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

Device and method for testing detachment of airplane flap actuator

Technical Field

The invention relates to the field of mechanical system structural design, in particular to a device and a method for disengaging an airplane flap actuator.

Background

The single side of the flap lift-increasing system of the large-scale aircraft is generally composed of two to three airfoils, each airfoil is respectively and synchronously driven by a coaxial mechanical system which is driven in a centralized way, and a single flap airfoil is simultaneously operated by two actuators. The flap actuators are generally designed for fail-safe design, i.e., when one of the actuators fails to disengage, the remaining intact actuators and/or the cross-linking mechanism need to be able to withstand the load on the entire airfoil, maintaining the entire airfoil angle of deflection, to maintain the left and right wing lift forces balanced. The fault scene has the problems of large deformation, rigid motion, plastic nonlinearity and the like, and the sufficient accuracy is difficult to ensure only by conventional finite element analysis, and the test verification is needed.

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

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

2) How to realize real flap installation supporting conditions such as bending of a wing box section;

3) The airfoil surface loading problem occurs in the process of large displacement/large deformation to buffering and energy absorbing braking within a very short time (30-50 ms) after disconnection.

The related disengaging technical proposal of the prior aircraft is designed aiming at structural members with disengaging requirements, such as guard plates, control surfaces, generators and the like. There is no technical solution which is specifically adapted to actively control the disengagement under load and which can simulate deformation of a wing box under high loads.

If the mechanical structure disengaging scheme is directly applied to the aircraft related disengaging scheme, the disengaging state is not instantaneously disconnected, and the structural design requirement is high, so that the mechanical structure disengaging scheme is not suitable for the design of the narrow part installation space of the actual aircraft.

The conventional test disconnect or disconnect protocol suffers from the following problems: 1) The real part actuator has high cost and long purchasing period, and the disconnection is basically not feasible by prefabricating defects on the real part; 2) If only the test of the uncoupling fault function of the unloaded and fixed clamping position is carried out, the simulation can be carried out by dismantling the actuator or a freely rotatable actuator dummy, and the simulated fault situation is limited.

The conventional disconnection/clutch in industry has overlarge volume and lower working load, and is not suitable for the conditions of small space and large working load (the maximum torque can reach 1-2 Kidney meters) of the aircraft flap actuator.

Thus, for extreme conditions such as flap disengagement, no solution is available that is too good in the prior art.

In addition, because the aerofoil box section of the airplane often has larger deformation under the fault of disconnection, in the test process, the deformation of the aerofoil box section cannot be simulated if fixed installation is adopted, and if a method for loading and deforming the aerofoil box section support piece is adopted, the loading workload is large, the cost is high, and the extreme displacement under the large load of the fault condition cannot be accurately simulated.

In view of the above-described deficiencies of the prior art, it would be desirable to provide an improved aircraft flap actuator release 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 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.

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

In some embodiments, the inside of the sleeve comprises a clamping groove and the clamping groove is connected with an output gear of the flap actuator, and 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 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 explosive bolts are fastened and connected with bolt holes 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 the lead of the explosive bolt is connected to a trip control unit that transmits a trigger signal via the lead such that the explosive bolt breaks in response to the trigger signal.

In some embodiments, the aircraft flap actuator release test apparatus further comprises one or more sensors for acquiring flap status data.

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 sliding track and a mounting support slidable on the sliding track, wherein the flap actuator is fixed to the mounting support.

In some embodiments, the apparatus further comprises a follower 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 load platform further comprises a position controlled actuator for controlling the attitude of the tabletop and a tabletop supporting the force controlled load ram.

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

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

The aircraft flap actuator release test device can transmit larger torque load before triggering release faults, so that simulation is closer to the actual situation. Meanwhile, by using the explosion bolt, the quick and controllable disconnection can be realized, so as to simulate the real instant disconnection fault condition. In addition, the aircraft flap actuator release test device does not occupy redundant space, the explosion 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 designate corresponding parts throughout the different 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 of an aircraft flap actuator release test apparatus of the present invention.

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

Fig. 3 shows a schematic diagram of the drive link of fig. 2.

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

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

FIG. 6 shows a schematic structural view of a follower loading platform of an aircraft flap actuator release test apparatus of the present invention.

FIG. 7 illustrates an example flow chart of a method of aircraft flap actuator release test of the present invention.

Detailed Description

The objects, technical solutions and advantages of the present invention will become more apparent by the following detailed description of the present invention with reference to the accompanying drawings. 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 in order to avoid unnecessarily obscuring the concepts of the present disclosure. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the scope of the invention. Meanwhile, the various aspects described in the embodiments may be arbitrarily combined without conflict.

Existing aircraft related decoupling schemes are designed for components requiring decoupling, such as shields, control surfaces, generators, etc. There is no solution for actively controlling the decoupling under load that is specifically adapted to simulate deformation of a wing box under high loads. The mechanical structure disconnecting scheme is directly carried over, the disconnecting state is not the condition of instantaneous disconnection, the structural design requirement is high, and the mechanical structure disconnecting scheme is not suitable for the design of the installation space of the narrow parts of the practical aircraft.

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

FIG. 1 shows an overall schematic of an aircraft flap actuator release test apparatus 100 of the present invention.

As shown, the device 100 comprises a displacement simulation unit 1, a decoupling mechanism 2, a simulated flap 3 and a follower 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 decoupling mechanism 2. The load sensor, the displacement sensor and the angle sensor are respectively arranged on each test part.

The displacement simulation unit 1 can be used to simulate the deformation/displacement of a flap actuator at the wing box installation location under high loads. The follower loading platform 4 can collect flap transient variation data.

At the moment of disengaging the flap, the disengaging load has a larger influence on the position of the flap due to the instantaneous change of the structure and the stress relation, and the follow-up loading platform 4 can carry out larger instantaneous displacement along with the flap 3. Sensors (not shown) installed everywhere collect relevant signals for verifying that the design has reached the target.

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

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

FIG. 2 is a conventional connection scheme for body components such as flap actuators, drive links, and flaps. As shown, one end of the drive link is connected to the flap actuator via a mounting flange, and the other end is connected to other components (not shown) such as a flap.

In normal operation, the flap actuator applies work through rotation, so that the drive connecting rod swings, and the flap is driven to extend or retract.

Fig. 3 shows a schematic diagram of the drive link of fig. 2.

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

It follows that in conventional aircraft flap connection schemes, 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 feature, thereby extending or retracting the flap.

In the design of flap actuation system structures, it is necessary to take into account the conditions in which extreme conditions occur which lead to a breakage of the drive links or to a disengagement of the actuators. Thus, there is a need to devise a test scheme for actively disengaging a drive link or actuator. In order to achieve the above test scheme, the present invention redesigns the conventional drive links to achieve active disengagement of the drive links or actuators.

Fig. 4 shows a schematic structural view of a release mechanism 400 of an aircraft flap actuator release test apparatus according to the invention.

The present invention redesigns the drive link configuration of fig. 3 in order to achieve active disengagement. The aim of this structural design is: the driving connecting rod structure can normally work before the disengaging instruction is not received, the driving system normally moves through the connecting structure, and after the disengaging instruction is received, the driving connecting rod structure can be separated into two parts and cannot be mechanically driven.

As shown in fig. 4, the release mechanism 400 includes 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 a flap (not shown) by a tab 13.

In the event of a non-triggered release failure, the sleeve 30 and the drive link 10 move integrally and the flap actuator drives the flap via the sleeve 30 and the drive link 10. Upon triggering a release failure, the explosive bolt 20 breaks to release the sleeve 30 and the drive link 10, and the flap actuator drives the sleeve 30 without entraining the drive link 10 and the flap.

As shown, the inside of the sleeve 30 includes a catch 31, and the catch 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, so that the drive link 10 moves integrally to drive the flap.

The outer wall of the sleeve 30 is engaged with the inner wall of the drive link 10, and the engagement 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, the explosive bolt 20 may include a plurality of explosive bolts uniformly distributed around the outer side of the sleeve 30. For example, 6 explosive bolts are shown in fig. 4, and are evenly distributed around the outside of the sleeve 30.

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

The outside of the sleeve 30 is provided with splines 32, and the explosive bolt 20 is fastened to a bolt hole (not shown in the figure) on the drive link 10 through the splines 32.

The release mechanism 400 also includes one or more lead holes through which the leads of the explosive bolt may be connected to a control device (e.g., a release 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 drive link 10 and the lead hole 12 is located on the rotational axis of the drive link 10 and the sleeve 30.

In the embodiment of the invention, the lead wire of the explosion bolt 20 penetrates into the driving connecting rod 10 through the lead wire hole 11, and then the driving connecting rod 10 is led out through the lead wire hole 12 and is connected to the disconnection control unit, so that the active control of the explosion bolt triggering is realized.

It should be noted that while fig. 4 shows two particular lead holes 11 and 12, this is by way of example only and not by way of limitation. In different implementations, the disengagement mechanism may include a different number of lead holes, and the lead holes may be arranged in a different manner than in fig. 4.

In use, when the decoupling fault is not triggered, the sleeve 30 and the drive link 10 can be considered as a whole, driven by the flap actuator via the sleeve annulus. The flap actuator is fixed to the wing box section, and in combination with the true mechanical connection of fig. 2, the entire release mechanism is connected to the output gear end of the flap actuator through the sleeve clamping 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 a trip fault is triggered, the explosive bolt 20 breaks and the spline 32 is disconnected from the drive link 10. Because the external contact surface of the sleeve 30 and the internal contact surface of the driving connecting rod 10 are both cylindrical and have a certain gap, the sleeve rotation driven by the gear cannot be transmitted to the driving connecting rod, thereby realizing effective disengagement 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 disconnect fault breaks, a larger torque load may be transferred, making the simulation closer to reality.

In addition, the release mechanism of fig. 4 is small in structure, easy to install, and does not affect surrounding components. Because the installation does not occupy redundant space and the explosive bolt is smaller, the influence of the generated explosion on surrounding parts can be basically ignored. In addition, by using explosive bolts, a quick and controllable release can be achieved to simulate a real transient release fault condition.

Fig. 5 shows a schematic structural view of a displacement simulation unit 500 of an aircraft flap actuator release test apparatus according to the invention.

According to the displacement simulation unit 500 disclosed by the invention, one sliding table is fixed on the test bed through the clamp, and the displacement of the supporting position of the wing box section in the surface of the driving rocker arm can be simulated 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 stand 50, and a slide table 60. The slide table 60 is fixed to the stage 50 by the jig 40.

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

The slide 67 is slidable on a slide rail 62, and the slide rail 62 is fixed to the fixed mount 61.

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

During the test, the worm gear 64 is 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) slides along the sliding rail 62, so that the required position of the mounting support 63 is obtained.

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

Fig. 6 shows a schematic structural view of a follower loading platform 600 of an aircraft flap actuator release test apparatus according to the invention.

The aircraft flap actuator release test apparatus of the present invention also includes a high load quick response follow-up loading platform 600. As shown, the follower loading platform 600 is divided into two layers, with a plurality of force-controlled loading rams 70 disposed on the upper layer for simulating pneumatic external loads. The force controlled load ram 70 is mounted on an intermediate table 80. The lower tier is provided with a plurality of position control actuators 90 for controlling the attitude of the intermediate table 80 to maintain the intermediate table 80 as consistent as possible with the flap airfoil motion to reduce displacement variation of the first tier load actuators 70. The base 100 is used to support and secure the entire platform.

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

Since the entire disengagement instant is completed within 50ms and the flap is rapidly moved in space at the disengagement instant, the response speed of the actuator is required to be high. By the follower loading platform 600 described above, a quick response under large loads can be simulated.

For a better understanding of the present invention, a method of performing a release test using the aircraft flap actuator release test apparatus of the present invention is described below in connection with FIG. 7.

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

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

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

When the sleeve and the drive link are disengaged, the disengagement mechanism is split into two parts and no mechanical drive is possible. At this time, the flap actuator does not drive the drive link and the flap when driving the sleeve.

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

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

In embodiments of the invention, the sensor may comprise one or more of the following: 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 when the trip failure is triggered.

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

In embodiments of the present invention, it may be verified whether the design meets the goal based on the collected data. In particular, a target state range of the flap after a decoupling failure can be set, in which the flap can still continue to operate. If this range is exceeded, it means that the flap may not continue to function.

It should be noted that the order of the above-described steps of method 700 is exemplary and not limiting. In 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 release 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 test device can transmit larger torque load, so that the simulation is closer to the actual situation.

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

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

5. Compared with the disconnection caused by the prefabricated defect, the parts of the invention can be reused except the explosion bolt, thereby being convenient for daily use and repeated test and having low cost.

6. By using explosive bolts, a quick and controllable release can be achieved to simulate a real transient release fault condition.

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

8. Through the follow-up loading platform, expected movement of the flap after release can be effectively realized, and the wing surface is maintained to be loaded, so that a 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 not "over 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, the use of such phrases may not merely refer to 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. The elements of each aspect described throughout this disclosure are all structural and functional equivalents that are presently or later 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. Additionally, 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 arrangements and instrumentalities 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.