CN114408211B - Device and method for testing detachment of airplane flap actuator - Google Patents

Abstract

The invention discloses an aircraft flap actuator release test device, which comprises: a flap; a flap actuator; a release mechanism comprising a sleeve, a drive link and an explosion bolt, the sleeve being connected to an output end of the flap actuator, a first end of the drive link being connected to the sleeve by the explosion bolt and a second end of the drive link being connected to the flap, wherein the sleeve and the drive link move integrally and drive the flap when the flap actuator drives the sleeve; and a disengagement control unit that transmits a trigger signal to the explosion bolt to break the explosion bolt, thereby disengaging the sleeve and the driving link.

Description

Aircraft flap actuator disengagement test device and method

Technical field

The invention relates to the field of mechanical system structural design, and in particular to an aircraft flap actuator disengagement test device and method.

Background technique

The flap lift system of large aircraft usually consists of two or three airfoils on one side. Each airfoil is synchronously driven by a centrally driven coaxial mechanical system. A single flap airfoil is controlled by two actuators simultaneously. Flap actuators are generally designed according to damage safety, that is, when one of the actuators fails to disengage, the remaining intact actuators and/or cross-linking mechanisms must be able to withstand the load on the entire wing surface and maintain the entire Wing deflection angle to maintain lift balance on the left and right wings. This fault scenario has problems such as large deformation, rigid body motion, and plastic nonlinearity. Conventional finite element analysis alone is difficult to ensure sufficient accuracy and requires experimental verification.

In order to verify the above system design, load strength test and functional test under disengagement failure need to be carried out, and there are three key issues during the test process:

1) How to realize actuator disconnection under large load;

2) How to realize real flap installation support conditions such as wing box segment bending;

3) After disconnection, large displacement/deformation occurs in a very short time (30-50ms) to the airfoil loading problem during the buffering and energy-absorbing braking process.

The current aircraft-related detachment technical solutions are designed for structural parts that require detachment, such as fenders, control surfaces, generators, etc. There is currently no technical solution that is specifically suitable for active control disengagement under load and that can simulate the deformation of the wing box segment under large loads.

If the mechanical structure disengagement scheme is directly applied to the aircraft-related disengagement scheme, there will be situations where the disengagement state is not instantaneous disconnection, and the structural design requirements are high, which is not suitable for the actual design of the installation space of small aircraft parts.

Commonly used test disconnection or disengagement solutions have the following problems: 1) The cost of real actuators is high and the procurement cycle is long, and it is basically unfeasible to disconnect by prefabricating defects on the real parts; 2) If it is only carried out without loading, Moreover, the functional test of the disengagement failure of the fixed clamping position can be simulated by dismantling the actuator or a freely rotating actuator dummy, and the simulated fault scenarios are limited.

The conventional disconnect/clutch in the industry is too large and has a low working load. It is not suitable for the small space and large working load of aircraft flap actuator installation (the maximum torque can reach 10,000-20,000 Nm).

Therefore, for extreme working conditions such as flap disengagement, the existing technology does not have a good solution to achieve it.

In addition, since the aircraft wing box section often undergoes large deformation under a disengagement failure, during the test process, if fixed installation is used, the deformation of the wing box section cannot be simulated. If the wing box section support is installed and loaded, the deformation of the wing box section cannot be simulated. The method of generating deformation requires a large loading workload and high cost, and the extreme displacement under large loads under fault conditions cannot be accurately simulated.

In view of the above-mentioned shortcomings of the prior art, it is desired to provide an improved aircraft flap actuator disengagement test device and method.

Contents of the invention

A brief overview of one or more aspects is given below to provide a basic understanding of these aspects. This summary is not an extensive overview of all contemplated aspects and is intended to neither identify key or critical elements of all aspects nor delineate the scope of any or all aspects. Its sole purpose is to present some concepts of one or more aspects in a simplified form as a prelude to the more detailed description that is presented later.

The invention provides an aircraft flap actuator disengagement test device, which includes: a flap; a flap actuator; and a disengagement mechanism. The disengagement mechanism includes a sleeve, a driving connecting rod and an explosive bolt. The sleeve Connected to the output end of the flap actuator, a first end of the drive link is connected to the sleeve through the explosive bolt and a second end of the drive link is connected to the flap, wherein the flap is actuated When the sleeve is driven by the driver, the sleeve and the driving link move integrally and drive the flap; and a disengagement control unit transmits a trigger signal to the explosive bolt to break the explosive bolt, thereby causing The sleeve and the drive link are disengaged.

In some embodiments, the explosive bolts include a plurality of explosive bolts evenly distributed around the outside of the sleeve.

In some embodiments, the inner side of the sleeve includes a slot and the slot is connected to the output gear of the flap actuator. The flap actuator drives the sleeve to rotate through the output gear.

In some embodiments, the first end of the drive link includes an outer sleeve surrounding the sleeve with a gap between the outer wall of the sleeve and the inner wall of the outer sleeve.

In some embodiments, splines are arranged on the outside of the sleeve, and the explosive bolt is tightly connected to the bolt hole on the driving connecting rod through the splines.

In some embodiments, the first end of the drive link includes one or more lead holes through which leads of the explosive bolt are connected to the disengagement control unit, and the disengagement control unit transmits the trigger signal via the leads, Causes the explosive bolt to break in response to a trigger signal.

In some embodiments, the aircraft flap actuator disengagement test device further includes one or more sensors for collecting flap status data.

In some embodiments, the one or more sensors include one or more of the following: a load sensor, a displacement sensor, and an angle sensor.

In some embodiments, the device further includes a displacement simulation unit, which includes a slide rail and a mounting bracket that can slide on the slide rail, wherein the flap actuator is fixed on the mounting bracket.

In some embodiments, the device further includes a follow-up loading platform that includes a force-controlled loading actuator, wherein the force-controlled loading actuator is connected to the flap to load the simulated load onto the flap.

In some embodiments, the follow-up loading platform further includes a position-controlled actuator and a table supporting a force-controlled loading actuator. The position-controlled actuator is used to control the posture of the table.

The invention also provides a method for carrying out the aircraft flap actuator disengagement test using the aforementioned aircraft flap actuator disengagement test device, which includes: driving the sleeve through the flap actuator so that the sleeve and The drive link moves integrally and drives the flap; the trigger signal is transmitted to the explosive bolt by the disengagement control unit to break the explosive bolt, thereby disengaging the sleeve and the drive link; and the flap actuator is disengaged by being installed on the aircraft. Sensors at various positions of the test device are turned on to collect status data of the flaps.

In some embodiments, the method further includes: determining whether the flap is in an allowed operating state range based on status data collected by the sensor.

The aircraft flap actuator disengagement test device of the present invention can transmit a large torque load before triggering a disengagement failure, thereby making the simulation closer to the real situation. At the same time, through the use of explosive bolts, rapid and controllable release can be achieved to simulate a real instantaneous release failure situation. In addition, the aircraft flap actuator disengagement test device of the present invention does not occupy extra space for installation, and the explosion bolt is small, so the impact of the explosion on surrounding components can be basically ignored.

Description of drawings

The characteristics, nature and advantages of the present invention will become more apparent when the detailed description set forth below is understood in conjunction with the accompanying drawings. In the drawings, identical reference characters are identified accordingly throughout. It is to be noted that the figures described are schematic and non-limiting only. In the drawings, the dimensions of some components may be exaggerated and not drawn to scale for illustrative purposes.

Figure 1 shows an overall schematic diagram of the aircraft flap actuator disengagement test device of the present invention.

Figure 2 shows a functional schematic diagram of an aircraft flap connection.

FIG. 3 shows a schematic structural diagram of the driving link in FIG. 2 .

Figure 4 shows a schematic structural diagram of the disengagement mechanism of the aircraft flap actuator disengagement test device of the present invention.

Figure 5 shows a schematic structural diagram of the displacement simulation unit of the aircraft flap actuator disengagement test device of the present invention.

Figure 6 shows a schematic structural diagram of the follow-up loading platform of the aircraft flap actuator disengagement test device of the present invention.

Figure 7 shows an example flow chart of the aircraft flap actuator disengagement test method of the present invention.

Detailed ways

In order to make the purpose, technical solutions and advantages of the present invention more clear, the present invention will be further described in detail below in conjunction with specific embodiments and with reference to the accompanying drawings. In the following detailed description, numerous specific details are set forth to provide a thorough understanding of the described exemplary embodiments. However, it will be apparent to one skilled in the art that the described embodiments may be practiced without some or all of these specific details. In other exemplary embodiments, well-known structures have not been described in detail to avoid unnecessarily obscuring the concepts of the present disclosure. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention. At the same time, various aspects described in the embodiments can be combined arbitrarily without conflict.

Existing aircraft-related detachment solutions are designed for structural parts that require detachment, such as fenders, control surfaces, generators, etc. There is currently no solution that is specifically suitable for active control disengagement under load and can simulate the deformation of the wing box segment under large loads. However, if the mechanical structure detachment scheme is directly copied, there will be situations where the detachment state is not instantaneous disconnection, and the structural design requirements are high, which is not suitable for the actual design of the narrow parts installation space of the aircraft.

To this end, the present invention provides an improved aircraft flap actuator disengagement test device, which realizes the instantaneous disengagement of the flap actuator under large load, making the simulation closer to the real situation. At the same time, the aircraft flap actuator disengagement test device of the present invention has a small structure, is easy to install, and has low test cost.

Figure 1 shows an overall schematic diagram of the aircraft flap actuator disengagement test device 100 of the present invention.

As shown in the figure, the device 100 includes a displacement simulation unit 1, a disengagement mechanism 2, a simulated flap 3 and a follow-up loading platform 4. The device 100 also includes a flap actuator (not shown in the figure), which can be installed in the displacement simulation unit 1 .

The displacement simulation unit 1 and the simulation flap 3 are connected through the disengagement mechanism 2 . Load sensors, displacement sensors, and angle sensors are installed on each test component respectively.

The displacement simulation unit 1 can be used to simulate the deformation/displacement of the flap actuator at the installation position of the wing box section under large loads. The follow-up loading platform 4 can collect instantaneous flap change data.

At the moment when the flap is disengaged, due to instantaneous changes in the structure and force relationship, the disengagement load will have a greater impact on the flap position, and the follow-up loading platform 4 will follow the flap 3 to undergo a large instantaneous displacement. Sensors (not shown) installed everywhere will collect relevant signals to verify whether the design achieves its goals.

The various components of the aircraft flap actuator disengagement test device 100 will be explained in further detail below.

Figure 2 shows a functional schematic diagram of a conventional aircraft flap connection.

Figure 2 shows the conventional connection scheme of airframe components such as flap actuators, drive links and flaps. As shown in the figure, one end of the driving link is connected to the flap actuator through the mounting flange, and the other end is connected to other components such as flaps (not shown in the figure).

In normal operation, the flap actuator performs work through rotation, causing the drive link to swing, thereby driving the flaps to extend or retract.

FIG. 3 shows a schematic structural diagram of the driving link in FIG. 2 .

As shown in Figure 3, one end of the driving link has a slot, which is connected to the output gear end of the flap actuator. The other end of the drive link connects to the flap via a connector lug.

It can be seen that in the conventional aircraft flap connection scheme, the flap and flap actuator are connected through a single part (driving link). The flap actuator transmits drive to the flaps through this single part, thereby extending or retracting the flaps.

When designing the structure of the flap actuator system, it is necessary to consider the working conditions in which extreme conditions may cause the drive link to break or the actuator to become disconnected. Therefore, a test scheme needs to be designed to actively disengage the drive linkage or actuator. In order to realize the above test plan, the present invention redesigns the conventional driving link to realize active disengagement of the driving link or actuator.

Figure 4 shows a schematic structural diagram of the disengagement mechanism 400 of the aircraft flap actuator disengagement test device of the present invention.

In order to achieve active disengagement, the present invention redesigns the driving link structure in Figure 3. The purpose of this structural design is: before receiving the disengagement command, the drive linkage structure can work normally and drive the system to move normally through the connecting structure. After receiving the disengagement command, the drive linkage structure will disengage into two parts. , cannot be mechanically driven.

As shown in FIG. 4 , the disengagement mechanism 400 includes a sleeve 30 , a driving connecting rod 10 and an explosive bolt 20 . The first end of the driving connecting rod 10 is connected to the sleeve 30 through the explosive bolt 20 , and the third end of the driving connecting rod 10 is connected to the sleeve 30 through the explosive bolt 20 . The two ends are connected to the flaps (not shown in the figure) through connector lugs 13 .

When the disengagement fault is not triggered, the sleeve 30 and the drive link 10 move integrally and the flap actuator drives the flap through the sleeve 30 and the drive link 10 . When a disengagement fault is triggered, the explosive bolt 20 breaks so that the sleeve 30 and the driving link 10 are disconnected, and the flap actuator drives the sleeve 30 without driving the driving link 10 and the flap.

As shown in the figure, the inner side of the sleeve 30 includes a clamping groove 31, and the clamping slot 31 is connected with the output gear of the flap actuator. When the disengagement fault is not triggered, the flap actuator drives the sleeve 30 to rotate through the output gear, thereby causing the drive link 10 to move integrally to drive the flap.

The outer wall of the sleeve 30 matches the inner wall of the driving connecting rod 10, and the matching part is a cylindrical surface. For example, the first end of the driving link 10 may include an outer cylinder surrounding the sleeve 30, and there is a certain gap between the outer wall of the sleeve 30 and the inner wall of the outer cylinder.

In embodiments of the present invention, the explosive bolts 20 may include a plurality of explosive bolts evenly distributed around the outside of the sleeve 30 . For example, six explosive bolts are shown in FIG. 4 , and these explosive bolts are evenly distributed around the outside of the sleeve 30 .

It should be noted that although a specific number of explosive bolts is shown in Figure 4, this is illustrative only and not limiting. In actual implementation, more or less than 6 explosive bolts can be used, and the explosive bolts can be distributed in different ways.

Splines 32 are arranged on the outside of the sleeve 30 , and the explosive bolt 20 is tightly connected to the bolt hole (not shown in the figure) on the driving connecting rod 10 through the splines 32 .

The release mechanism 400 also includes one or more lead holes through which leads of the explosive bolt can be connected to a control device (eg, a release control unit).

For example, FIG. 4 shows two lead holes 11 and 12 , where the lead hole 11 is located at the first end of the driving link 10 and the lead hole 12 is located on the rotation axis of the driving link 10 and the sleeve 30 .

In the embodiment of the present invention, the lead wire of the explosive bolt 20 penetrates into the driving connecting rod 10 through the lead hole 11, and then is led out of the driving connecting rod 10 through the lead hole 12, and is connected to the disengagement control unit to achieve active control of the triggering of the explosive bolt. .

It should be noted that although FIG. 4 shows two specific lead holes 11 and 12, this is only illustrative and not limiting. In different implementations, the decoupling mechanism may include a different number of lead holes, and the lead holes may be arranged differently than in FIG. 4 .

In use, when the disengagement fault is not triggered, the sleeve 30 and the drive link 10 can be regarded as a whole, and are driven by the flap actuator through the internal gear of the sleeve. The flap actuator is fixed on the wing box section. Combined with the real aircraft connection relationship in Figure 2, the entire disengagement mechanism will be connected to the output gear end of the flap actuator through the sleeve slot 31. Under normal operating conditions, the flap actuator drives the connecting rod to control the flaps through the gear-connected slot 31 .

When a disengagement fault is triggered, the explosion bolt 20 breaks and the spline 32 is disconnected from the drive link 10 . Since the outer contact surface of the sleeve 30 and the inner contact surface of the drive link 10 are both cylindrical and have a certain gap, the rotation of the sleeve driven by the gear cannot be transmitted to the drive link, thereby achieving effective disengagement and completing fault simulation. .

When the fault simulation is not triggered, the combination of the driving link and sleeve in Figure 4 can replace the function of the original driving link and has the same mechanical properties. Before the simulated disconnection fault is disconnected, a larger torque load can be transmitted, making the simulation closer to the real situation.

In addition, the disengagement mechanism in Figure 4 has a small structure, is easy to install, and will not affect surrounding components. Since the installation does not take up extra space and the explosive bolts are small, the impact of the explosion on surrounding components can be basically ignored. In addition, by using explosive bolts, rapid and controlled release can be achieved to simulate a real instantaneous release failure situation.

Figure 5 shows a schematic structural diagram of the displacement simulation unit 500 of the aircraft flap actuator disengagement test device of the present invention.

The displacement simulation unit 500 of the present invention fixes a slide table on the test bench through a fixture. Through the angle of the fixture and the worm gear mechanism on the slide table, it can simulate the displacement of the wing box section support position in the driving rocker arm plane.

As shown in FIG. 5 , the displacement simulation unit 500 includes a fixture 40 , a test bench 50 and a slide table 60 . The sliding table 60 is fixed on the stand 50 through the clamp 40 .

In the embodiment of the present invention, the sliding table 60 includes a fixed support 61 , a slide rail 62 , a mounting support 63 , a worm gear 64 , a driving device 65 , a worm 66 and a sliding member 67 .

The sliding member 67 can slide on the slide rail 62 , and the slide rail 62 is fixed on the fixed support 61 .

The flap actuator is fixed on the mounting bracket 63, wherein the mounting bracket 63 is fixed on the slider 67 by fasteners. At the same time, the sliding member 67 is connected to one end of the worm 66 .

During the test, the driving device 65 can be used to drive the worm gear 64 to rotate, thereby driving the worm 66 to move up and down, so that the sliding member 67 (and the flap actuator) slides with the slide rail 62 to obtain the required installation support 63. Location.

Through the above-mentioned displacement simulation unit 500, the displacement of the wing box section of the actuator flap installation interface under large loads can be accurately simulated.

Figure 6 shows a schematic structural diagram of the follow-up loading platform 600 of the aircraft flap actuator disengagement test device of the present invention.

The aircraft flap actuator disengagement test device of the present invention also includes a follow-up loading platform 600 that responds quickly to large loads. As shown in the figure, the follow-up loading platform 600 is divided into two layers, and a number of force-controlled loading actuators 70 are arranged on the upper layer for simulating aerodynamic external loads. The force-controlled loading actuator 70 is installed on the middle table 80 . There are a number of position-controlled actuators 90 arranged on the lower floor, which are used to control the attitude of the middle table 80 so that the motion of the middle table 80 and the flap airfoil are as consistent as possible, so as to reduce the displacement change of the first layer loading actuator 70. The base 100 is used to support and fix the entire platform.

In actual implementation, the follow-up loading platform 600 can be controlled through external instructions. Specifically, the force-controlled loading actuator 70 can be controlled by instructions to apply a predetermined load to the wing, thereby simulating a real load. At the same time, the position control actuator 90 can be controlled through instructions to control the attitude of the intermediate table 80 .

Since the entire disengagement moment is completed within 50ms, and the flap moves rapidly in space at the disengagement moment, the response speed of the actuator is required to be high. Through the above-mentioned dynamic loading platform 600, rapid response under large loads can be simulated.

In order to better understand the present invention, the method of using the aircraft flap actuator disengagement test device of the present invention to conduct a disengagement test will be explained below with reference to FIG. 7 .

FIG. 7 shows an example flow diagram of the aircraft flap actuator disengagement test method 700 of the present invention. In a preferred embodiment, method 700 may be performed by the aircraft flap actuator disengagement test apparatus 100 of FIG. 1 .

Method 700 begins at step 705. In step 705, the flap actuator driving sleeve of the test device 100 is disengaged through the aircraft flap actuator, so that the sleeve and the driving link move integrally and drive the flap.

At step 710, a trigger signal is transmitted to the explosive bolt through the disengagement control unit to cause the explosive bolt to break, thereby disengaging the sleeve and the driving link.

When the sleeve and drive link become disengaged, the disengagement mechanism separates into two parts, rendering mechanical actuation impossible. At this time, when the flap actuator drives the sleeve, it does not drive the drive link and flaps.

In embodiments of the invention, a disengagement fault may be triggered by disengagement from the control unit. Specifically, the lead wire of the explosive bolt can be connected to the disengagement control unit through the lead hole to achieve the triggering of the disengagement fault. In a specific implementation, the disconnection control unit can be manually controlled to trigger the disconnection fault, or the disconnection control unit can be controlled by a computer to trigger the disconnection fault.

In step 715, flap status data is collected through sensors installed at various locations of the device.

In embodiments of the present invention, the sensor may include one or more of the following: a load sensor, a displacement sensor, and an angle sensor.

For example, the status data of the flap may include load data, displacement data, angle data, etc. when the flap triggers a disengagement failure.

In step 720, it is determined whether the flap is in the allowed operating state range based on the status data collected by the sensor.

In embodiments of the present invention, it can be verified based on the collected data whether the design achieves the goals. Specifically, the target state range of the flap after a disengagement failure can be set, within which the flap can still continue to work. If it exceeds this range, it means that the flaps may not continue to work.

It should be noted that the order of the above steps of method 700 is illustrative and not limiting. In various embodiments of the present invention, the above steps may be performed in different orders or in parallel, or new steps may be added according to actual conditions.

The aircraft flap actuator disengagement test device of the present invention has the following advantages:

1. When fault simulation is not triggered, the combination of the driving connecting rod and the sleeve of the present invention can replace the function of the conventional driving connecting rod in the prior art, and the mechanical properties of the two are consistent.

2. Before simulating a disconnection fault, the test device of the present invention can transmit a larger torque load, making the simulation closer to the real situation.

3. The overall component connection structure of the test device of the present invention is small, easy to install, and will not affect surrounding components.

4. Since the installation does not take up extra space and the explosive bolts are small, the impact of the explosion on surrounding components can be basically ignored.

5. Compared with disconnection caused by prefabricated defects, the components of the present invention can be reused except for the explosive bolts, which is convenient for daily use and repeated testing and has low cost.

6. By using explosive bolts, rapid and controllable release can be achieved to simulate real instantaneous release failure conditions.

7. Through the wing box segment displacement simulation device, the displacement of the wing box segment at the actuator flap installation interface under large loads can be simulated.

8. Through the follow-up loading platform, the expected movement after the flaps are disengaged can be effectively realized and the wing surface can be maintained to be loaded to simulate the real stress state.

The detailed description set forth above in connection with the appended drawings describes examples and does not represent all examples that may be implemented or that fall within the scope of the claims. The terms "example" and "exemplary" when used in this specification mean "serving as an example, instance, or illustration" and do not mean "better or superior to other examples."

Reference throughout this specification to "one embodiment" or "an embodiment" means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the invention. Therefore, use of these phrases may refer to more than one embodiment. Furthermore, the described features, structures or characteristics may be combined in any suitable manner in one or more embodiments.

The previous description is provided to enable any person skilled in the art to practice the various aspects described herein. Various modifications to these aspects will be readily apparent to those skilled in the art, and the general principles defined herein may be applied to other aspects. Accordingly, the claims are not intended to be limited to the aspects shown herein, but are to be accorded the full scope consistent with the language of the claims, wherein reference to the singular form of an element is not intended to be limited unless expressly stated otherwise. It means "there is and only one", but "one or more". Unless specifically stated otherwise, the term "some" refers to one or more. All structural and functional equivalents to the elements of the various aspects described throughout this invention that are now or hereafter known to those of ordinary skill in the art are expressly incorporated herein by reference and are intended to be covered by the claims.

It should also be noted that these embodiments may be described as processes depicted as flowcharts, flowcharts, structural diagrams, or block diagrams. Although a flowchart may describe operations as a sequential process, many of these operations can be performed in parallel or concurrently. Additionally, the order of these operations can be rearranged.

While various embodiments have been illustrated and described, it is to be understood that the embodiments are not limited to the precise configurations and components described above. Various modifications, substitutions and improvements apparent to those skilled in the art may be made in the arrangement, operation and details of the apparatus disclosed herein without departing from the scope of the claims.

Claims (11)

1. An aircraft flap actuator release test apparatus comprising:

a flap;

a flap actuator;

a release mechanism comprising a sleeve, a drive link and an explosion bolt, the sleeve being connected to an output end of the flap actuator, a first end of the drive link being connected to the sleeve by the explosion bolt and a second end of the drive link being connected to the flap, wherein the sleeve and the drive link move integrally and drive the flap when the flap actuator drives the sleeve;

a disengagement control unit that transmits a trigger signal to the explosion bolt to break the explosion bolt, thereby disengaging the sleeve and the drive link;

one or more sensors for acquiring status data of the flap; and

a follower loading platform comprising a force controlled loading ram, wherein the force controlled loading ram is connected to the flap to load a simulated load to the flap.

2. The apparatus of claim 1, wherein the explosive bolt comprises a plurality of explosive bolts evenly distributed around the outside of the sleeve.

3. The device of claim 1, wherein the sleeve includes a slot inside and the slot is coupled to an output gear of the flap actuator, the flap actuator rotating the sleeve via the output gear.

4. The device of claim 1, wherein the first end of the drive link comprises an outer barrel surrounding the sleeve, a gap being present between an outer wall of the sleeve and an inner wall of the outer barrel.

5. The device according to claim 1, characterized in that the sleeve is provided with splines on the outside, through which splines the explosive bolts are fastened to the bolt holes on the drive link.

6. The apparatus of claim 1, wherein the first end of the drive link includes one or more lead holes through which leads of the explosive bolt are connected to the disengagement control unit, the disengagement control unit transmitting the trigger signal via the leads such that the explosive bolt breaks in response to the trigger signal.

7. The apparatus of claim 1, wherein the one or more sensors comprise one or more of: load sensor, displacement sensor, angle sensor.

8. The device of claim 1, further comprising a displacement simulation unit comprising a slide rail and a mounting support slidable on the slide rail, wherein the flap actuator is fixed to the mounting support.

9. The apparatus of claim 1, wherein the slave load platform further comprises a position controlled actuator and a table top supporting the force controlled load ram, the position controlled actuator for controlling the attitude of the table top.

10. A method of performing an aircraft flap actuator release test using the apparatus of any one of claims 1 to 9, comprising:

driving the sleeve through the flap actuator to integrally move the sleeve and the drive link and drive the flap;

transmitting a trigger signal to the explosive bolt through the uncoupling control unit to fracture the explosive bolt, thereby uncoupling the sleeve and the drive link; and

status data of the flap is acquired by sensors mounted at various positions of the device.

11. The method as recited in claim 10, further comprising:

determining whether the flap is in an allowable operating state range based on state data collected by the sensor.