CN115901163A - Wind tunnel test method for aerodynamic characteristics of landing of helicopter - Google Patents
Wind tunnel test method for aerodynamic characteristics of landing of helicopter Download PDFInfo
- Publication number
- CN115901163A CN115901163A CN202310221939.XA CN202310221939A CN115901163A CN 115901163 A CN115901163 A CN 115901163A CN 202310221939 A CN202310221939 A CN 202310221939A CN 115901163 A CN115901163 A CN 115901163A
- Authority
- CN
- China
- Prior art keywords
- helicopter
- ship
- acquisition module
- wind tunnel
- acquisition
- 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.)
- Granted
Links
Images
Classifications
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention discloses a wind tunnel test method for the aerodynamic characteristics of helicopter landing, which adopts a six-degree-of-freedom parallel mechanism to realize ship attitude simulation, adopts a master-slave mode synchronization method to realize synchronous acquisition of helicopter model load and ship attitude, realizes the combined motion simulation of single degree of freedom or multiple degrees of freedom of a ship through computer control, improves the accuracy of ship attitude simulation, realizes the synchronous acquisition of ship attitude and helicopter load through the master-slave mode, can acquire rich data, and has important significance for analyzing and researching the influence relationship between the ship attitude and the helicopter load.
Description
Technical Field
The invention relates to the field of wind tunnel tests, in particular to a wind tunnel test method for the landing aerodynamic characteristics of a helicopter.
Background
When a ship sails on the sea, the meteorological conditions are complex, the wind direction is changeable, the vortexes of the rotor and the ship are mixed with each other along with the flying of the helicopter above a deck, a complex coupling flow field is formed, especially, the ship can generate spatial six-degree-of-freedom motion under the high sea condition, the motion can undoubtedly interfere with the flow field condition in the taking-off and landing process of the ship-borne helicopter, and adverse effects can be brought to the pneumatic characteristics of the rotor and the operation of a driver. The research on the aerodynamic characteristics of the helicopter landing process by the ship motion under the high sea condition has important significance for the safe landing of the helicopter. At present, many domestic scholars study the flow field structure of a ship deck by means of CFD numerical simulation, but the domestic published experimental study on the ship coupling wind tunnel is less, and particularly the experimental study on the aerodynamic characteristics of a helicopter under ship coupling under the condition of ship body dynamic motion is less.
The helicopter landing aerodynamic characteristic experiment method needs to simulate six-degree-of-freedom motion of a ship under a high sea condition, and simultaneously needs to obtain helicopter load and ship attitude data in order to analyze the correlation between the helicopter landing aerodynamic characteristic and the ship attitude. Therefore, a test method capable of simulating ship motion and synchronous data acquisition is needed, so that research on the aerodynamic characteristics of the helicopter in the ship landing process is realized, and the influence rule between the load of the helicopter model and the ship posture is revealed.
Disclosure of Invention
The invention aims to realize a wind tunnel test method for helicopter landing aerodynamic characteristics, which adopts a six-degree-of-freedom parallel mechanism to realize ship attitude simulation and adopts a master-slave mode synchronization method to realize synchronous acquisition of helicopter model loads and ship attitudes.
In order to achieve the above object, the scheme of the invention is as follows:
a wind tunnel test method for aerodynamic characteristics of landing of a helicopter comprises the following steps:
s1: a six-degree-of-freedom motion mechanism is installed on a floor of a wind tunnel test section, a ship model is installed at the upper part of the six-degree-of-freedom motion mechanism, and heaving, swaying, surging, pitching, head shaking and rolling six-degree-of-freedom motion of a ship is realized through the six-degree-of-freedom motion mechanism;
s2: a groove is opened on the aviation floor parallel to the floor of the wind tunnel test section, the ship model can pass through the through groove,
lifting the aviation floor to the ship model waterline position by lifting, wherein the aviation floor is used for simulating the sea surface;
s3: connecting a helicopter model to an aviation floor through a support rod, wherein the helicopter model is connected with the support rod through a six-component balance;
s4: the top of the wind tunnel test section is provided with a ship model attitude acquisition module, and the ship model attitude acquisition module acquires ship model attitude information in real time by adopting an optical trajectory tracking system;
s5: the helicopter load acquisition module acquires the load of a helicopter model in real time according to the analog signal on the six-component balance through the data line;
s6: and synchronously acquiring attitude information of the optical track tracking system through the control system according to acquisition feedback of the analog signal.
In the technical scheme, a gap is reserved between the ship model and the aviation floor in the S2, and the ship model and the aviation floor do not interfere with each other when moving mutually.
In the above technical solution, in S6, the specific method for synchronous acquisition includes the following steps:
a1: starting a helicopter load acquisition module and acquiring an initial reading;
a2: monitoring the load and the rotating speed in real time by utilizing a helicopter load acquisition module;
a3: controlling the rotation speed of a helicopter model to meet the test requirement, and controlling the total pitch angle of a rotor wing to enable the load to meet the test requirement;
a4: controlling a six-degree-of-freedom motion mechanism to move according to given parameters of amplitude and oscillation frequency;
a5: starting a wind tunnel power system, and controlling the wind speed to reach the test required wind speed;
a6: the ship model attitude acquisition module prepares to acquire data and waits for synchronous acquisition of signals;
a7: the helicopter load acquisition module starts to acquire data according to the set sampling frequency and the number of sampling points, and simultaneously sends synchronous acquisition signals to the ship model attitude acquisition module;
a8: the ship model attitude acquisition module receives the synchronous acquisition signal and starts to acquire data at the same time;
a9: after the helicopter load acquisition module finishes acquisition according to the set number of sampling points, sending an acquisition stop signal to the ship model attitude acquisition module, and after receiving the acquisition stop signal, stopping acquisition and storing data by the ship model attitude acquisition module;
a10: changing the amplitude and frequency of the six-degree-of-freedom motion mechanism, and repeating the steps A6-A9 until the test is finished;
a11: and after the test is finished, stopping the six-freedom-degree motion mechanism and stopping the wind tunnel power system.
In the above technical solution, in A3, the rotation speed and the collective pitch angle of the helicopter model are wirelessly controlled by a remote controller.
In the above technical solution, the six-degree-of-freedom motion includes a single degree of freedom of heave, sway, surge, pitch, rock head, and roll, or a combination motion of any several degrees of freedom.
In the technical scheme, the acquisition frequency of the ship model attitude acquisition module is determined according to more than 20 times of the oscillation frequency of the six-degree-of-freedom movement mechanism, the sampling frequency of the helicopter load acquisition module is determined according to 64 times of the ship model attitude acquisition frequency, and 64 points are acquired by each rotation of a helicopter rotor wing.
In the technical scheme, the synchronous acquisition mode adopts a master-slave synchronous mode, the helicopter load acquisition module is the master, and the ship model attitude acquisition module is the slave.
In the above technical solution, the synchronization signal is a TTL level signal output by the data acquisition module, the high level is 5V, and the low level is 0V.
In summary, due to the adoption of the technical scheme, the invention has the beneficial effects that: the wind tunnel test method for the aerodynamic characteristics of landing of the helicopter adopts the parallel six-degree-of-freedom mechanism to simulate the movement of a ship on the sea surface, realizes the combined movement simulation of single degree of freedom or multiple degrees of freedom of the ship through computer control, and improves the accuracy of ship attitude simulation. Through a master-slave synchronization mode, synchronous acquisition of the ship attitude and the helicopter load is realized, abundant data can be acquired, and the method has important significance for analyzing and researching the influence relationship between the ship attitude and the helicopter load.
Drawings
The invention will now be described, by way of example, with reference to the accompanying drawings, in which:
FIG. 1 is a schematic layout of the test apparatus;
FIG. 2 is a block diagram of a signal synchronization acquisition process;
FIG. 3 is a timing diagram of signal synchronous acquisition;
FIG. 4 is a graph of typical condition experimental data.
Detailed Description
All of the features disclosed in this specification, or all of the steps in any method or process so disclosed, may be combined in any combination, except combinations of features and/or steps that are mutually exclusive.
Any feature disclosed in this specification (including any accompanying claims, abstract and drawings), may be replaced by alternative features serving equivalent or similar purposes, unless expressly stated otherwise. That is, unless expressly stated otherwise, each feature is only an example of a generic series of equivalent or similar features.
As shown in fig. 1, the test device mainly comprises a ship six-degree-of-freedom attitude simulation mechanism, a ship attitude acquisition module and a helicopter load acquisition module, wherein the ship six-degree-of-freedom attitude simulation mechanism is installed on the ground of a wind tunnel test section, an aviation floor is adopted to simulate a sea surface, a through groove is formed in the aviation floor and used for lifting the aviation floor to a waterline of the ship model in a test through the ship model, and the aviation floor is used for simulating the sea surface. The ship model is arranged on the six-degree-of-freedom attitude simulation mechanism, and the six-degree-of-freedom mechanism can realize the combined motion of single degree of freedom or multiple degrees of freedom and can simulate the ship motion under different sea conditions.
The helicopter model is connected to the aviation floor through a support rod, the helicopter model is connected with the support rod through a six-component balance, when in a blowing test state, the load of the helicopter model can be obtained through the component balance, the six-component balance converts the load into an analog signal, the analog signal enters a data acquisition module, and the data acquisition module converts the analog signal into a digital signal through A/D conversion and stores the digital signal. In order to meet the test requirements, the ship attitude and the helicopter load must be synchronously acquired, and because the attitude of the ship six-degree-of-freedom motion simulation device is obtained by resolving the displacement of six servo driving rods, the attitude data acquired by the six-degree-of-freedom device control system has low frequency and cannot be synchronously acquired with the helicopter load data. Therefore, the space displacement of the ship body mark point is obtained through a plurality of infrared cameras arranged at the top of the wind tunnel, and then the six-degree-of-freedom attitude of the ship is obtained through conversion matrix calculation.
As shown in fig. 2, the synchronous acquisition steps of the ship attitude and the helicopter load include parameter setting, load monitoring, synchronous signal output and the like. The realization process is as follows:
firstly, setting the sampling frequency and the number of sampling points of a helicopter load acquisition module according to test requirements, starting to continuously monitor the load and the rotating speed, and waiting for the stability of the load and the rotating speed of a helicopter; starting the ship attitude simulation mechanism and the wind tunnel, and after reaching a test state, enabling the ship attitude acquisition module to enter a ready acquisition state; the helicopter load acquisition module acquires data according to a set sampling frequency and a point number and outputs a TTL level signal through a digital I/O interface; after detecting the TTL level signal, the ship attitude acquisition module starts to acquire according to a set sampling frequency; and stopping collecting after the collection is finished, and entering a next-wheel collection state. The above process was repeated until the end of the test.
In this embodiment, the signal acquisition needs to be strictly acquired according to the acquisition timing sequence, as shown in fig. 3,
when the rising edge of the pulse starts to be collected, the attitude of the ship and the load of the helicopter start to be collected at the same time, and the starting point of the recorded data is ensured to be the same. The ship attitude acquisition frequency is determined according to more than 20 times of the ship oscillation frequency so as to ensure enough attitude resolution, the highest ship oscillation frequency in the embodiment is 3Hz, the ship attitude sampling frequency is set to be 60Hz, and the ship attitude acquisition clock has 60 rising edges per second. Helicopter load is a dynamic signal, the load of a rotor changes periodically every turn, and in order to obtain stable and accurate load, 64 points of each turn need to be averaged into one point. Therefore, one datum is acquired by the ship attitude, 64 points need to be acquired by the helicopter load, namely the sampling frequency of the helicopter load is 64 times of the acquisition frequency of the ship attitude, and the helicopter load acquisition clock has 60 multiplied by 64=3840 rising edges per second.
In this embodiment, the attitude of the ship and the lift data of the helicopter in a typical state are obtained through the above experimental method, and as shown in fig. 4, the data curve of the course of the helicopter model stress varying with the attitude is obtained when the ship moves in a longitudinal and transverse direction (pitch and roll combined motion) at 3HZ, the pitch amplitude is 2 °, and the roll amplitude is 5 °. When the ship moves, the distance between a paddle disc at the hovering position of the helicopter and a ship deck is periodically changed, the lift force of the helicopter is periodically changed under the same frequency due to the ground effect, and a phase lag exists, so that the aerodynamic law of the helicopter is met.
In the embodiment, the ship attitude motion is controlled by a computer, and different motion states can be realized by changing parameters so as to meet different test requirements and improve the test efficiency.
The invention is not limited to the foregoing embodiments. The invention extends to any novel feature or any novel combination of features disclosed in this specification and any novel method or process steps or any novel combination of features disclosed.
Claims (7)
1. A wind tunnel test method for aerodynamic characteristics of landing of a helicopter is characterized by comprising the following steps:
s1: a six-degree-of-freedom motion mechanism is installed on a floor of a wind tunnel test section, a ship model is installed at the upper part of the six-degree-of-freedom motion mechanism, and heaving, swaying, surging, pitching, head shaking and rolling six-degree-of-freedom motion of a ship is realized through the six-degree-of-freedom motion mechanism;
s2: opening a groove on an aviation floor parallel to a floor of a wind tunnel test section, wherein the ship model can pass through the through groove, and the aviation floor is lifted to a ship model waterline position by lifting and is used for simulating the sea surface;
s3: connecting a helicopter model to an aviation floor through a support rod, wherein the helicopter model is connected with the support rod through a six-component balance;
s4: the top of the wind tunnel test section is provided with a ship model attitude acquisition module, and the ship model attitude acquisition module acquires ship model attitude information in real time by adopting an optical trajectory tracking system;
s5: the helicopter load acquisition module acquires the load of a helicopter model in real time according to the analog signal on the six-component balance through the data line;
s6: and synchronously acquiring the attitude information of the optical track tracking system through the control system according to the acquisition feedback of the analog signal.
2. The wind tunnel test method for the aerodynamic characteristics of landing on a ship of a helicopter according to claim 1, characterized in that: and a gap is reserved between the ship model and the aviation floor in the S2, and the ship model and the aviation floor are not interfered with each other when the ship model and the aviation floor move mutually.
3. The wind tunnel test method for the aerodynamic characteristics of landing on a ship of a helicopter according to claim 1, characterized in that in S6, the specific method for synchronous acquisition comprises the following steps:
a1: starting a helicopter load acquisition module and acquiring an initial reading;
a2: monitoring the load and the rotating speed in real time by utilizing a helicopter load acquisition module;
a3: controlling the rotation speed of a helicopter model to meet the test requirement, and controlling the total pitch angle of a rotor wing to enable the load to meet the test requirement;
a4: controlling a six-degree-of-freedom motion mechanism to move according to given parameters of amplitude and oscillation frequency;
a5: starting a wind tunnel power system, and controlling the wind speed to reach the test required wind speed;
a6: the ship model attitude acquisition module prepares to acquire data and waits for synchronous acquisition of signals;
a7: the helicopter load acquisition module starts to acquire data according to the set sampling frequency and the number of sampling points, and simultaneously sends synchronous acquisition signals to the ship model attitude acquisition module;
a8: the ship model attitude acquisition module receives the synchronous acquisition signal and starts to acquire data at the same time;
a9: after the helicopter load acquisition module finishes acquisition according to the set number of sampling points, sending an acquisition stop signal to the ship model attitude acquisition module, and after receiving the acquisition stop signal, stopping acquisition and storing data by the ship model attitude acquisition module;
a10: changing the amplitude and frequency of the six-degree-of-freedom motion mechanism, and repeating the steps A6-A9 until the test is finished;
a11: and after the test is finished, stopping the six-freedom-degree motion mechanism and stopping the wind tunnel power system.
4. The wind tunnel test method for the aerodynamic characteristics of landing on a helicopter according to claim 3, wherein in A3, the rotating speed and the collective pitch angle of the helicopter model are wirelessly controlled by a remote controller.
5. A wind tunnel test method for aerodynamic characteristics of landing on a helicopter according to claim 1 or 3, characterized in that the six-degree-of-freedom motion comprises a single degree of freedom of heave, sway, surge, pitch, rock head, roll or a combination of any several degrees of freedom.
6. The wind tunnel test method for the aerodynamic characteristics of landing on a helicopter as claimed in claim 3, wherein the acquisition frequency of the ship model attitude acquisition module is determined according to more than 20 times of the oscillation frequency of the six-degree-of-freedom movement mechanism, the sampling frequency of the helicopter load acquisition module is determined according to 64 times of the ship model attitude acquisition frequency, and 64 points are acquired per rotation of a helicopter rotor.
7. The wind tunnel test method for the aerodynamic characteristics of landing on a ship of a helicopter according to claim 1 or 3, characterized in that the synchronous acquisition mode adopts a master-slave synchronous mode, the helicopter load acquisition module is the master, and the ship model attitude acquisition module is the slave.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310221939.XA CN115901163B (en) | 2023-03-09 | 2023-03-09 | Helicopter landing aerodynamic characteristic wind tunnel test method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310221939.XA CN115901163B (en) | 2023-03-09 | 2023-03-09 | Helicopter landing aerodynamic characteristic wind tunnel test method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN115901163A true CN115901163A (en) | 2023-04-04 |
| CN115901163B CN115901163B (en) | 2023-06-16 |
Family
ID=86473080
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310221939.XA Active CN115901163B (en) | 2023-03-09 | 2023-03-09 | Helicopter landing aerodynamic characteristic wind tunnel test method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115901163B (en) |
Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH06144390A (en) * | 1992-11-16 | 1994-05-24 | Kawasaki Heavy Ind Ltd | Flight test method for automatic landing experimental plane |
| US20140350754A1 (en) * | 2013-05-23 | 2014-11-27 | Honeywell International Inc. | Aircraft precision approach and shipboard landing control system and method |
| CN106125765A (en) * | 2016-08-03 | 2016-11-16 | 中国人民解放军总参谋部第六十研究所 | A kind of boat-carrying depopulated helicopter vehicle-mounted landing analog systems |
| CN109033628A (en) * | 2018-07-25 | 2018-12-18 | 南京航空航天大学 | A kind of helicopter dynamic warship wind limit figure production method and system |
| CN110749412A (en) * | 2019-10-09 | 2020-02-04 | 中国空气动力研究与发展中心低速空气动力研究所 | Ship swaying table for wind tunnel test |
| CN110836760A (en) * | 2019-11-06 | 2020-02-25 | 南京航空航天大学 | Ship attitude dynamic simulation system for wind tunnel test and working method thereof |
| CN111881632A (en) * | 2020-07-30 | 2020-11-03 | 南京航空航天大学 | Helicopter wind limit diagram determining method and system |
| CN112173158A (en) * | 2020-09-25 | 2021-01-05 | 中国直升机设计研究所 | Landing/ship-borne load calculation method for wheeled landing gear helicopter |
| CN114169068A (en) * | 2021-11-23 | 2022-03-11 | 中国直升机设计研究所 | Carrier landing flight characteristic analysis method suitable for coaxial rigid rotor helicopter |
-
2023
- 2023-03-09 CN CN202310221939.XA patent/CN115901163B/en active Active
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH06144390A (en) * | 1992-11-16 | 1994-05-24 | Kawasaki Heavy Ind Ltd | Flight test method for automatic landing experimental plane |
| US20140350754A1 (en) * | 2013-05-23 | 2014-11-27 | Honeywell International Inc. | Aircraft precision approach and shipboard landing control system and method |
| CN106125765A (en) * | 2016-08-03 | 2016-11-16 | 中国人民解放军总参谋部第六十研究所 | A kind of boat-carrying depopulated helicopter vehicle-mounted landing analog systems |
| CN109033628A (en) * | 2018-07-25 | 2018-12-18 | 南京航空航天大学 | A kind of helicopter dynamic warship wind limit figure production method and system |
| CN110749412A (en) * | 2019-10-09 | 2020-02-04 | 中国空气动力研究与发展中心低速空气动力研究所 | Ship swaying table for wind tunnel test |
| CN110836760A (en) * | 2019-11-06 | 2020-02-25 | 南京航空航天大学 | Ship attitude dynamic simulation system for wind tunnel test and working method thereof |
| CN111881632A (en) * | 2020-07-30 | 2020-11-03 | 南京航空航天大学 | Helicopter wind limit diagram determining method and system |
| CN112173158A (en) * | 2020-09-25 | 2021-01-05 | 中国直升机设计研究所 | Landing/ship-borne load calculation method for wheeled landing gear helicopter |
| CN114169068A (en) * | 2021-11-23 | 2022-03-11 | 中国直升机设计研究所 | Carrier landing flight characteristic analysis method suitable for coaxial rigid rotor helicopter |
Non-Patent Citations (3)
| Title |
|---|
| 曹普孙;段广战;陈平剑;: "舰船流场对舰载直升机起降特性影响研究", 航空计算技术, no. 03, pages 72 - 74 * |
| 赵鹏程;李海旭;宗昆;王金玲;: "直升机舰上起降安全的影响因素研究", 船舶工程, no. 09, pages 92 - 96 * |
| 陈华健;徐国华;史勇杰;: "直升机动态着舰流场下的旋翼气动载荷分析", 南京航空航天大学学报, no. 02, pages 66 - 73 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN115901163B (en) | 2023-06-16 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN102305699A (en) | Wind tunnel experiment system for free flight model | |
| CN102180270B (en) | Microminiature rotorcraft experiment platform and application thereof | |
| CN105547676A (en) | Multifunctional swing-arm type rotor wing test stand | |
| CN110849576B (en) | Wind tunnel testing device of deformable wing with adjustable attack angle | |
| CN109443689B (en) | Wind tunnel test measuring device and measuring method for dynamic aerodynamic force of radar antenna during rotating work | |
| CN106017847A (en) | Observation system and method for aerodynamic force test and flapping wing flow field of flapping-wing micro air vehicle | |
| CN111623951B (en) | Wind tunnel model multi-attitude coupling real-time force measuring device and test method | |
| CN104199462B (en) | Sea wave sensing based water unmanned plane autonomous taking-off and landing control system | |
| CN103010485A (en) | Simulation modeling method for tilt-rotor unmanned plane and system thereof | |
| CN103837322A (en) | Hydrodynamic performance testing method of water surface aircraft | |
| CN109883642B (en) | Vehicle-mounted force measuring system of low-speed aircraft | |
| CN107525647A (en) | A kind of dynamical bifurcation generating means of aerodynamic stalling | |
| CN108216695A (en) | A kind of pendulum model Mars unmanned plane rotor system hovering characteristic test device | |
| CN109823566A (en) | A kind of vertically taking off and landing flyer flight control system test platform | |
| CN110749412A (en) | Ship swaying table for wind tunnel test | |
| CN115901163B (en) | Helicopter landing aerodynamic characteristic wind tunnel test method | |
| CN205719468U (en) | A kind of platform realizing the gentle dynamic test of miniature ornithopter flapping wing Flow visualisation | |
| CN111216921A (en) | Test system and test method for ground-imitating flight of unmanned aerial vehicle | |
| CN109635376B (en) | Modeling method and system based on unmanned aerial vehicle | |
| CN203870468U (en) | Mechanical failure debugging device used for rotorcraft | |
| CN202219839U (en) | Omnidirectional flight wing structure of vertically taking-off and landing aircraft | |
| CN114879533A (en) | eVTOL aircraft control surface load simulation test method | |
| Wang et al. | A feasibility verification scheme with flight analysis for balloon-borne launched unmanned aerial vehicles | |
| CN115290282B (en) | Wind tunnel test method for researching course control efficiency of control surface of flying wing aircraft | |
| CN114063629A (en) | Portable two-degree-of-freedom flight attitude control experiment table |
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 |