Fixed-wing aircraft flutter excitation and testing system and method

Abstract

A fixed-wing aircraft flutter excitation and testing system and method. The fixed-wing aircraft flutter excitation and testing system comprises an excitation apparatus and a testing system; the excitation apparatus comprises miniature excitation rockets arranged at excitation points of wingtips of elastic structures on an aircraft, and an ignition control system electrically connected to the miniature excitation rockets; the testing system comprises an airborne signal acquisition subsystem and a vibration signal acquisition subsystem; the airborne signal acquisition subsystem comprises a data acquisition recorder and a GPS antenna, and is used for recording flight state parameters and GPS time of the aircraft; and the vibration signal acquisition subsystem comprises low-frequency vibration acceleration sensors arranged at testing points of rigid planes of the elastic structures on the aircraft, and a data acquisition recorder. According to the fixed-wing aircraft flutter excitation and testing system and method, various flutter modes can be accurately, conveniently and efficiently obtained by means of the excitation apparatus, and flight state parameters and vibration parameters required for flutter analysis are obtained by means of the testing system.

Description

A fixed-wing aircraft flutter excitation test system and method Technical Field

[0001] The invention belongs to the technical field of flight testing, and in particular relates to a flutter excitation test system and method for a fixed-wing aircraft. Background Art

[0002] The Civil Aviation Administration of China's "Airworthiness Regulations for Normal Category Airplanes" (CCAR-23 R4 version) stipulates the aeroelastic requirements for normal category aircraft (here referring to aircraft with 19 or fewer passenger seats and a maximum certified takeoff weight of 8,618 kg or less). Among them, Section 23.2245(a) (Section 23.629 in CCAR-23 R3 version) provides relevant conditions for flutter testing, but does not provide specific design methods for the excitation method and specific construction methods for the measurement system in the flutter test.

[0003] Summary of the Invention

[0004] The purpose of the present invention is to provide a fixed-wing aircraft flutter excitation test system and method, which can complete the generation of flutter excitation and the collection of test data.

[0005] Technical solution of the present invention: In order to achieve the above-mentioned purpose, according to the first aspect of the present invention, a fixed-wing aircraft flutter excitation test system is proposed, including an excitation device and a test system; the excitation device includes an excitation small rocket arranged at an excitation point at the wingtip of an elastic structure on the aircraft, and an ignition control system electrically connected to the excitation small rocket; the test system includes an airborne signal acquisition subsystem and a vibration signal acquisition subsystem; the airborne signal acquisition subsystem includes a data acquisition recorder and a GPS antenna for recording the flight status parameters and GPS time of the aircraft; the vibration signal acquisition subsystem includes a low-frequency vibration acceleration sensor arranged at a test point on the rigid plane of the elastic structure on the aircraft, and a data acquisition recorder.

[0006] In a possible embodiment, the excitation rocket is rigidly connected to the wing tip of each elastic structure on the aircraft using bolts via a transition structure.

[0007] In one possible embodiment, the elastic structure includes a wing, a horizontal stabilizer, and a vertical stabilizer.

[0008] In a possible embodiment, each excitation point is provided with one or more small excitation rockets.

[0009] In a possible embodiment, the low-frequency vibration acceleration sensor is hard-connected to the rigid plane of each elastic structure on the aircraft using bolts through a transition structure, and the installation position is selected at one point each at the front and rear of the fuselage; and one point each at the front and rear of the elastic structure.

[0010] In one possible embodiment, the ignition control system includes a mounting plate, a double-pole switch, and a cable. Different combinations of the double-pole switches enable single-launch mode, salvo-only mode, and a combination of single and dual-launch modes. This improves the synchronization of ignitions during a salvo, reducing or even eliminating the time difference between the two rockets ignited during a salvo. The double-pole switch is secured to the mounting plate, which is then installed along with the test system on a dedicated fixed equipment rack designed and installed within the cabin. This allows for highly integrated control switches and secure mounting of the test equipment, while also facilitating easy operation by the flight crew.

[0011] In one possible embodiment, in the single-launch mode only, the double-pole switch is respectively connected to the double leads of each excitation small rocket; directly operating and closing the double-pole switch of a single small rocket can realize the launch of the small rocket in the single-launch mode only.

[0012] In one possible embodiment, in the simultaneous launch mode only, the two excitation small rockets are first connected in series and then connected to the main double-pole switch; directly operating the main double-pole switch that closes the two excitation small rockets can realize the launch of the small rockets in the simultaneous launch mode only.

[0013] In one possible embodiment, in the coexistence mode of single launch and dual launch, separate double-pole switches are respectively set on the series circuits of the two excitation small rockets and then connected to the main double-pole switch; first operate to close the main double-pole switch of the two excitation small rockets, and then operate to close the separate double-pole switch of the single excitation small rocket, which can realize the single launch of the excitation small rocket; first operate to close the separate double-pole switch of the single excitation small rocket, and then operate to close the main double-pole switch of the two excitation small rockets, which can realize the simultaneous launch of the small rockets.

[0014] According to a second aspect of the present invention, a flutter excitation test method for a fixed-wing aircraft is provided, comprising the following steps:

[0015] At the beginning of the test, after starting the aircraft engine, turn on the power supply of the excitation device and the test system;

[0016] During the test, after the aircraft speed reaches the predetermined speed point, the ignition control system is operated according to the predetermined small rocket ignition method to obtain the target flutter mode;

[0017] After the experiment, the data from the data acquisition recorder was collected, processed and analyzed.

[0018] Compared with the prior art, the present invention has the following beneficial effects:

[0019] The aircraft flight state parameters and vibration parameters obtained through the present invention can achieve GPS time synchronization, facilitating the determination of the start and end times of the time period required for data analysis, as well as the determination of the flight state parameters corresponding to the vibration parameters. The flutter modes obtained can meet the modal requirements of theoretical analysis, and the excitation device can achieve the various modes required for flutter analysis. The double-pole switch design controls the simultaneous disconnection and connection of both positive and negative poles, thereby improving the reliability and safety of the system. The present invention can accurately, conveniently, and efficiently obtain various flutter modes through the excitation device, and obtain the flight state parameters and vibration parameters required for flutter analysis through the test system. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] FIG1 is a schematic diagram of the architecture of a fixed-wing aircraft flutter excitation test system according to the present invention;

[0021] FIG2 is a schematic structural diagram of an ignition control system in a fixed-wing aircraft flutter excitation test system according to the present invention;

[0022] Figure 3 is a schematic diagram of the actual installation of the excitation rocket;

[0023] Figure 4 is a schematic diagram of the actual installation of the low-frequency acceleration sensor. DETAILED DESCRIPTION

[0024] In order to further understand the present invention, preferred embodiments of the present invention are described below in conjunction with examples. However, it should be understood that these descriptions are only for further illustrating the features and advantages of the present invention rather than limiting the patent claims of the present invention.

[0025] 1. Implementation method of the incentive device:

[0026] Figure 1 shows the structure of the stimulus device and the test system architecture.

[0027] Figure 2 shows the structure of the ignition control system.

[0028] The excitation rocket in the excitation device is an off-the-shelf product, encased in a metal casing filled with a combustible material. Two ignition fuses are reserved for the exterior. When the two fuses are connected, an electric spark is generated to ignite the excitation rocket. The rocket has a designed thrust of 100 kg (multiplied by the acceleration due to gravity of 9.8 meters per second squared, which converts to 0.98 kilonewtons), a designed force duration of 20 or 60 milliseconds, and a cylindrical shape. Because explosives are used in the production process, this product must be custom-manufactured by a qualified military enterprise.

[0029] The excitation rocket is hard-connected to the wingtip of each elastic structure (wing, horizontal stabilizer, vertical stabilizer, etc.) on the aircraft using bolts through a transfer structure. Depending on the requirements of unilateral flutter mode and contralateral flutter mode, one or more excitation rockets can be installed at the wingtip of each elastic structure and ignited on demand to achieve the excitation of multiple flutter modes in the same flight.

[0030] The ignition control system in the excitation device is a self-developed product, consisting of a mounting plate, several double-pole switches and cables. The excitation of the unilateral flutter mode is achieved by a single shot of a small rocket, and the excitation of the symmetrical flutter mode is achieved by a salvo of small rockets. Through different combinations of double-pole switches, the synchronization of the ignition action during the salvo of small rockets can be improved, and the time difference between the ignition of the two small rockets during the salvo can be reduced or even eliminated. The double-pole switch is fixed on the mounting plate, and the mounting plate is installed together with the test system on a special fixed equipment rack designed and installed separately in the cabin to achieve high integration of the control switch and firm installation of the test equipment, while being convenient for the flight crew to operate.

[0031] The excitation rocket and the ignition control system are physically connected through a shielded twisted-pair cable. Internal routing is preferred for cable layout and fixation. If the internal routing conditions cannot be met, a card block must be installed at a certain interval on the external routing path to fix the cable to prevent the cable from being too long and causing large swings during flight, affecting test flight safety.

[0032] The 28-volt power supply in the excitation device is drawn from the onboard power supply and is connected in series with the ignition fuse and double-pole switch of each small rocket to form a small rocket ignition control circuit. The ignition control circuits of each small rocket are connected in parallel.

[0033] Taking the ignition control system structure in Figure 2 as an example, the operation method of the ignition control system is explained in detail:

[0034] (1) Confirmation of launch mode: The number of small rockets for single launch only is 2 (XHJ1, XHJ2), the number of small rockets for salvo launch only is 2 (XHJ3 and XHJ4, XHJ5 and XHJ6), and the number of small rockets for both single launch and salvo launch is 4 (XHJ7, XHJ8, XHJ7 and XHJ8, XHJ9, XHJ10, XHJ9 and XHJ10).

[0035] (2) Single-shot mode operation: Directly closing the double-pole switch of a single rocket can realize the launch of a single rocket. For example, closing the double-pole switch SD1 can realize the single launch of the rocket XHJ1; closing the double-pole switch SD2 can realize the single launch of the rocket XHJ2.

[0036] (3) Operation method for salvo-only mode: Directly closing the main double-pole switch of the two small rockets can realize the salvo-only launch of the small rockets. For example, closing the double-pole switch SD3 can realize the salvo launch of small rockets XHJ3 and XHJ4; closing the double-pole switch SD4 can realize the salvo launch of XHJ5 and XHJ6.

[0037] (4) Operation method for single and double firing: First close the main double-pole switch of the two small rockets, then close the sub-double-pole switch of the single small rocket, to achieve a single firing of the small rocket; first close the sub-double-pole switch of the single small rocket, then close the main double-pole switch of the two small rockets, to achieve a simultaneous firing of the small rockets. For example: first close the main double-pole switch SD5, then close the sub-double-pole switch SD6 (or SD7), to achieve a single firing of the small rocket XHJ7 (or XHJ8); first close the main double-pole switch SD8, then close the sub-double-pole switch SD9 (or SD10), to achieve a single firing of the small rocket XHJ9 (or XHJ10); first close the sub-double-pole switches SD6 and SD7 (or SD9 and SD10), then close the main double-pole switch SD5 (or SD8), to achieve a simultaneous firing of the small rockets XHJ7 and XHJ8 (or XHJ9 and XHJ10).

[0038] 2. Implementation method of the airborne signal acquisition subsystem in the test system:

[0039] Figure 1 shows the structure of the stimulus device and the test system architecture.

[0040] The airborne signal acquisition subsystem consists of an airborne signal data acquisition recorder and a GPS antenna, and is powered by an external onboard power supply (28 volts).

[0041] The airborne signal data acquisition recorder is used to collect aircraft flight status parameters. The flight status parameters are collected by extracting the onboard bus signal and are directly obtained and stored by the airborne signal data acquisition recorder. The data in the airborne signal data acquisition recorder is the same as the onboard signal. It is installed on a special fixed equipment rack designed and installed separately in the cabin. The airborne signal data acquisition recorder can use the KAM-500 data acquisition recorder produced by Ireland's ACRA company.

[0042] The GPS antenna provides GPS time to the test system. A power splitter can be used to split the GPS input signal into two outputs, one for each data acquisition and recorder (an airborne signal data acquisition and recorder and an airborne vibration data acquisition and recorder). The GPS antenna is mounted via an extension cable on a specially designed bracket installed in front of the aircraft cockpit windshield to enhance GPS satellite signal search capabilities. An L1 antenna manufactured by Canadian company NovAtel can be used.

[0043] 3. Implementation method of vibration signal acquisition subsystem in test system:

[0044] Figure 1 shows the structure of the stimulus device and the test system architecture.

[0045] The vibration measurement subsystem consists of a low-frequency vibration acceleration sensor and a data acquisition recorder, and is powered by an external on-board power supply (28 volts).

[0046] Low-frequency vibration accelerometers are rigidly connected to the aircraft's elastic structures (wings, horizontal stabilizers, vertical stabilizers, etc.) using bolts via adapters. Installation locations include one point each at the front and rear of the elastic structures (wings, horizontal stabilizers, vertical stabilizers, etc.), and one point each at the front and rear of the fuselage. For engine modal supplements, typical engine locations (such as the reducer casing) are selected. The 3741B1230G accelerometer manufactured by PCB (USA) can be used.

[0047] The vibration signal data acquisition recorder is used to record and store the electrical signal converted from the vibration signal by each low-frequency vibration acceleration sensor (GPS time is obtained by the GPS antenna in the airborne signal acquisition subsystem through a one-to-two method). It is installed on a special fixed equipment rack designed and installed separately in the cabin. The vibration signal data collector can use the SQLABⅡ data acquisition recorder produced by the German Acoustics company.

[0048] The low-frequency acceleration sensor and the vibration signal data acquisition recorder are physically connected through a dedicated cable. Internal routing is preferred for cable layout and fixation. If the internal routing conditions cannot be met, a card block must be installed at a certain interval on the external routing path to fix the cable to prevent the cable from being too long and causing large swings during flight, affecting test flight safety.

[0049] Example 1

[0050] The Yun-12F aircraft is a commuter, multi-purpose, all-metal, semi-monocoque structure small transport aircraft with a maximum take-off weight of 8,400 kg. This type of aircraft adopts a twin-engine, high-wing, single vertical tail, and retractable front three-point landing gear overall layout. The left and right engine compartments are respectively equipped with a PT6A-65B turboprop engine produced by Pratt & Whitney Canada (P&WC). Each engine is connected to a HC-C-B5MP-3D / M10876ANSK five-blade metal propeller produced by Hartzell Corporation of the United States, with constant speed, feathering and reverse propeller functions.

[0051] (1) Installation location and method of the excitation rockets: During the flutter test of the Yun-12F aircraft, the left wing, right wing, left horizontal stabilizer, and right horizontal stabilizer were selected as the excitation locations. The rockets were installed at the wingtip front beams of the above structures (the direction of force applied was perpendicular to the ground). Three excitation rockets were installed at each installation location through a transfer structure to stimulate different modes and provide backup.

[0052] (2) Installation location and method of low-frequency acceleration sensors: During the flutter test flight of the Yun-12F aircraft, the left wingtip front beam, the left wingtip rear beam, the right wingtip front beam, the right wingtip rear beam, the left horizontal stabilizer wingtip front beam, the left horizontal stabilizer wingtip rear beam, the right horizontal stabilizer wingtip front beam, the right horizontal stabilizer wingtip rear beam, the vertical stabilizer wingtip front beam, the vertical stabilizer wingtip rear beam, the left engine reducer case, and the right engine reducer case were selected. Acceleration sensors were installed at each installation location through a transition structure to collect and record acceleration signals.

[0053] (3) Flutter mode classification and excitation method:

[0054] ① By firing two small rockets at the left and right wingtips, the "wing symmetrical 1 bend" mode is obtained.

[0055] ② By firing two small rockets at the left and right horizontal stabilizer wingtips, the "horizontal stabilizer symmetrical single bend" and "launch house symmetrical pitch" modes are obtained.

[0056] ③ By firing a small rocket at the left wingtip, the modes of "antisymmetric wing bend", "lateral fuselage bend" and "horizontal bending of the engine room" are obtained.

[0057] ④ By firing a small rocket from the left horizontal stabilizer wingtip, the modes of “vertical tail lateral bend 1” and “horizontal tail antisymmetric bend 1” are obtained.

[0058] Example 2

[0059] Specific operation methods of the stimulus device and test system:

[0060] Figure 3 illustrates the actual installation of the excitation rocket.

[0061] Figure 4 illustrates the practical installation of a low-frequency accelerometer.

[0062] ① Before the test begins, a test equipment rack is installed inside the cabin. The equipment rack has three layers, from top to bottom: the first layer is the ignition control system, the second layer is the vibration signal data collector and the airborne signal data collector, and the third layer is the spare equipment layer; install the excitation small rocket and low-frequency acceleration sensor at the predetermined position, and use cables to connect the small rocket to the ignition control system and the acceleration sensor to the vibration signal data collector.

[0063] ② At the beginning of the test, after starting the aircraft engine, turn on the power supply of the excitation device and the test system.

[0064] ③ During the test, after the aircraft speed reaches the predetermined speed point, the ignition control system is operated according to the predetermined small rocket ignition method to obtain the target flutter mode.

[0065] ④After the test, collect, process and analyze the data from the data acquisition recorder.

[0066] ⑤ Determine whether to conduct additional flights based on the test results.

[0067] Test results:

[0068] The results show that the Y-12F aircraft has a symmetrical 1-bend, a symmetrical 1-bend in the horizontal stabilizer, a symmetrical 1-bend in the engine room, an anti-symmetrical 1-bend in the wing, a lateral 1-bend in the fuselage, a horizontal bend in the engine room, a lateral 1-bend in the vertical tail, and an anti-symmetrical 1-bend in the horizontal tail. The test data under each mode are analyzed to obtain the amplitude-frequency curve, frequency-speed curve and damping ratio-speed curve. The test results show that the Y-12F aircraft has a symmetrical 1-bend in the dive speed V D Under the conditions of small rocket excitation, the aircraft did not experience any major vibration or buffeting. Under the conditions of small rocket excitation, the aircraft did not experience any flutter or control reversal or divergence.

Claims

1. A fixed-wing aircraft flutter excitation test system, characterized in that, It includes an excitation device and a test system; the excitation device includes an excitation small rocket arranged at an excitation point at the wing tip of an elastic structure on an aircraft, and an ignition control system electrically connected to the excitation small rocket; the test system includes an airborne signal acquisition subsystem and a vibration signal acquisition subsystem; the airborne signal acquisition subsystem includes a data acquisition recorder and a GPS antenna for recording the flight status parameters and GPS time of the aircraft; the vibration signal acquisition subsystem includes a low-frequency vibration acceleration sensor arranged at a test point on the rigid plane of the elastic structure on the aircraft, and a data acquisition recorder.

2. The flutter excitation test system for a fixed-wing aircraft according to claim 1, wherein The small excitation rocket is rigidly connected to the wing tip of each elastic structure on the aircraft by bolts through a transition structure.

3. A fixed-wing aircraft flutter excitation test system according to any one of claims 1 or 2, characterized in that, The elastic structure includes wings, horizontal stabilizers, and vertical stabilizers.

4. A fixed-wing aircraft flutter excitation test system according to claim 1, characterized in that, Each excitation point is provided with one or more small excitation rockets.

5. The flutter excitation test system for a fixed-wing aircraft according to claim 1, characterized in that, The low-frequency vibration acceleration sensor is hard-connected to the rigid plane of each elastic structure on the aircraft using bolts through a transfer structure. The installation position is selected at one point each at the front and rear of the fuselage; and one point each at the front and rear of the elastic structure.

6. The flutter excitation test system for a fixed-wing aircraft according to claim 1, characterized in that, The ignition control system includes a mounting plate, a double-pole switch and a cable; the double-pole switch is fixed on the mounting plate, and the double-pole switch is connected to the excitation small rocket through a cable. Through different combinations of the double-pole switches, a single-launch mode, a simultaneous-launch mode, and a single-launch and double-launch coexistence mode can be realized.

7. A fixed-wing aircraft flutter excitation test system according to claim 6, characterized in that, In the single-launch mode only, the double-pole switch is respectively connected to the double leads of each excitation small rocket; directly operating and closing the double-pole switch of a single small rocket can realize the launch of the small rocket in the single-launch mode only.

8. The flutter excitation test system for a fixed-wing aircraft according to claim 6, wherein In the salvo-only mode, the two excitation small rockets are first connected in series and then connected to the main double-pole switch; the main double-pole switch of the two excitation small rockets is directly operated to close to realize the launch of the small rockets in salvo-only mode.

9. The flutter excitation test system for a fixed-wing aircraft according to claim 6, characterized in that, In the single-launch and dual-launch coexistence mode, separate double-pole switches are respectively set on the series circuits of the two excitation small rockets and then connected to the main double-pole switch; first operate to close the main double-pole switch of the two excitation small rockets, and then operate to close the separate double-pole switch of the single excitation small rocket to achieve a single excitation small rocket; first operate to close the separate double-pole switch of the single excitation small rocket, and then operate to close the main double-pole switch of the two excitation small rockets to achieve a simultaneous launch of the small rockets.

10. A flutter excitation test method for a fixed-wing aircraft, which uses a flutter excitation test system for a fixed-wing aircraft according to any one of claims 1-9, characterized in that, The steps include: At the beginning of the test, after the aircraft engine is started, the power supply of the excitation device and the test system is turned on; During the test, after the aircraft speed reaches the predetermined speed point, the ignition control system is operated to obtain the target flutter mode according to the predetermined small rocket ignition method; After the experiment, the data from the data acquisition recorder was collected, processed and analyzed.

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