CN114408211A - Airplane flap actuator disengagement test device and method - Google Patents

Abstract

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

Description

Airplane flap actuator disengagement test device and method

Technical Field

The invention relates to the field of mechanical system structure design, in particular to a device and a method for a release test of an aircraft flap actuator.

Background

The single side of a large-scale airplane flap lift-rising system generally consists of two to three wing surfaces, each wing surface is respectively and synchronously driven by a coaxial mechanical system driven in a centralized way, and a single flap wing surface is simultaneously operated by two actuators. Flap actuators are typically designed for fail-safe, i.e., when one of the actuators fails to disengage, the remaining intact actuator and/or cross-linking mechanism(s) must be able to bear the load across the entire airfoil to maintain the entire airfoil tip angle to maintain the lift balance of the left and right airfoils. The fault scene has the problems of large deformation, rigid motion, plastic nonlinearity and the like, and the conventional finite element analysis is difficult to ensure enough accuracy and needs to be tested and verified.

To verify the above system design, a load strength test and a functional test under a disconnection fault are required, and there are 3 key problems in the test process:

1) how to realize the disconnection of an actuator under a large load;

2) how to realize the actual flap installation supporting conditions such as the bending of the wing box section and the like;

3) after the brake is switched off, the problem of wing surface loading during the process from large displacement/large deformation to buffering, energy absorption and braking occurs in a very short time (30-50 ms).

The existing technical scheme of the airplane related disconnection is designed for structural members with disconnection requirements, such as a guard plate, a control surface, a generator and the like. There has been no solution specifically adapted for actively controlled disengagement under load and simulation of wing box deformation under high loads for a while.

If the mechanical structure disconnection scheme is directly applied to the airplane-related disconnection scheme, there are cases where the disconnection state is not instantaneous disconnection and the structural design requirement is high, which is not suitable for the actual airplane narrow component installation space design.

The following problems exist with the commonly used trial break-off or disengagement schemes: 1) the real part actuator has high cost and long purchase period, and the real part actuator is basically infeasible to be disconnected in a mode of prefabricating defects on the real part; 2) if only the release fault function test without loading and fixed clamping position is carried out, the simulation can be carried out by removing the actuator or a free-rotating actuator dummy piece, and the simulated fault situation is limited.

The conventional breaking/clutch in the industry has overlarge volume and low working load, and is not suitable for the situation of narrow space and large working load (the maximum torque can be 1-2 ten thousand Newton meters) for installing the flap actuator of the airplane.

Thus, for extreme conditions such as flap disengagement, the prior art has not been able to achieve a good solution.

In addition, because the aircraft wing box often has great deformation under the disconnection fault, in the test process, if fixed installation is adopted, the deformation of the aircraft wing box cannot be simulated, and if a method for installing a wing box support and loading the wing box support to generate deformation is adopted, the loading workload is large, the cost is high, and the extreme displacement under the large load under the fault condition cannot be accurately simulated.

In response to the above-identified deficiencies of the prior art, it would be desirable to provide an improved aircraft flap actuator disengagement test apparatus and method.

Disclosure of Invention

The following presents a simplified summary of one or more aspects in order to provide a basic understanding of such 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 a disengagement test device for an aircraft flap actuator, which comprises: a flap; a flap actuator; a release mechanism comprising a sleeve, a drive link, and an explosive bolt, the sleeve being connected to an output of the flap actuator, a first end of the drive link being connected to the sleeve by the explosive bolt and a second end of the drive link being connected to the flap, wherein the sleeve and the drive link move integrally and entrain the flap when the flap actuator drives the sleeve; and a release control unit that transmits a trigger signal to the explosive bolt to break the explosive bolt, thereby releasing the sleeve and the driving link.

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

In some embodiments, the sleeve comprises a slot on the inside and the slot is connected to an output gear of the flap actuator, which rotates the sleeve via the output gear.

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

In some embodiments, splines are arranged on the outer side of the sleeve, and the explosion bolt is tightly connected with a bolt hole on the driving connecting rod through the splines.

In some embodiments, the first end of the drive link includes one or more leadthroughs through which the leadthroughs of the explosive bolt are connected to a disconnect control unit, the disconnect control unit transmitting a trigger signal via the leadthroughs such that the explosive bolt breaks in response to the trigger signal.

In some embodiments, the aircraft flap actuator disengagement testing apparatus further comprises one or more sensors for collecting status data of the flap.

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

In some embodiments, the device further comprises 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.

In some embodiments, the apparatus further comprises a follow-up loading platform comprising a force controlled loading ram, wherein the force controlled loading ram is connected to the flap to load the flap with the simulated load.

In some embodiments, the slave loading platform further comprises a position control actuator for controlling the attitude of the table and a table supporting the force control loading ram.

The invention also provides a method for using the airplane flap actuator disengagement test device to carry out the airplane flap actuator disengagement test, which comprises the following steps: driving the sleeve by a flap actuator to move the sleeve and the drive link integrally and drive the flap; transmitting a trigger signal to the explosive bolt through a disengagement control unit to break the explosive bolt, so that the sleeve and the driving connecting rod are disengaged; and acquiring state data of the flap through sensors arranged at various positions of the aircraft flap actuator disengagement test device.

In some embodiments, the method further comprises: it is determined whether the flap is in the allowable operating state range based on the state data collected by the sensor.

The aircraft flap actuator release test device can transfer larger torque load before triggering release fault, so that simulation is closer to real condition. Meanwhile, by using the explosive bolt, quick and controllable separation can be realized, so that the real instantaneous separation fault condition can be simulated. In addition, the airplane flap actuator disengagement testing device does not occupy redundant space when being installed, the explosive bolts are small, and the influence of the generated explosion on surrounding parts can be basically ignored.

Drawings

The features, nature, and advantages of the present invention will become more apparent from the detailed description set forth below when taken in conjunction with the drawings. In the drawings, like reference numerals are used to designate corresponding parts throughout the several views. It is noted that the drawings described are only schematic and are non-limiting. In the drawings, the size of some of the elements may be exaggerated and not drawn on scale for illustrative purposes.

Fig. 1 shows an overall schematic view of an aircraft flap actuator disengagement test device according to the invention.

Fig. 2 shows a functional schematic of an aircraft flap connection.

Fig. 3 shows a schematic view of the structure of the drive link in fig. 2.

Fig. 4 shows a schematic structural diagram of a release mechanism of the aircraft flap actuator release test device according to the invention.

Fig. 5 shows a schematic structural diagram of a displacement simulation unit of the aircraft flap actuator disengagement test device according to the invention.

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

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

Detailed Description

In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention will be described in further detail with reference to the accompanying drawings in conjunction with specific embodiments. In the following detailed description, numerous specific details are set forth in order to provide a thorough understanding of the described exemplary embodiments. It will be apparent, however, 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 merely illustrative of the invention and are not intended to limit the invention. Meanwhile, the various aspects described in the embodiments may be arbitrarily combined without conflict.

The existing airplane-related disengagement schemes are designed for structural members with disengagement requirements, such as a fender, a steering surface, a generator and the like. There has been no solution specifically adapted for actively controlled disengagement under load and that can simulate deformation of the wing box under high loads for a while. The mechanical structure disengagement scheme is directly carried over, so that the disengagement state is not instantaneous disconnection, the structural design requirement is high, and the mechanical structure disengagement scheme is not suitable for the installation space design of the actual narrow parts of the airplane.

Therefore, the invention provides an improved disengagement test device for an aircraft flap actuator, which realizes instantaneous disengagement of the flap actuator under a large load and enables simulation to be closer to a real situation. Meanwhile, the airplane flap actuator disengagement testing device is small in structure, easy to install and low in testing cost.

Fig. 1 shows an overall schematic view of an aircraft flap actuator disengagement testing apparatus 100 of the present invention.

As shown, the device 100 comprises a displacement simulation unit 1, a tripping mechanism 2, a simulation flap 3 and a follow-up loading platform 4. The device 100 further comprises a flap actuator (not shown in the figures), which may be mounted in the displacement simulation unit 1.

The displacement simulation unit 1 and the simulation flap 3 are connected by a release mechanism 2. The load sensor, the displacement sensor and the angle sensor are respectively arranged on each test component.

The displacement simulation unit 1 can be used to simulate the deformation/displacement of the wing flap actuator at the wing box installation position under a large load. The follow-up loading platform 4 can acquire flap transient variation data.

At the moment of flap release, due to the fact that the structure and the stress relation are instantaneously changed, release load can greatly affect the position of the flap, and the follow-up loading platform 4 can follow the flap 3 to generate large instantaneous displacement. Sensors (not shown) installed at various locations collect relevant signals for verifying that the design has achieved the target.

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

Fig. 2 shows a functional schematic of a conventional aircraft flap connection.

Fig. 2 shows a conventional connection scheme of body parts such as a flap actuator, a drive link and a flap. As shown, the drive link is connected at one end to the flap actuator via a mounting flange and at the other end to other components (not shown) such as a flap.

In normal operation, the flap actuator applies work through rotation to enable the driving connecting rod to swing, so that the flap is driven to extend or retract.

Fig. 3 shows a schematic view of the structure of the drive link in fig. 2.

As shown in fig. 3, one end of the drive link has a catch slot that is connected to the output gear end of the flap actuator. The other end of the drive link is connected to the flap by a joint tab.

It follows that in conventional aircraft flap connection solutions, the flap and the flap actuator are connected by a single part (drive link). The flap actuator transmits drive to the flap through the single piece to extend or retract the flap.

When designing the structure of the flap actuating system, the working condition that the extreme condition causes the breakage of the driving connecting rod or the disengagement of the actuator needs to be considered. Therefore, it is desirable to devise a test scheme to actively disengage the drive link or actuator. To accomplish this, the present invention redesigns the conventional drive link to achieve active disengagement of the drive link or actuator.

Fig. 4 shows a schematic structural diagram of a release mechanism 400 of the aircraft flap actuator release test device according to the invention.

The present invention redesigns the drive linkage configuration of fig. 3 in order to achieve active disengagement. The purpose of this structural design lies in: the driving connecting rod structure can work normally before a disconnection instruction is not received, the connecting structure drives the system to move normally, and after the disconnection instruction is received, the driving connecting rod structure can be disconnected into two parts and cannot be driven mechanically.

As shown in fig. 4, the disconnect mechanism 400 comprises a sleeve 30, a drive link 10 and an explosion bolt 20, wherein a first end of the drive link 10 is connected to the sleeve 30 by the explosion bolt 20 and a second end of the drive link 10 is connected to the flap (not shown in the figures) by a joint tab 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. Upon a trigger release failure, the explosive bolt 20 breaks to disengage the sleeve 30 and the drive link 10, and the flap actuator drives the sleeve 30 without driving the drive link 10 and the flap.

As shown, the inner side of the sleeve 30 comprises a catch 31, and the catch 31 is connected with the output gear of the flap actuator. When the release failure is not triggered, the flap actuator rotates the sleeve 30 through the output gear, and then integrally moves the drive link 10 to drive the flap.

The outer wall of the sleeve 30 is fitted with the inner wall of the drive link 10, and the fitted portion is a cylindrical surface. For example, the first end of the drive link 10 may include an outer barrel surrounding the sleeve 30 with a gap between the outer wall of the sleeve 30 and the inner wall of the outer barrel.

In an embodiment of the present invention, explosive bolt 20 may comprise a plurality of explosive bolts evenly distributed around the outside of sleeve 30. For example, 6 explosive bolts are shown in fig. 4 and evenly distributed around the outside of the sleeve 30.

It should be noted that while a specific number of explosive bolts are shown in fig. 4, this is merely exemplary and not limiting. In a practical implementation, more or less than 6 explosive bolts may be employed, and the explosive bolts may be distributed in different ways.

The spline 32 is arranged on the outer side of the sleeve 30, and the explosion bolt 20 is tightly connected with a bolt hole (not shown) on the driving connecting rod 10 through the spline 32.

The disconnect mechanism 400 also includes one or more feedthrough holes through which the feedthrough of the explosive bolt can be connected to a control device (e.g., a disconnect control unit).

For example, fig. 4 shows two lead holes 11 and 12, wherein the lead hole 11 is located at a 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 invention, the lead of the explosive bolt 20 is led into the driving connecting rod 10 through the lead hole 11 and then led out of the driving connecting rod 10 through the lead hole 12, and is connected to the disconnection control unit, so as to realize active control of the triggering of the explosive bolt.

It should be noted that while fig. 4 shows two specific wire holes 11 and 12, this is merely exemplary and not limiting. In different implementations, the disengagement mechanism may include a different number of wire holes, and the wire holes may be arranged in a manner different than that of fig. 4.

In use, when a disengagement fault is not triggered, the sleeve 30 and drive link 10 can be considered as a unit, driven by the flap actuator through the internal gear of the sleeve. The flap actuator is fixed on the wing box, and then, in combination with the real aircraft connection relationship of fig. 2, the whole release mechanism is connected with the output gear end of the flap actuator through the sleeve slot 31. Under normal working conditions, the flap actuator drives the connecting rod to control the flap through the clamping groove 31 connected with the gear.

When the trigger release failure occurs, the explosive bolt 20 breaks and the spline 32 disconnects 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 clearance, the rotation of the sleeve by the gear cannot be transmitted to the drive link, thereby achieving effective decoupling and completing fault simulation.

When the fault simulation is not triggered, the combination of the driving connecting rod and the sleeve in fig. 4 can replace the function of the original driving connecting rod, and the mechanical properties are consistent. Before the simulated disconnection fault is disconnected, a large torque load can be transmitted, so that the simulation is closer to the real condition.

In addition, the release mechanism of fig. 4 is small in structure, easy to install, and does not affect surrounding parts. Because the installation does not occupy redundant space, and the explosive bolt is less, the influence of the generated explosion on surrounding parts can be basically ignored. In addition, by using explosive bolts, a fast and controlled release can be achieved to simulate real instantaneous release failure conditions.

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

The displacement simulation unit 500 of the invention fixes a sliding table on the test bench through a clamp, and can simulate the displacement of the supporting position of the wing box section in the driving rocker arm plane through the angle of the clamp and the worm and gear mechanism on the sliding table.

As shown in fig. 5, the displacement simulation unit 500 includes a jig 40, a test bed 50, and a slide table 60. The slide table 60 is fixed to the stage 50 by a jig 40.

In the embodiment of the present invention, the slide table 60 includes a fixed mount 61, a slide rail 62, a mounting mount 63, a worm wheel 64, a driving device 65, a worm 66, and a slider 67.

The slide 67 can slide on the slide 62, the slide 62 being fixed to the fixed support 61.

The flap actuator is fixed to a mounting bracket 63, wherein the mounting bracket 63 is fixed to a slide 67 by means of fasteners. Meanwhile, a slider 67 is connected to one end of the worm 66.

During the test, the worm wheel 64 can be driven to rotate by the driving device 65, so that the worm 66 is driven to move up and down, and the sliding piece 67 (and the flap actuator) can slide along the sliding rail 62 to obtain the required position of the mounting support 63.

Through the displacement simulation unit 500, the displacement of the actuator flap mounting interface wing box under a large load can be accurately simulated.

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

The aircraft flap actuator disengagement test apparatus of the present invention further includes a high load, fast response, compliant loading platform 600. As shown, the servo loading platform 600 is divided into two layers, and a plurality of force-controlled loading actuators 70 are arranged on the upper layer for simulating the pneumatic external load. The force controlled load ram 70 is mounted on a middle deck 80. The lower layer is provided with a plurality of position control actuators 90 for controlling the attitude of the intermediate table 80 so that the intermediate table 80 and flap wing surface motions are kept as consistent as possible to reduce the displacement variation of the first layer loading ram 70. The base 100 is used to support and fix the entire platform.

In practical implementation, the follow-up loading platform 600 can be controlled by an external command. Specifically, the force controlled load ram 70 may be commanded to apply a predetermined load to the wing to simulate a real load. Meanwhile, the position control actuator 90 can be controlled by a command to control the posture of the middle table 80.

Since the entire moment of disengagement is completed within 50ms and the flap moves rapidly in space at the moment of disengagement, the demand on actuator response speed is high. By the above-described follow-up loading platform 600, a rapid response under a large load can be simulated.

For a better understanding of the invention, a method for performing a release test using the aircraft flap actuator release test apparatus of the invention is explained below with reference to fig. 7.

Fig. 7 illustrates an example flow chart of an aircraft flap actuator disengagement testing method 700 of the invention. In a preferred embodiment, the method 700 may be performed by the aircraft flap actuator disengagement testing apparatus 100 of FIG. 1.

Method 700 begins at step 705. In step 705, the flap actuator drive sleeve of the test device 100 is disengaged by the aircraft flap actuator to move the sleeve and drive link together and bring the flap.

In step 710, a trigger signal is transmitted to the explosive bolt by the disconnect control unit to break the explosive bolt, thereby disconnecting the sleeve and the drive link.

When the sleeve and drive link are disengaged, the disengagement mechanism is split into two parts, thereby disabling mechanical actuation. At this time, the flap actuator drives the sleeve without driving the drive link and the flap.

In an embodiment of the invention, the disconnection fault may be triggered by a disconnection control unit. In particular, the lead of the explosive bolt can be connected to the disconnection control unit through a lead hole to enable the triggering of a disconnection fault. In a specific implementation, the disconnection control unit can be controlled manually to trigger the disconnection fault, and can also be controlled by a computer to trigger the disconnection fault.

At step 715, status data for the flap is collected via sensors mounted at various locations of the device.

In embodiments of the invention, the sensor may comprise one or more of: load sensor, displacement sensor, angle sensor.

For example, the status data of the flap may include load data, displacement data, angle data, etc. of the flap at the time of the triggering disengagement fault.

At step 720, it is determined whether the flap is in the allowable operating state range based on the status data collected by the sensors.

In embodiments of the present invention, whether a design meets a target may be verified based on the collected data. In particular, a target state range of the flap after a disengagement fault can be set, within which the flap can still continue to operate. If this range is exceeded, this indicates that the flap may not continue to operate.

It should be noted that the order of the above-described steps of method 700 is exemplary and not limiting. In various embodiments of the present invention, the above steps may be performed in a different order or in parallel, or new steps may be added, depending on the actual situation.

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

1. when the fault simulation is not triggered, the combination of the driving connecting rod and the sleeve can replace the functions of the conventional driving connecting rod in the prior art, and the mechanical properties of the driving connecting rod and the sleeve are consistent.

2. Before the simulated disconnection fault is disconnected, the testing device can transmit larger torque load, so that the simulation is closer to the real condition.

3. The test device has the advantages of small whole part connecting structure, easy installation and no influence on surrounding parts.

4. Because the installation does not occupy redundant space, and the explosive bolt is less, the influence of the generated explosion on surrounding parts can be basically ignored.

5. Compared with the disconnection caused by the prefabrication defect, the part of the invention can be repeatedly used except the explosive bolt, is convenient for daily use and repeated tests, and has low cost.

6. By using explosive bolts, a fast and controllable trip can be achieved for simulating real instantaneous trip fault conditions.

7. The displacement of the wing box of the interface for mounting the actuator flap under a large load can be simulated by the wing box displacement simulation device.

8. Through the follow-up loading platform, the expected movement of the flap after being disengaged can be effectively realized, and the wing surface loading is maintained, so that the real stress state is simulated.

The detailed description set forth above in connection with the appended drawings describes examples and is not intended to represent all examples that may be implemented or 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 "superior 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 present invention. Thus, usage of such phrases may not refer to only 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 generic principles defined herein may be applied to other aspects. Thus, the claims are not intended to be limited to the aspects shown herein, but is to be accorded the full scope consistent with the language claims, wherein reference to an element in the singular is not intended to mean "one and only one" unless specifically so stated, but rather "one or more. The term "some" means one or more unless specifically stated otherwise. All structural and functional equivalents to the elements of the various aspects described throughout this disclosure that are known or later come to be known to those of ordinary skill in the art are expressly incorporated herein by reference and are intended to be encompassed by the claims.

It is also noted that the embodiments may be described as a process which is depicted as a flowchart, a flow diagram, a structure diagram, or a block diagram. Although a flowchart may describe the operations as a sequential process, many of the operations can be performed in parallel or concurrently. In addition, the order of the operations may be rearranged.

While various embodiments have been illustrated and described, it is to be understood that the embodiments are not limited to the precise configuration 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 devices disclosed herein without departing from the scope of the claims.

Claims (13)

1. An aircraft flap actuator release test device comprising:

a flap;

a flap actuator;

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

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

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 on an inner side thereof 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, wherein a gap exists between an outer wall of the sleeve and an inner wall of the outer barrel.

5. The device of claim 1, wherein splines are arranged on the outside of the sleeve, and the explosion bolt is tightly connected with a bolt hole on the driving connecting rod through the splines.

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

7. The device of claim 1, wherein the aircraft flap actuator disengagement testing device further comprises one or more sensors for collecting status data of the flap.

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

9. 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.

10. The apparatus of claim 1, further comprising a follow-up loading platform comprising a force controlled loading ram, wherein the force controlled loading ram is connected to the flap to load the flap with a simulated load.

11. The device of claim 10, wherein the compliant loading platform further comprises a position control actuator and a table supporting the force control loading ram, the position control actuator for controlling the attitude of the table.

12. A method of performing an aircraft flap actuator release test using the apparatus of any one of claims 1 to 11, comprising:

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

transmitting a trigger signal to the explosive bolt through the disengagement control unit to break the explosive bolt, thereby disengaging the sleeve and the drive link; and

the status data of the flap are acquired by sensors installed at various positions of the device.

13. The method of claim 12, further comprising:

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

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