CN112078830A - Aircraft track control method and tail skirt - Google Patents
Aircraft track control method and tail skirt Download PDFInfo
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- CN112078830A CN112078830A CN202010773607.9A CN202010773607A CN112078830A CN 112078830 A CN112078830 A CN 112078830A CN 202010773607 A CN202010773607 A CN 202010773607A CN 112078830 A CN112078830 A CN 112078830A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/242—Orbits and trajectories
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
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Abstract
The invention relates to an aircraft track control method and a tail skirt, wherein the method comprises the following steps: a. determining flight parameters of an aircraft cabin with an adjustable tail skirt during flight, and judging whether the aircraft cabin deviates from a set track; b. if the aircraft cabin deviates from the set track, adjusting the flare angle and the deflection angle of the tail skirt to correct the track, otherwise, keeping the current flight state; c. repeating steps (a) and (b) during the flight of the aircraft cabin until the end of its flight phase. The method of the invention is to modify the orbit of the tail skirt by adjusting the tail skirt. This way the adjustment is smoother and more continuous and saves material.
Description
Technical Field
The invention relates to the field of aircraft track control, in particular to an aircraft track control method and a tail skirt.
Background
A return satellite is a vehicle that can bring items from a space orbit back to the ground. At present, when the aircraft returns to the atmosphere, ballistic and semi-ballistic reentry modes are mainly adopted. Wherein, the return satellite adopting the ballistic reentry mode is not controlled after braking and separating from the orbit; the returning satellite adopting the semi-ballistic reentry mode only controls the reentry orbit through spinning after braking to leave the orbit, and the control capability is extremely limited. Due to the fact that the defects of the control capability of the aircraft orbit cause great deviation of the landing point of the recoverable satellite, the landing area range is enlarged, and difficulty is increased for searching and recovering the recoverable satellite. Therefore, when the recoverable satellite meets an emergency, the time for searching personnel to find the recoverable satellite is greatly increased, and the possibility of danger of articles in the recoverable satellite is also greatly increased. Therefore, in the process of the aircraft, the track of the aircraft is effectively controlled, the ground personnel search time can be effectively shortened, and the safety is improved.
Because the dynamic pressure and the thermal load on the surface of the return satellite are extremely severe in the reentry process, the current return satellite is basically in a rotating body configuration in order to reduce the influence of the dynamic pressure and the thermal load on articles in the cabin. It is difficult to incorporate a conventional aerodynamic control surface for attitude trajectory control in this configuration. For some return satellites that use semi-ballistic reentry, orbit correction is also controlled by using an orbit control engine with better control capability, which can achieve orbit correction but consumes a lot of fuel, and smooth and continuous adjustment of the engine cannot be achieved by start-stop control of the orbit changing. It is of course possible to incorporate a tail skirt arrangement at the trailing edge of the star, but such tail skirts are generally not adjustable and serve only to stabilise the trailing airflow and not for orbital control. Some current aircraft (such as missiles) with smaller bottom surface radius are provided with an adjustable tail skirt at the tail part. For example, patent CN110230958A discloses an expandable tail skirt mechanism, which uses long and short rods to cooperate with inner and outer skirt pieces to form a three-bar mechanism, i.e. moving by linear guide rails. It is known that the skirt is dimensioned to be relatively large in this manner, and this is not suitable for aircraft with a large base radius, such as a recoverable satellite. Forcing its application to a return satellite will cause the tail skirt to increase in size and weight significantly beyond acceptable limits.
Disclosure of Invention
The invention aims to provide an aircraft track control method and a tail skirt, which can continuously and smoothly adjust a track.
In order to achieve the purpose, the invention provides an aircraft track control method and a tail skirt, wherein the method comprises the following steps:
a. determining flight parameters of an aircraft cabin with an adjustable tail skirt during flight, and judging whether the aircraft cabin deviates from a set track;
b. if the aircraft cabin deviates from the set track, adjusting the flare angle and the deflection angle of the tail skirt to correct the track, otherwise, keeping the current flight state;
c. repeating steps (a) and (b) during the flight of the aircraft cabin until the end of its flight phase.
According to one aspect of the invention, in the step (a), the flight parameters include position, attitude and flight direction, which are measured by an inertial measurement unit, and the position includes altitude and latitude and longitude.
According to one aspect of the invention, the navigation system is used to determine whether the aircraft cabin deviates from the predetermined trajectory, and the deviation from the predetermined trajectory is determined when the speed deviation reaches 0.5m/s or the position deviation reaches 1 m.
According to one aspect of the invention, in step (b), if the aircraft cabin deviates from the predetermined trajectory, the current altitude of the aircraft cabin is further obtained, as well as the speed deviation and/or the position deviation.
According to one aspect of the invention, the direction and magnitude of the force required to be applied to the aircraft cabin are analyzed according to the current speed deviation or position deviation of the aircraft cabin, and the formula for calculating the force by using the speed deviation is as follows:
F＝mΔV/t；
wherein F is force, Δ V is speed deviation, t is tail skirt action time, and m is mass of the recoverable satellite;
the formula for calculating force using positional deviation is:
F＝mX/t2；
wherein F is force, X is position deviation, and the calculation formula is
t is the acting time of the tail skirt, and m is the mass of the recoverable satellite;
and calculating to obtain the dynamic pressure at the moment according to the height and the speed of the current aircraft cabin body, wherein the calculation formula is as follows:
dp＝1/2·ρV2；
wherein rho is air density, obtained according to the current altitude and an atmosphere database, V is satellite speed, and dp is dynamic pressure;
the coefficient of force is calculated using the following calculation formula:
Cf＝F/(S·dp)；
wherein S is a reference area of the recoverable satellite;
and obtaining the angle required to be adjusted by the current tail skirt according to the fitting relation of the force coefficient and the opening angle and deflection angle of the tail skirt.
The tail skirt, including regulating plate and driving piece, the regulating plate is arranged along the circumferencial direction interval, the driving piece is located the shape that the regulating plate encloses is inboard, still including adjusting the ring, is located the shape that the regulating plate encloses is inboard, adjust the ring and every the regulating plate is all connected, the driving piece is arranged and rather than the facing along the circumferencial direction interval the regulating plate is connected.
According to one aspect of the invention, a through groove for guiding and limiting the adjusting ring is arranged on the plate surface of the adjusting plate, and the guiding direction of the through groove is parallel to the length direction of the adjusting plate;
the adjusting ring is provided with a clamping structure which can move in the through groove on the adjusting plate.
According to one aspect of the invention, the clamping structure is composed of a cylinder fixedly connected to the outer side of the adjusting ring and a baffle positioned at one end of the cylinder far away from the adjusting ring.
According to one aspect of the invention, the snap-fit structure is a bolt screwed onto the adjustment ring from the outside thereof.
According to one aspect of the invention, the drive member is a hydraulic ram;
one end of the adjusting plate is connected with a ball hinge.
According to one aspect of the invention, the device further comprises a controller for controlling the driving member, and an inertial measurement unit and a navigation unit connected with the controller.
According to one aspect of the invention, the connecting point of the driving member and the adjusting plate is located at 10% -30% of the length direction of the adjusting plate, and the connecting point is lower than the lowest end of the movable range of the adjusting ring.
According to one aspect of the invention, the number of said regulating plates is between 30 and 60;
3 adjusting plates are arranged between adjacent driving parts at intervals;
the adjusting ring is made of aluminum alloy or titanium alloy.
According to one scheme of the invention, the flight orbit of the aircraft is corrected by adjusting the opening angle and the deflection angle of the tail skirt, so that compared with an orbit control engine, the invention saves more fuel and can realize smooth and continuous orbit adjustment.
According to one aspect of the invention, the trailing skirt utilizes a hydraulic ram as a drive member for rotating the adjustment plate. This reduces the size requirements for the adjustment plate. The adjusting ring is arranged in the adjusting plate, is positioned on the inner side of the shape surrounded by the adjusting plates and is connected with all the adjusting plates. The adjusting plate is also provided with a through groove, so that the limiting and guiding of the adjusting ring can be completed by matching with a clamping structure on the adjusting ring. Therefore, when one driving part drives the adjusting plate connected with the driving part to act, the adjusting ring can drive the other adjusting plates to follow up, and the quantity requirement of the driving parts is reduced.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings needed in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings without creative efforts.
FIG. 1 schematically illustrates a flow chart of an aircraft trajectory control method of one embodiment of the present invention;
FIG. 2 is a schematic representation of an axial force coefficient versus trailing skirt opening angle using a fit in an aircraft trajectory control method according to an embodiment of the present invention;
FIG. 3 is a schematic representation of a lateral force coefficient versus tailskirt drift angle using a fit in an aircraft trajectory control method in accordance with an embodiment of the present invention;
FIG. 4 is a schematic representation of pressure clouds and a mach number cloud of a symmetric surface of a returnable satellite at different tail skirt opening angles;
FIG. 5 schematically illustrates a return satellite isometric view utilizing a tail skirt of one embodiment of the present invention;
figure 6 schematically illustrates a bottom block diagram of a recoverable satellite utilizing a tailskirt according to one embodiment of the present invention;
FIG. 7 is a schematic representation of the inboard connection of the tail skirt in accordance with one embodiment of the present invention;
fig. 8 is a schematic diagram showing the connection relationship between the outer sides of the tail skirt according to the embodiment of the present invention.
Detailed Description
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the embodiments will be briefly described below. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
In describing embodiments of the present invention, the terms "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in an orientation or positional relationship that is based on the orientation or positional relationship shown in the associated drawings, which is for convenience and simplicity of description only, and does not indicate or imply that the referenced device or element must have a particular orientation, be constructed and operated in a particular orientation, and thus, the above-described terms should not be construed as limiting the present invention.
The present invention is described in detail below with reference to the drawings and the specific embodiments, which are not repeated herein, but the embodiments of the present invention are not limited to the following embodiments.
Referring to fig. 1, the aircraft trajectory control method of the present invention is implemented by adjusting the angle of the tail skirt. The method is mainly used for re-entry orbit correction of the recoverable satellite, but can also be used for controlling orbits of other aircrafts. Because the method utilizes the tail skirt to control the flight attitude of the aircraft, the aircraft to which the method aims is required to be provided with a flight main body and the tail skirt. In the present invention, the main flying body is an aircraft cabin, and in the present embodiment, the main flying body is a satellite cabin. The stability of recoverable satellite flight can be improved to the flare angle of adjusting the tail skirt. And the satellite orbit can be changed by adjusting the deflection angle of the tail skirt. In fact, the opening angle and the deflection angle can cause the aerodynamic force of the whole recoverable satellite to change, and the reentry control of the recoverable satellite into the orbit can be realized through the change of the aerodynamic force, so that the accuracy of the landing point of the recoverable satellite is improved, the recovery time of the recoverable satellite is shortened, and the reentry and return safety is improved.
In the method, the flight parameters of the aircraft cabin are measured in real time during the whole flight period of the aircraft cabin. The flight parameters include position, attitude and flight direction. All three are determined using an Inertial Measurement Unit (IMU). The invention uses the height and longitude and latitude of the flight cabin body to represent the position information. While the flight parameters are measured, a navigation unit (which may be a GPS or the like) determines whether the satellite capsule deviates from a predetermined orbit. If the satellite is judged at a certain momentIf the cabin deviates from the predetermined orbit, the altitude, position deviation and/or speed deviation at that moment are determined. The height is already included in the position information, and two deviations are measured by the navigation unit, so that only three parameters at this moment need to be acquired. Due to the passing of the formula
Since the positional deviation can be obtained by the speed deviation, only the speed deviation may be measured. Since the tail skirt has a short operating time, the above formula may be X ═ Δ V t. In the invention, when the speed deviation of the cabin body monitored by the return satellite control system (namely the master control system) reaches 0.5m/s or the position deviation reaches 1m, the deviation from the orbit is judged, and a subsequent adjusting program is started. Before that, the satellite cabin interior processing system performs calculation to convert the acquired height and the two deviations into the tail skirt angle to be adjusted. Specifically, the direction and magnitude of the force required to be applied to the satellite capsule (i.e., the force to be applied to the adjustment orbit) are analyzed according to the current velocity deviation or position deviation of the satellite capsule, and the required force is divided into an axial force and a lateral force, which have the same calculation formula and are different only in parameters, so the following description will be given by taking the calculation of the axial force as an example.
The formula for calculating the magnitude of the axial force by using the speed deviation is as follows:
F＝mΔV/t；
wherein F is an axial force, Δ V is a speed deviation, t is a tail skirt acting time, and m is the mass of the recoverable satellite;
the formula for calculating the magnitude of the axial force by using the position deviation is as follows:
F＝mX/t2；
wherein F is an axial force, X is a position deviation, t is the acting time of the tail skirt, and m is the mass of the recoverable satellite; (the formula is the constant velocity situation)
And calculating to obtain the dynamic pressure at the moment according to the height and the speed of the current satellite cabin, wherein the calculation formula is as follows:
dp＝1/2·ρV2；
where ρ is the air density, obtained from the current altitude and the atmospheric database, V is the satellite velocity (measured at Δ V above) and dp is the dynamic pressure;
calculating an axial force coefficient according to the axial force, wherein the calculation formula is as follows:
Cf＝F/(S·dp)；
where S is the reference area of the recoverable satellite.
The velocity and the respective deviations in the above calculation are components in the axial direction. And obtaining an axial force coefficient through the calculation, and then obtaining the angle required to be adjusted by the current tail skirt according to the axial force coefficient. In the invention, the axial force coefficient and the flare angle of the tail skirt are fitted to obtain a relation graph shown in figure 2. In fig. 2, a refined CFD simulation is performed on reentry states of the recoverable satellite with different tail skirt opening angle configurations, where the simulation condition is a reentry angle of 2 °, the flight speed is mach 1.6, and the height is 10 km. The relationship between the axial force coefficient and the flare angle of the tail skirt is obtained through simulation, the axial force coefficient is approximately linearly increased along with the increase of the flare angle of the tail skirt, and the controllability of the swinging tail skirt can be improved through the monotonous linear relationship. The reason for this change is shown in fig. 4, the distribution of the pressure on the surface of the satellite cabin does not change greatly as the opening angle of the tail skirt increases, but the pressure on the outer surface of the skirt sheet of the tail skirt increases as the opening angle of the tail skirt increases, and the axial force generated by the integral of the partial surface pressure in the axial projection also increases. The calculation process and the relation chart can obtain the opening angle required to be adjusted, and the adjustment of the deflection angle needs to calculate the lateral force coefficient. The calculation process of the lateral force coefficient is the same as that of the axial force coefficient, only the axial components of the speed and each deviation are changed into the lateral components, and finally the relationship between the lateral force and the deflection angle is obtained as shown in fig. 3. Accordingly, the axial force coefficient and the lateral force coefficient obtained by the calculation correspond to two relation maps to obtain the currently required tail skirt angle, wherein the angle comprises the opening angle and the deflection angle. The adjustment of the opening angle plays a role in stabilizing the flight, and the deflection angle plays a leading role in orbital transfer. In this step, if the satellite capsule does not deviate from the predetermined orbit, the current flight state is maintained until the satellite capsule deviates from the orbit and then the adjustment is performed. In this way, before the satellite capsule arrives at the destination, the measurement is continuously carried out, and the flight orbit of the satellite capsule is adjusted and corrected repeatedly according to the measurement result until the flight phase of the satellite capsule arrives at the destination.
Referring to fig. 5, the tail skirt of the present invention comprises an adjusting plate 1, a driving member 2 and an adjusting ring 3, which form the tail skirt of the satellite capsule. Some aircraft are also generally not equipped with a controller for controlling the tail skirt if they are not equipped with a tail skirt. The tail skirt of the present invention may also include a controller. As shown in fig. 5, the tail skirt device is disposed at the tail of the satellite cabin a, the structure of the satellite cabin a is not improved, and only the controller for controlling the tail skirt device is integrated into the original control system of the satellite cabin a. The adjusting plate 1 serves as the main component of the tail skirt, which can also be understood as a skirt piece of the tail skirt. Different from the traditional tail skirt, the adjusting plates 1 are arranged at intervals along the circumferential direction, so that the tail skirt is of a single-layer structure and is simple in design. In the present embodiment, the shape of the adjustment plate 1 is almost flat. According to the method, the satellite orbit correction process simultaneously involves the adjustment of the opening angle and the deflection angle of the tail skirt, and the opening angle and the deflection angle of the tail skirt can be changed by rotating the adjusting plate 1. Therefore, the invention provides a ball hinge at the end of the adjusting plate 1, so that it can be hinged to the satellite capsule a. And the respective adjusting plates 1 are spaced apart to prevent interference thereof when rotated. From this configuration, the adjustment accuracy is higher as the number of adjustment plates 1 is larger. However, an excessive number also places a certain burden on other components, for example, the driving force or the number requirement of the drive element 2 increases. The diameter of the tail edge of a general satellite cabin is 1-2m, so that the number of the adjusting plates 1 is 30-60, and the burden of other parts can be reduced on the premise of ensuring the precision.
Referring to fig. 6, according to the present invention, the adjusting plates 1 are driven to rotate by the driving member 2, so as to complete the adjustment of the opening angle and the deflection angle. In this embodiment, the driving member 2 is a hydraulic actuator cylinder, one end of the hydraulic actuator cylinder is hinged to the adjusting plate 1, and the other end of the hydraulic actuator cylinder is connected to the bottom of the satellite cabin a, which has the advantage that the sizes of the adjusting plate 1 and the driving member 2 do not need to be too large, so that the tail skirt device can be applied to the satellite when the satellite returns. In the invention, the driving pieces 2 are arranged along the circumferential direction and are all arranged inside the shape surrounded by the adjusting plate 1. Since too high a connection point of the driver 2 to the adjustment plate 1 results in an oversized adjustment plate 1, while too low a connection point results in an increased drive torque requirement for the driver 2, the connection point of the driver 2 to the adjustment plate 1 in the present invention is located at 10% -30%, especially 20%, of the length of the adjustment plate 1. In addition, since each adjusting plate 1 is independent and has a large number, if each adjusting plate 1 is connected to one driving member 2, the weight and cost of the tail skirt device are increased. For this purpose, an adjusting ring 3 is also provided in the invention to assist the adjustment.
Referring to fig. 7, the adjusting ring 3 is a standard circular ring, and the material thereof is a rigid material, such as an aluminum alloy or a titanium alloy. An adjusting ring 3 is connected with each adjusting plate 1 at the inner side of the adjusting plate 1. Therefore, when the driving piece 2 drives one adjusting plate 1 to act, the adjacent adjusting plates 1 can be driven to act together. In this way, the driving members 2 can be spaced apart from each other, so that the number of the driving members can be reduced, and uniform linkage adjustment can be realized. The number of the adjusting plates 1 spaced between the adjacent driving members 2 depends on the precision requirement of the tail skirt deflection angle, and as mentioned above, the smaller the number of the adjusting plates 1 spaced, the more precise the adjustment, but the smaller the number of the spacing, the less the number of the adjusting plates 1 spaced, the less the number of the driving members 2. Therefore, in the present embodiment, 3 adjusting plates 1 are spaced between adjacent drivers 2. From this function, the adjusting plate 1 directly connected to the driving member 2 can be called as a driving plate, and the rest is called as a driven plate.
In order to match with the adjusting function of the adjusting ring 3, the adjusting plate 1 is provided with a through groove 11 with guiding and limiting functions. The through groove 11 penetrates through the thickness of the adjusting plate 1 and is in a strip shape. The length direction is the guiding direction, and the guiding direction is parallel to the length direction of the adjusting plate 1. Meanwhile, a clamping structure 31 matched with the through groove 11 is also arranged on the adjusting ring 3. Referring to fig. 8, in the present embodiment, the latch structure 31 has a column 31a and a stop 31b at the end of the column 31 a. The other end of the cylinder 31a is fixed to the outside of the adjustment ring 3, and the cylinder 31a is movable in the through groove 11. According to the structure, the clamping structure 31 can also be directly a bolt, so that the bolt directly penetrates through the through groove 11 and is screwed into the outer side of the adjusting ring 3. According to the design, when the tail skirt is adjusted, the controller sends a command to the driving piece 2 so as to drive the adjusting plate 1 to actuate. When the opening angle is adjusted, the driving adjusting plate 1 is driven to move by the driving piece 2, so that the adjusting ring 3 moves along with the driving adjusting plate. The moving track of the satellite cabin body A is that the satellite cabin body A makes linear motion along the axial direction of the satellite cabin body A, so that the tail skirt is driven to contract or expand integrally. During the declination angle adjustment, the driving part 2 drives the active adjusting plate 1 to drive the adjusting ring 3 to move, at the moment, the moving track of the adjusting ring 3 does not only do the linear motion, but also do the translation or the inclined motion in the radial direction of the satellite cabin body A, so that the adjusting plate 1 on one side rotates outwards, the adjusting plate on the other side rotates inwards, the declination angle adjustment is realized, the existence of the ball hinge can enable the adjusting plate 1 to have higher degree of freedom and rotate in all directions, and the locking and the fracture are avoided. Therefore, the tail skirt device controls the track by utilizing the adjusting mode of the tail skirt piece, and compared with an orbit control engine, the adjusting process is smoother and continuous. From the above, the present invention provides a limitation on the height of the position of the connection point of the driving member 2 and the adjusting plate 1, and the connection point of the driving member 2 and the adjusting plate 1 is lower than the lowest point of the movable range of the adjusting ring 3 due to the existence of the adjusting ring 3, so as to avoid collision between the driving member and the adjusting plate during the adjustment process.
In summary, the tail skirt of the present invention can be used as a tail skirt for correcting the re-entry orbit after being installed on the satellite cabin a. The invention is corresponding to the satellite cabin, and the general satellite cabin is provided with an inertia measurement unit and a navigation unit, so that when the satellite cabin is installed, the controller is only embedded into the satellite cabin and connected with the two units and/or the cabin master control system to form a closed loop with feedback regulation, and when the two units or the master control system detect that the satellite deviates from the orbit, the satellite master control can automatically send an instruction to the controller to control the driving piece 2 to actuate, thereby realizing the accurate control of the orbit. If the applied aircraft does not have the two measuring units, the tail skirt can be additionally provided with an inertial measuring unit and a navigation unit which are connected with the controller, so that the rail control method is realized. According to this concept, if the corresponding aircraft itself is also provided with a tail skirt structure, such as the missile described above. It can also be directly modified from its original structure to minimize its weight and cost. Moreover, the adjusting process can be completed in a short time, so that the re-orbit of the returning satellite can be quickly and accurately controlled, and the returning satellite finally lands at a preset place.
The above description is only one embodiment of the present invention, and is not intended to limit the present invention, and it is apparent to those skilled in the art that various modifications and variations can be made in the present invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (13)
1. An aircraft trajectory control method comprising the steps of:
a. determining flight parameters of an aircraft cabin with an adjustable tail skirt during flight, and judging whether the aircraft cabin deviates from a set track;
b. if the aircraft cabin deviates from the set track, adjusting the flare angle and the deflection angle of the tail skirt to correct the track, otherwise, keeping the current flight state;
c. repeating steps (a) and (b) during the flight of the aircraft cabin until the end of its flight phase.
2. The trajectory control method according to claim 1, wherein in the step (a), the flight parameters include a position, an attitude, and a flight direction, which are measured by an inertial measurement unit, and the position includes an altitude and a latitude and longitude.
3. The trajectory control method according to claim 2, wherein a navigation system is used to determine whether the aircraft cabin deviates from the predetermined trajectory, and the deviation is determined when the speed deviation reaches 0.5m/s or the position deviation reaches 1 m.
4. The trajectory control method according to claim 3, wherein in step (b), the current altitude of the aircraft cabin and the speed deviation and/or the position deviation are further obtained if the aircraft cabin deviates from the predetermined trajectory.
5. The track control method according to claim 4, wherein the direction and magnitude of the force applied to the aircraft cabin are analyzed according to the current speed deviation or position deviation of the aircraft cabin, and the formula for calculating the force by using the speed deviation is as follows:
F＝mΔV/t；
wherein F is force, Δ V is speed deviation, t is tail skirt action time, and m is mass of the recoverable satellite;
the formula for calculating force using positional deviation is:
F＝mX/t2；
wherein F is force, X is position deviation, and the calculation formula is
t is the acting time of the tail skirt, and m is the mass of the recoverable satellite;
and calculating to obtain the dynamic pressure at the moment according to the height and the speed of the current aircraft cabin body, wherein the calculation formula is as follows:
dp＝1/2·ρV2；
wherein rho is air density, obtained according to the current altitude and an atmosphere database, V is satellite speed, and dp is dynamic pressure;
the coefficient of force is calculated using the following calculation formula:
Cf＝F/(S·dp)；
wherein S is a reference area of the recoverable satellite;
and obtaining the angle required to be adjusted by the current tail skirt according to the fitting relation of the force coefficient and the opening angle and deflection angle of the tail skirt.
6. A tail skirt for implementing the track control method according to any one of claims 1 to 5, comprising adjusting plates (1) and driving members (2), wherein the adjusting plates (1) are arranged at intervals along the circumferential direction, the driving members (2) are arranged inside the shape surrounded by the adjusting plates (1), and the tail skirt is characterized by further comprising adjusting rings (3) arranged inside the shape surrounded by the adjusting plates (1), the adjusting rings (3) are connected with each adjusting plate (1), and the driving members (2) are arranged at intervals along the circumferential direction and are connected with the adjusting plates (1) facing the adjusting rings.
7. The tail skirt according to claim 6, characterized in that the adjusting plate (1) is provided with a through groove (11) on the plate surface for guiding and limiting the adjusting ring (3), and the guiding direction of the through groove (11) is parallel to the length direction of the adjusting plate (1);
the adjusting ring (3) is provided with a clamping structure (31) which can move in the through groove (11) on the adjusting plate (1) on the outer side.
8. The tailskirt according to claim 7, characterized in that the clamping structure (31) comprises a cylinder (31a) fixedly connected to the outside of the adjusting ring (3) and a baffle (31b) located at an end of the cylinder (31a) far away from the adjusting ring (3).
9. The tailskirt according to claim 7, characterized in that the snap-in structure (31) is a bolt screwed onto the adjusting ring (3) from outside thereof.
10. A tail skirt according to claim 6, wherein the drive member (2) is a hydraulic ram;
one end of the adjusting plate (1) is connected with a ball hinge.
11. The tailskirt according to claim 6, characterized in that it further comprises a controller for controlling the drive (2) and an inertial measurement unit and a navigation unit connected to the controller.
12. The tailskirt according to claim 6, characterized in that the connecting point of the driving member (2) and the adjusting plate (1) is located at 10% -30% of the length of the adjusting plate (1) and is lower than the lowest end of the movable range of the adjusting ring (3).
13. The tailskirt according to claim 6, characterized in that the number of regulating plates (1) is between 30 and 60;
3 adjusting plates (1) are arranged between adjacent driving pieces (2);
the adjusting ring (3) is made of aluminum alloy or titanium alloy.
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| SU1818283A1 (en) * | 1990-12-20 | 1993-05-30 | K B Salyut | Ballistic recoverable capsule |
| CN1530288A (en) * | 2003-03-12 | 2004-09-22 | 秦为国 | Circular magnetic suspension flying device |
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| CN107891979A (en) * | 2017-09-28 | 2018-04-10 | 中国运载火箭技术研究院 | A kind of hypersonic aircraft can adjust tranquilizer |
| CN110230958A (en) * | 2019-07-22 | 2019-09-13 | 哈尔滨工业大学 | A kind of expandable type Wei Qun mechanism |
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| SU1818283A1 (en) * | 1990-12-20 | 1993-05-30 | K B Salyut | Ballistic recoverable capsule |
| CN1530288A (en) * | 2003-03-12 | 2004-09-22 | 秦为国 | Circular magnetic suspension flying device |
| CN1978279A (en) * | 2005-12-05 | 2007-06-13 | 罗进南 | Efficient flying boat and missile with tail-wing-skirt self-stabilizing return capsule |
| CN107891979A (en) * | 2017-09-28 | 2018-04-10 | 中国运载火箭技术研究院 | A kind of hypersonic aircraft can adjust tranquilizer |
| CN110230958A (en) * | 2019-07-22 | 2019-09-13 | 哈尔滨工业大学 | A kind of expandable type Wei Qun mechanism |
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