of the middle temperature measuring point occurs at the 1 st moment, namely
，
I.e. not meeting critical anti-icing conditions.
Meanwhile, when the preset time period is 12s, the preset temperature difference value is 5 ℃, and the preset temperature measuring points are the middle temperature measuring point T0, the nth upper temperature measuring point and the nth lower temperature measuring point, only the highest temperature value of the middle temperature measuring point
Minimum temperature value of middle temperature measuring point
I.e. by
(ii) a And the highest temperature value of the nth upper temperature measuring point
The lowest temperature value of the nth upper temperature measuring point
I.e. by
(ii) a Meanwhile, the highest temperature value of the nth lower temperature measuring point
The lowest temperature value of the nth lower temperature measuring point
I.e. by
(ii) a At this time, middle temperature measurementDotT 0 The nth upper temperature measuring point
And the nth lower temperature measuring point
And the temperature difference values in the preset time period are all smaller than the preset temperature difference value, namely the temperature difference values of the preset temperature measuring points in the preset time period are all smaller than the preset temperature difference value. And, when the icing information is not monitored by the icing measuring component 120 within the preset time of 12s, the test condition meeting the critical anti-icing is obtained, i.e. the test condition is stable.
The stability of the hot gas anti-icing test can be judged before and after spraying, before spraying, the surface of the lip 110 cannot be iced, icing information obtained by the icing measuring component 120 can be ignored in the actual production process, and the icing information obtained by the icing measuring component 120 can be taken into consideration so as to ensure the test accuracy. After the spraying is started, ice may be formed on the surface of lip 110, and the ice formation information is required as a precondition for determining whether the test conditions are stable.
With the advance of time, the current time is the last time of the preset time period in each calculation, and the time of the preset time period is traced forwards, namely the time is the data basis for judging the temperature difference value of the preset temperature measuring point.
Further, the preset temperature measuring points are all the temperature measuring points. In order to improve the measurement accuracy, all temperature measuring points can be selected as preset temperature measuring points.
Further, the hot gas parameters include hot gas temperature and/or hot gas flow.
In order to efficiently and reasonably use hot air energy, the invention searches the critical anti-icing hot air quantity through an icing wind tunnel test aiming at different icing environments, and the critical anti-icing condition is reflected by whether the lip 110 is iced or not and the surface temperature, thereby being used as a reference basis for designing and optimizing a hot air anti-icing system.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents and improvements made within the spirit and principle of the present invention are intended to be included within the scope of the present invention.
Claims (10)
1. A stability judging method for a hot gas anti-icing test is characterized by comprising the following steps: step S10: starting an aircraft air inlet channel test system;
step S20: controlling hot gas parameters until temperature difference values of preset temperature measuring points in a preset time period are smaller than a preset temperature difference value, and judging that the test condition is stable when the icing measuring component (120) does not monitor icing information in the preset time period; the preset temperature measuring points comprise a middle temperature measuring point, at least part of upper temperature measuring points and at least part of lower temperature measuring points;
the aircraft air inlet test system comprises a control module and a data monitoring module which are connected with each other, the data monitoring module comprises an aircraft air inlet test device (100), and the data monitoring module is connected with a monitoring structure;
in aircraft intake duct test device (100), the windward end of intake duct is provided with lip (110), lip (110) inside is provided with hot air chamber (112), be provided with a plurality of monitoring structures along leading edge (111) extending direction on the lateral wall of lip (110), the monitoring structure includes temperature measurement component (130) and icing measurement component (120), temperature measurement component (130) with icing measurement component (120) is located the same cross-section of leading edge (111), temperature measurement component (130) are used for obtaining the temperature of lip (110) surface, icing measurement component (120) are used for obtaining the information that freezes of lip (110) surface.
2. The method for determining stability in the hot gas anti-icing test according to claim 1, wherein the preset temperature measuring points are all the temperature measuring points.
3. The method for determining stability of hot gas anti-icing test according to claim 1, wherein the hot gas parameters comprise hot gas temperature and/or hot gas flow.
4. An aircraft inlet test device (100) for performing the hot gas anti-icing test stability judging method according to any one of claims 1 to 3, wherein a lip (110) is arranged at the windward end of the inlet, a hot gas chamber (112) is arranged inside the lip (110), a plurality of monitoring structures are arranged on the side wall of the lip (110) along the extension direction of the front edge (111), each monitoring structure comprises a temperature measuring component (130) and an icing measuring component (120), the temperature measuring components (130) and the icing measuring components (120) are located on the same cross section of the front edge (111), the temperature measuring components (130) are used for obtaining the temperature of the outer surface of the lip (110), and the icing measuring components (120) are used for obtaining the icing information of the outer surface of the lip (110).
5. The aircraft inlet test device (100) according to claim 4, wherein the temperature measurement component (130) comprises a middle temperature measurement point, an upper temperature measurement point and a lower temperature measurement point, the middle temperature measurement point is arranged on the front edge (111), the upper temperature measurement point is arranged on the upper side wall of the lip (110), and the lower temperature measurement point is arranged on the lower side wall of the lip (110).
6. Aircraft inlet test unit (100) according to claim 5, characterised in that the number of upper temperature measurement points and lower temperature measurement points is equal.
7. The aircraft inlet test device (100) of claim 5, wherein the temperature measurement component (130) is embedded in a skin (113), the skin (113) being disposed on an outer surface of the lip (110).
8. The aircraft inlet test device (100) according to claim 5, wherein the icing measuring assembly (120) comprises a receiving end (121) and a transmitting end (122), the receiving end (121) being arranged between the middle temperature point and the upper temperature point, the transmitting end (122) being arranged between the middle temperature point and the lower temperature point; or the receiving end (121) is arranged between the middle temperature measuring point and the lower temperature measuring point, and the transmitting end (122) is arranged between the middle temperature measuring point and the upper temperature measuring point.
9. Aircraft air inlet test system, characterized in that it comprises a control module and a data monitoring module connected to each other, said data monitoring module comprising an aircraft air inlet test device (100) according to one of claims 4 to 8, said data monitoring module being connected to a monitoring structure.
10. The aircraft inlet testing system of claim 9, wherein the data monitoring module is coupled to the temperature measurement assembly (130) and the icing measurement assembly (120).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210129678.4A CN114166460B (en) | 2022-02-11 | 2022-02-11 | Aircraft air inlet passage test device and system and hot gas anti-icing test stability judgment method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210129678.4A CN114166460B (en) | 2022-02-11 | 2022-02-11 | Aircraft air inlet passage test device and system and hot gas anti-icing test stability judgment method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN114166460A CN114166460A (en) | 2022-03-11 |
| CN114166460B true CN114166460B (en) | 2022-04-19 |
Family
ID=80489799
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210129678.4A Active CN114166460B (en) | 2022-02-11 | 2022-02-11 | Aircraft air inlet passage test device and system and hot gas anti-icing test stability judgment method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114166460B (en) |
Families Citing this family (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114757109B (en) * | 2022-06-14 | 2022-08-23 | 中国空气动力研究与发展中心低速空气动力研究所 | Method and system for testing relation between internal and external icing parameters of air inlet channel |
| CN114771871B (en) * | 2022-06-14 | 2022-10-04 | 中国空气动力研究与发展中心低速空气动力研究所 | Method for debugging icing sensor of air inlet channel, sensor and control system |
| CN115575083B (en) * | 2022-12-12 | 2023-03-21 | 中国空气动力研究与发展中心低速空气动力研究所 | Verification device and verification method for control strategy of air bag controller |
| CN115599144A (en) * | 2022-12-12 | 2023-01-13 | 中国空气动力研究与发展中心低速空气动力研究所（Cn） | Dynamic temperature feedback electric heating anti-icing method and device for air inlet passage |
Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103308271A (en) * | 2012-03-13 | 2013-09-18 | 波音公司 | Supercooled large drop icing condition simulation system |
| EP2703291A2 (en) * | 2012-08-29 | 2014-03-05 | The Boeing Company | Inlet icing protection simulation system |
| CN104457078A (en) * | 2014-11-05 | 2015-03-25 | 大连理工大学 | Airplane engine blade frozen ice making system |
| CN107985607A (en) * | 2017-11-02 | 2018-05-04 | 成都飞机工业（集团）有限责任公司 | A kind of unmanned plane icing protective device |
| CN108691704A (en) * | 2017-04-10 | 2018-10-23 | 清华大学 | Engine intake icing detecting system and deicing system |
| CN109522627A (en) * | 2018-11-01 | 2019-03-26 | 西安电子科技大学 | Fan blade icing prediction technique based on SCADA data |
| CN109661348A (en) * | 2016-08-22 | 2019-04-19 | 挪威科技大学 | Icing control system |
| CN110702419A (en) * | 2019-10-11 | 2020-01-17 | 中国直升机设计研究所 | Anti-icing conformance test system and method for engine air inlet system |
| CN111655990A (en) * | 2017-11-14 | 2020-09-11 | 联合发动机制造集团股份公司 | Method for controlling an anti-icing system for an aircraft gas turbine engine air intake |
Family Cites Families (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2898939B1 (en) * | 2006-03-22 | 2008-05-09 | Snecma Sa | SYSTEM FOR DEFROSTING A TURBOMOTEUR INPUT CONE FOR AIRCRAFT |
| CN102313510B (en) * | 2010-07-02 | 2013-09-11 | 中国商用飞机有限责任公司 | Image icing detector |
| US20140014776A1 (en) * | 2012-07-13 | 2014-01-16 | Kelly Aerospace Thermal Systems Llc | System containing an electric heating element and method for installation and use thereof |
| CN103047010B (en) * | 2012-11-28 | 2015-04-22 | 中国商用飞机有限责任公司 | Anti-icing system and anti-icing control method for inlet lip of engine |
| US20160011105A1 (en) * | 2014-06-25 | 2016-01-14 | Timothy J. Bencic | Light Extinction Tomography for Measurement of Ice Crystals and Other Small Particles |
| CN105644791B (en) * | 2015-08-31 | 2018-06-22 | 中国商用飞机有限责任公司 | Icing detection system and aircraft with icing system |
| CN107271134B (en) * | 2017-06-28 | 2023-04-07 | 中国空气动力研究与发展中心低速空气动力研究所 | High-precision simulation method for icing wind tunnel hot gas anti-icing test |
| US11125157B2 (en) * | 2017-09-22 | 2021-09-21 | The Boeing Company | Advanced inlet design |
| CN107560822B (en) * | 2017-10-31 | 2020-05-05 | 中国商用飞机有限责任公司 | Anti-icing wind tunnel test equipment for front edge of engine air inlet and test method thereof |
| CN111794862B (en) * | 2020-07-15 | 2021-10-22 | 山东大学 | Deicing device and deicing method for air inlet lip of aircraft engine |
| CN112197932B (en) * | 2020-12-07 | 2021-03-12 | 中国空气动力研究与发展中心低速空气动力研究所 | Surface temperature correction method and measurement method for airplane anti-icing and deicing test model |
-
2022
- 2022-02-11 CN CN202210129678.4A patent/CN114166460B/en active Active
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103308271A (en) * | 2012-03-13 | 2013-09-18 | 波音公司 | Supercooled large drop icing condition simulation system |
| EP2703291A2 (en) * | 2012-08-29 | 2014-03-05 | The Boeing Company | Inlet icing protection simulation system |
| CN104457078A (en) * | 2014-11-05 | 2015-03-25 | 大连理工大学 | Airplane engine blade frozen ice making system |
| CN109661348A (en) * | 2016-08-22 | 2019-04-19 | 挪威科技大学 | Icing control system |
| CN108691704A (en) * | 2017-04-10 | 2018-10-23 | 清华大学 | Engine intake icing detecting system and deicing system |
| CN107985607A (en) * | 2017-11-02 | 2018-05-04 | 成都飞机工业（集团）有限责任公司 | A kind of unmanned plane icing protective device |
| CN111655990A (en) * | 2017-11-14 | 2020-09-11 | 联合发动机制造集团股份公司 | Method for controlling an anti-icing system for an aircraft gas turbine engine air intake |
| CN109522627A (en) * | 2018-11-01 | 2019-03-26 | 西安电子科技大学 | Fan blade icing prediction technique based on SCADA data |
| CN110702419A (en) * | 2019-10-11 | 2020-01-17 | 中国直升机设计研究所 | Anti-icing conformance test system and method for engine air inlet system |
Non-Patent Citations (2)
| Title |
|---|
| 关于飞机结冰的多重安全边界问题;桂业伟等;《航空学报》;20161110(第02期);全文 * |
| 姿控发动机真空深冷环境中传热研究;肖泽娟等;《热科学与技术》;20050601(第02期);全文 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN114166460A (en) | 2022-03-11 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN114166460B (en) | Aircraft air inlet passage test device and system and hot gas anti-icing test stability judgment method | |
| CA2867782C (en) | Total air temperature sensors | |
| EP2317293B1 (en) | Total air temperature probe and method for reducing de-icing/anti-icing heater error | |
| US10006813B2 (en) | Temperature sensors | |
| US9156557B2 (en) | Icing sensor system and method | |
| CN107560822B (en) | Anti-icing wind tunnel test equipment for front edge of engine air inlet and test method thereof | |
| CN103101626B (en) | Icing detector | |
| EP2969762B1 (en) | Engine inlet total air temperature sensor | |
| US9476690B2 (en) | Method for controlling the clearance at the tips of blades of a turbine rotor | |
| CN103048109B (en) | Wing type test piece for icing wind tunnel test of aircraft wing anti-icing system | |
| WO2013034083A1 (en) | Freezing condition detector | |
| CN205203396U (en) | Aircraft engine nacelle steam anti -icing system | |
| US10850857B2 (en) | Ice detection | |
| US20110226903A1 (en) | De-icing device comprising means for detection of a leakage in a hot air supply system | |
| CN110702419B (en) | Anti-icing conformance test system and method for engine air inlet system | |
| CN116605414A (en) | Aeroengine fairing icing simulation measurement method and deicing method | |
| CN115452312B (en) | Hot gas flow measurement and control method and system | |
| CN216246492U (en) | Temperature-control type ice-preventing and removing atmospheric data probe | |
| CN105509917B (en) | A kind of aircraft tail fairing thermometric experimental rig | |
| CN111553084B (en) | Helicopter icing detection position optimization method, electronic product and storage device | |
| BR112020009553A2 (en) | method for controlling the antifreeze system of the air intake of the aircraft gas turbine engine | |
| CN110589026B (en) | Ground test device of closed-loop aircraft electric anti-icing system | |
| CN219694541U (en) | Total temperature and total pressure measuring device | |
| CN115599144A (en) | Dynamic temperature feedback electric heating anti-icing method and device for air inlet passage | |
| CN115932500A (en) | Temperature rise curve-based composite insulator point temperature rise defect diagnosis method and device |
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 |