WO2021143428A1 - 用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法 - Google Patents

用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法 Download PDF

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Publication number
WO2021143428A1
WO2021143428A1 PCT/CN2020/136506 CN2020136506W WO2021143428A1 WO 2021143428 A1 WO2021143428 A1 WO 2021143428A1 CN 2020136506 W CN2020136506 W CN 2020136506W WO 2021143428 A1 WO2021143428 A1 WO 2021143428A1
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WO
WIPO (PCT)
Prior art keywords
satellite
locking
ejection
boss
electromagnet
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PCT/CN2020/136506
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English (en)
French (fr)
Inventor
王兆魁
何云瀚
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清华大学
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Publication date
Application filed by 清华大学 filed Critical 清华大学
Priority to US17/295,047 priority Critical patent/US11597540B2/en
Publication of WO2021143428A1 publication Critical patent/WO2021143428A1/zh

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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
    • B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
    • B64G1/645—Separators
    • B64G1/6457—Springs; Shape memory actuators
    • 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
    • B64G1/64—Systems for coupling or separating cosmonautic vehicles or parts thereof, e.g. docking arrangements
    • B64G1/645—Separators
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/002—Launch systems
    • B64G1/005—Air launch

Definitions

  • the invention belongs to the field of aerospace technology, and specifically relates to an electromagnetic lock release mechanism and an electromagnetic lock release method for separating stars and arrows.
  • the key design elements of the satellite lock release mechanism are the mechanism form and active components.
  • the mechanism form mainly adopts the docking ring and the belt-type separation mechanism.
  • the main problem is: the docking ring and the belt-type separation mechanism require the satellite to have a docking surface connected to the separation mechanism for installation, which cannot be applied to Irregularly shaped satellites that require point positioning and locking.
  • Active components mainly use fire attack products, hot knives, memory alloys, motors, etc.
  • the main problems are: the high cost of fireworks, complicated management requirements, and destructive separation, which cannot be repeated on the ground; hot knives can
  • the load is very small, it is difficult to use for the fixation of the entire satellite, and the action time is uncertain, and the synchronization cannot be guaranteed; the memory alloy action time is uncertain, and the synchronization cannot be guaranteed; the motor needs continuous power supply, and the micro-satellites are generally launched on rockets. Rockets generally cannot provide continuous current, so it is difficult to apply them to tiny satellites.
  • the present invention provides an electromagnetic lock release mechanism and an electromagnetic lock release method for separating stars and arrows, which can effectively solve the above-mentioned problems.
  • the invention provides an electromagnetic lock release mechanism for separating stars and arrows, including: a frame (1), an ejection unit, a satellite unit, a lock release unit, and a lock release drive unit; the horizontal direction is set to the X direction; and the vertical direction is set to be the X direction.
  • the direction is Y direction;
  • the satellite unit includes a satellite (7), a positioning pin (8) and a satellite boss (9); the satellite (7) is fixedly connected with the satellite boss (9) through the positioning pin (8);
  • the ejection unit is arranged below the satellite unit and is used to provide the satellite unit with ejection force in the positive Y direction;
  • the ejection unit includes an ejection spring (2), an ejection jack (3), and a spring sleeve ( 4);
  • the spring sleeve (4) is fixedly connected to the frame (1); the spring sleeve (4) is provided with the ejection spring (2); the bottom of the ejection rod (3) Is located in the spring sleeve (4) and presses on the ejection spring (2); the top of the ejection pin (3) passes through the spring sleeve (4) and extends to the Under the satellite unit and in contact with the lower surface of the satellite unit; in the locked state, the ejection spring (2) is pressed by the ejection mandrel (3), and the ejection spring (2) passes through the
  • the ejection jack (3) provides ejection force in the positive direction of Y to the satellite unit; in the released state, under the ejection force of the
  • the locking release unit includes a locking pin (10), a locking slider (11), an unlocking spring (12) and a base (13);
  • the base (13) is located on the left side of the satellite unit and is fixedly connected to the frame (1); the right side of the base (13) is provided with a first cavity, and the locking pin ( 10) is a D shape; the upper surface of the satellite boss (9) is designed as a slope; in the locked state, the locking pin (10) is partially located in the first cavity, and the locking pin
  • the cylindrical surface of (10) presses on the inclined surface of the satellite boss (9), thereby restricting the movement of the satellite boss (9) in the positive direction of Y; in addition, the inclined surface of the satellite boss (9)
  • the locking pin (10) provides contact elastic force along the normal direction of the inclined plane; the contact elastic force is decomposed into the contact elastic force in the positive direction of Y and the contact elastic force in the negative direction of X; the contact elastic force in the positive direction of Y causes the locking pin (10) to compress the position The base (13); the contact elastic force in the negative X direction causes the locking pin (10) to press the locking slider (11) behind it;
  • a guide groove along the Y direction is provided on the left side of the base (13), and the locking slide block (11) is arranged in the guide groove, so that the locking slide block (11) can only move in the Y direction Movement;
  • the unlocking spring (12) is arranged between the locking slider (11) and the base (13), and the unlocking spring (12) provides Y positive to the locking slider (11) Thrust in direction;
  • the lower part of the locking slider (11) is provided with a cavity along the X direction; the locking slider (11) is located behind the locking pin (10), and in the locked state, the locking The upper plane part of the slider (11) is pressed behind the locking pin (10); in the released state, the locking slider (11) moves in the positive Y direction under the action of the unlocking spring (12) Move, so that the cavity of the locking slider (11) rises, and the locking pin (10) moves along the negative X direction under the action of the contact elastic force in the negative X direction provided by the satellite boss (9) Move into the cavity of the locking slider (11), so that the locking pin (10) releases the obstacle to the movement of the satellite unit;
  • the lock release drive unit includes an electromagnet limit nut (14), an electromagnet moving core (15) and an electromagnet (16); the electromagnet (16) is fixed to the frame (1);
  • the electromagnet (16) has a built-in coil; the electromagnet moving core (15) is arranged along the X direction.
  • the limit bracket (17) is fixedly installed above the locking slider (11) to limit the locking slider (11) to Y positive The distance moved in the direction.
  • it also includes a pre-tensioning unit
  • the pre-tensioning unit includes a pre-tensioning member bracket (5) and a pre-tensioning member (6);
  • the pretensioner support (5) is fixedly connected to the frame (1); the pretensioner (6) is connected to the pretensioner support (5) by a pretension screw; and, the pretensioner
  • the piece (6) is located under the satellite boss (9); by rotating the pre-tightening screw, the pre-tensioning piece (6) is moved in the positive direction of Y until the upper part of the pre-tensioning piece (6) is The surface is in close contact with the lower surface of the satellite boss (9); at the same time, during the tightening of the pre-tightening screw, the pre-tensioning member (6) pushes the satellite boss (9) to move in the positive Y direction , The contact between the satellite boss (9) and the locking pin (10) is elastically deformed, and the contact gap is eliminated.
  • the present invention provides an electromagnetic lock release method for an electromagnetic lock release mechanism for separating stars and arrows, which includes the following steps:
  • Step 1 Before the rocket reaches the predetermined orbit, the satellite and the rocket are locked in the following ways:
  • Step 1.1 the built-in coil of the electromagnet (16) is energized, so that the right end of the electromagnet moving core (15) is pressed against the top of the locking slider (11), and the locking slider (11) is located below the guide groove;
  • Step 1.2 when the locking slider (11) is located below the guide groove, the upper flat part of the locking slider (11) is pressed behind the locking pin (10) to restrict the locking pin (10) from being negative along X Move in the direction so that the right end of the locking pin (10) protrudes from the first cavity of the base (13);
  • Step 1.3 therefore, the cylindrical surface of the locking pin (10) is pressed against the inclined surface of the satellite boss (9), thereby restricting the satellite boss (9) from moving in the positive direction of Y; at the same time, the ejection pin (3) is in the ejection spring Under the action of (2), the ejection force in the positive direction of Y is provided to the satellite unit, but the ejection force is less than the force exerted by the locking pin (10) to the satellite boss (9); through the locking pin (10) and ejection rod
  • the joint action of (3) makes the satellite unit locked on the rack (1);
  • Step 2 After the rocket enters orbit, a separation instruction is given, so that the satellite can be released safely, which specifically includes the following steps:
  • Step 2.1 after the rocket enters orbit, give a separation instruction, so that the built-in coil of the electromagnet (16) is de-energized;
  • the electromagnet movable core (15) moves in the negative X direction, so that the right end of the electromagnet movable core (15) leaves the top of the locking slider (11), releasing the The limit function of the locking slider (11) in the Y direction;
  • Step 2.2 when the electromagnet moving core (15) releases the limiting effect on the locking slider (11) in the Y direction, the locking slider (11) moves in the positive Y direction under the action of the unlocking spring (12) , And limit the movement distance of the locking slider (11) through the limit bracket (17), so that the cavity of the locking slider (11) rises to a certain height;
  • Step 2.4 when the locking pin (10) is out of the limit on the satellite boss (9) along the positive direction of Y, under the ejection force of the ejection spring (2), make the ejection pin (3) along the positive direction of Y Movement, so that the satellite unit including the satellite boss (9) is ejected and launched in the positive direction of Y, so as to realize the safe release of the satellite.
  • the invention is a point-type positioning and locking release mechanism that can be used to separate small satellites and rockets for repeatable tests. It can achieve reliable locking and reliable separation of satellites and rockets in a complex mechanical environment, and can be repeatedly separated on the ground.
  • Figure 1 is a cross-sectional view of the electromagnetic lock release mechanism for separating stars and arrows provided by the present invention in a locked state;
  • FIG. 2 is a cross-sectional view of the electromagnetic lock release mechanism for separating stars and arrows provided by the present invention in a released state;
  • Figure 3 is an assembly diagram of the electromagnetic lock release mechanism for separating stars and arrows provided by the present invention
  • Figure 4 is a schematic diagram of the bottom plate
  • Figure 5 is a schematic diagram of a support arm
  • Figure 6 is an assembly diagram of the support arm and the bottom plate
  • Figure 7 is a structural diagram formed after assembly in step 3.
  • Figure 8 is a structural diagram formed after step 4 assembly
  • Figure 9 is a structural diagram formed after step 5 assembling.
  • Figure 10 is a diagram of the placement position of the pretensioner tooling
  • Figure 11 is a diagram of the placement position of the gasket
  • Figure 12 is a structural diagram formed after step 6 assembly
  • Figure 13 is a structural diagram formed after step 7 assembly
  • Figure 14 is a structural diagram formed after step 8 assembly
  • Figure 15 is a structural diagram formed after step 9 assembly
  • Figure 16 is a structural diagram formed after assembly in step 10.
  • Figure 17 is a structural diagram formed after step 11 assembly
  • Figure 18 is a structural diagram formed after step 12 assembly
  • Figure 19 is a structural diagram formed after assembly in step 13;
  • Figure 20 is the spacing control diagram
  • Figure 21 is a structural diagram formed after step 14 is assembled.
  • Figure 22 is a structural diagram formed after assembly in step 15;
  • Figure 23 is a schematic diagram of the distance between the outermost side of the satellite and the innermost side of the base opposite to the surface of 0.9mm;
  • Fig. 24 is a schematic diagram when the distance between the side surface of the pin and the inner wall of the base groove is 6 mm.
  • the present invention provides an electromagnetic lock release mechanism for the separation of satellites and rockets, which is used for the lock release of satellites and rockets. It is a point-type positioning lock release mechanism for the separation of micro satellites and rockets that can be tested repeatedly. : In the normal state, it is the locked state, and after the command is given, it becomes the separated release state.
  • the frame, the stationary fixed part, is fixedly connected with the rocket, plays a supporting role, and changes as the shape of the satellite 7 changes.
  • the satellite unit includes a satellite 7, a positioning pin 8 and a satellite boss 9.
  • the satellite 7 is fixedly connected to the satellite boss 9 through the positioning pin 8, and interacts with the rest of the organization through the satellite boss 9, that is: the use of this organization requires Find a suitable position on the satellite 7 to install the positioning pin 8 and the satellite boss 9.
  • the satellite 7, the positioning pin 8 and the satellite boss 9 are integrally connected to form a whole, and when the satellite is separated, the satellite unit is ejected and launched as a whole.
  • the satellite 7 is a moving part, which is the target of the locking and releasing mechanism, and the locking and releasing mechanism does not limit the specific form of the satellite.
  • the ejection unit is arranged below the satellite unit and is used to provide the satellite unit with ejection force in the positive Y direction;
  • the ejection unit includes an ejection spring 2, an ejection jack 3 and a spring sleeve 4;
  • the spring sleeve 4 is fixedly connected with the frame 1;
  • the spring sleeve 4 is provided with an ejection spring 2;
  • the bottom of the ejection pin 3 is located in the spring sleeve 4 and presses on the ejection spring 2;
  • the top of the ejection pin 3 passes through the spring sleeve 4 and extends to the satellite unit Below and in contact with the lower surface of the satellite unit, specifically in contact with the lower surface of the satellite boss 9;
  • the ejection spring 2 In the locked state, the ejection spring 2 is compressed by the ejection plunger 3, the ejection spring is in a compressed state, and the elastic potential energy stored after compression is the energy source of the satellite separation. Therefore, the ejection spring 2 is directed to the satellite unit through the ejection plunger 3 Provides the ejection force in the positive Y direction, but in the locked state, the locking function of the lock release unit provides the force in the negative Y direction to the satellite unit, thereby overcoming the ejection force of the ejection spring 2 and making the satellite unit and the rack 1 Lock; In the released state, the lock release unit releases the force on the satellite unit. Therefore, under the ejection force of the ejection spring 2, the satellite unit moves in the positive direction of Y and separates from the frame 1.
  • the lock release unit mainly includes a lock function and a release function.
  • the general principle is:
  • the upper surface of the satellite boss 9 is designed as an inclined plane, and the inclined plane angle ⁇ is a design value.
  • the inclined surface of the satellite boss 9 is in contact with the locking pin 10, and the locking pin 10 restricts the satellite boss 9 from moving in the positive Y direction.
  • the locking pin 10 has a "D" shape, and the part that contacts the inclined surface of the satellite boss 9 is a cylindrical surface, so that the locking pin 10 receives the contact elastic force given by the satellite boss 9 along the normal direction of the inclined surface.
  • the contact elastic force can be decomposed into The components of +Y and -X, the component of +Y makes the locking pin 10 press the base 13, and the component of -X makes the locking pin 10 press the locking slider 11, the base 13 and
  • the frame 1 is fixedly connected, and the locking slider 11 is installed in the guide groove of the base 13 and can only move in the Y direction, so the locking pin 10 is under the action of the satellite boss 9, the locking slider 11, and the base 13.
  • the satellite boss 9 is stationary under the action of the locking pin 10 and the base 13, so as to realize the function of locking the satellite.
  • the release function of the lock release mechanism refers to the release of the locking pin 10 from blocking the movement of the satellite boss 9.
  • the principle is:
  • the locking release unit includes a locking pin 10, a locking slider 11, an unlocking spring 12 and a base 13;
  • the left side of the base 13 is provided with a guide groove along the Y direction, the locking slider 11 is arranged in the guide groove, so that the locking slider 11 can only move in the Y direction; the locking slider 11 and the base 13 are provided
  • the unlocking spring 12 provides a thrust in the positive Y direction to the locking slider 11; when in the locked state, the unlocking spring is in a compressed state.
  • limit bracket 17 is fixedly installed above the locking slider 11 to limit the distance of the locking slider 11 moving in the positive Y direction.
  • the lower part of the locking slider 11 is provided with a cavity along the X direction; the locking slider 11 is located behind the locking pin 10, in the locked state, the upper flat part of the locking slider 11 is pressed behind the locking pin 10 ; In the released state, the locking slider 11 moves in the positive Y direction under the action of the unlocking spring 12, so that the cavity of the locking slider 11 rises, and the locking pin 10 is in the negative X direction provided by the satellite boss 9 Under the action of the contact elastic force, it moves in the negative X direction and enters the cavity of the locking slider 11, so that the locking pin 10 releases the obstacle to the movement of the satellite unit;
  • the key to the locking process and the releasing process is the movement of the locking slider 11 in the +Y direction.
  • An unlocking spring 12 is installed between the locking slider 11 and the base 13. The unlocking spring 12 is in a compressed state and will push the locking slider 11 to move to +Y.
  • the upper surface of the locking slider 11 is in contact with the electromagnet moving core 15, and the electromagnet moving core 15 can only move in the X direction. Therefore, only the electromagnet moving core 15 moves to -X for a period of time. When contacting again, the locking slider 11 can move.
  • the electromagnet moving core 15 and the electromagnet 16 are matched with a shaft hole, and a return spring is installed between the electromagnet moving core 15 and the electromagnet 16, and the return spring is a compression spring, giving the electromagnet moving core 15 +X direction elastic force.
  • the electromagnet 16 is energized, and the electromagnetic force generated by the electromagnet 16 will attract the electromagnet moving core 15, so that the electromagnet moving core 15 will overcome the elastic force of the return spring and move towards -X, thereby releasing the lock Tighten the constraints of the slider 11.
  • the locking slider 11 moves to +Y under the action of the unlocking spring 12 to release the constraint on the locking pin 10.
  • the locking pin 10 moves along -X under the action of the satellite boss 9. When the satellite boss 9 completely pushes the locking pin 10, the separation action can be completed.
  • the limit bracket 17 will limit the movement stroke of the locking slider 11.
  • the lock release drive unit includes an electromagnet limit nut 14, an electromagnet moving core 15 and an electromagnet 16;
  • the electromagnet 16 is fixed to the frame 1; the electromagnet 16 has a built-in coil; the electromagnet moving core 15 is arranged along the X direction.
  • the built-in coil of the electromagnet 16 is not energized, the right end of the electromagnet moving core 15 is pressed against the locking slider 11
  • the moving core 15 of the electromagnet is attracted to move in the negative X direction, releasing the limit on the locking slider 11 to make the lock
  • the tightening slider 11 moves in the positive direction of Y under the action of the unlocking spring 12.
  • Preloading method During the rocket launch process, the entire mechanism faces a harsh mechanical environment, including vibration, shock, and overload. In order to prevent gaps in the mechanism during launch and eliminate assembly errors and deviations caused by vibration, a pre-tightening unit is designed.
  • the pre-tensioning unit includes a pre-tensioning member bracket 5 and a pre-tensioning member 6;
  • the pretensioner bracket 5 is fixedly connected to the frame 1; the pretensioner 6 is connected to the pretensioner support 5 by a pretension screw; and the pretensioner 6 is located under the satellite boss 9;
  • the tightening member 6 moves in the positive direction of Y until the upper surface of the pre-tensioning member 6 is in close contact with the lower surface of the satellite boss 9; at the same time, during the tightening of the pre-tightening screw, the pre-tensioning member 6 pushes the satellite boss 9 along the positive Y direction.
  • Directional movement makes the contact between the satellite boss 9 and the locking pin 10 elastically deform, eliminates the contact gap, and together with the ejector pin 3, provides an initial separation force after the mechanism is unlocked.
  • the present invention also provides an electromagnetic lock release method for separating stars and arrows, which includes the following steps:
  • Step 1 Before the rocket reaches the predetermined orbit, the satellite and the rocket are locked in the following ways:
  • Step 1.1 the built-in coil of the electromagnet 16 is energized, so that the right end of the electromagnet moving core 15 is pressed against the top of the locking slider 11, so that the locking slider 11 is located below the guide groove;
  • Step 1.2 When the locking slider 11 is located below the guide groove, the upper flat part of the locking slider 11 is pressed behind the locking pin 10, restricting the locking pin 10 from moving in the negative X direction, thereby making the locking pin The right end of 10 protrudes from the first cavity of the base 13;
  • Step 1.3 Therefore, the cylindrical surface of the locking pin 10 is pressed against the inclined surface of the satellite boss 9, thereby restricting the movement of the satellite boss 9 in the positive direction of Y; at the same time, the ejection pin 3 is provided to the satellite unit under the action of the ejection spring 2.
  • the ejection force in the positive direction of Y but the ejection force is less than the force exerted by the locking pin 10 to the satellite boss 9; through the joint action of the locking pin 10 and the ejection rod 3, the satellite unit is locked on the frame 1;
  • Step 2 After the rocket enters orbit, a separation instruction is given, so that the satellite can be released safely, which specifically includes the following steps:
  • Step 2.1 after the rocket enters orbit, give a separation instruction, so as to power off the built-in coil of the electromagnet 16;
  • Step 2.2 when the electromagnet moving core 15 releases the limiting effect on the locking slider 11 in the Y direction, the locking slider 11 moves in the positive direction of Y under the action of the unlocking spring 12, and is aligned by the limit bracket 17 The movement distance of the locking slider 11 is limited, so that the cavity of the locking slider 11 rises to a certain height;
  • Step 2.3 when the cavity of the locking slider 11 rises to a certain height, release the hindrance to the movement of the locking pin 10 in the negative X direction;
  • the locking pin 10 moves in the negative X direction under the contact elastic force provided by the inclined surface of the satellite boss 9 in the negative X direction, and partially enters the cavity of the locking slider 11, so that the locking pin 10 Break away from the limit of the satellite boss 9 in the positive direction of Y;
  • Step 2.4 when the locking pin 10 is out of the limit on the satellite boss 9 in the positive direction of Y, under the ejection force of the ejection spring 2, the ejector pin 3 is moved in the positive direction of Y so as to include the satellite boss
  • the satellite unit of 9 is ejected and launched in the positive direction of Y to realize the safe release of the satellite.
  • Frame 1 Aluminum alloy is generally used, and conductive oxidation treatment must be carried out;
  • Ejection ejector rod 3 the same material as the spring sleeve 4 and satellite boss 9 cannot be used, and the hardness should be greater than 1;
  • Pre-tensioning member bracket 5 Use stronger materials, such as stainless steel, titanium alloy, and not the same material as the satellite boss 9;
  • Satellite boss 9 The surface needs to be polished
  • Locking pin 10 The hardness is greater than that of all other parts, the surface must be polished, and it can be plated with molybdenum disulfide. It cannot be the same material as the satellite boss 9 and base 13;
  • Locking slider 11 It cannot be the same material as the locking pin 10 and base 13;
  • Electromagnet moving core 15 It must be made of iron material, and the force-bearing part is stainless steel. The gap between the electromagnet 16 and the electromagnet 16 must be strictly controlled within ⁇ 0.3mm of the design value during assembly.
  • Electromagnet 16 Internal coil parameters need to be calculated according to the load
  • Figure 3 is an embodiment of this mechanism.
  • 1 is a rack connected and fixedly installed with the rocket
  • 18 is the lock release mechanism of the present invention
  • 19 is a spherical satellite.
  • the spherical satellite is connected with the frame 1 through two sets of symmetrical locking and releasing mechanisms, so as to realize the star-arrow connection and the star-arrow separation.
  • the two sets of lock release mechanisms act at the same time, and the satellite and the rocket are separated.
  • the mass of the satellite is 22 kg
  • the diameter of the sphere is 550 mm
  • the total weight of the frame and the lock release mechanism is 15 kg.
  • the inclined plane angle ⁇ of the satellite boss 9 is 30°
  • the selected ejection spring 2 has a force of 200N
  • the unlocking spring 12 has a force of 100N
  • the electromagnet return spring has a force of 40N.
  • the selected electromagnet moving core 15 has a power of 100W
  • the selected frame 1 is 2A12 aluminum alloy
  • the ejector pin 3 is 1Cr18 stainless steel
  • the spring sleeve 4 is 2A12 aluminum alloy
  • the pretensioner bracket 5 and the pretensioner 6 are 1Cr18 Stainless steel
  • positioning pin 8 and satellite boss 9 are TC4
  • locking pin 10 is 9Cr18 stainless steel (quenched)
  • locking slider 11 is TC4
  • base 13 is 1Cr18 stainless steel
  • limit bracket 17 is 2A12 aluminum alloy.
  • the diameter of the locking pin 10 is 16 mm
  • the movement stroke of the locking slider 11 is 2 mm
  • the movement stroke of the electromagnet 16 is 4 mm.
  • the separation speed of the satellite in this embodiment is 0.8 m/s
  • the mechanical environment adaptability is shown in Table 1 to Table 3
  • the separation action time is about 40 ms, which enables repeated ground tests.
  • Step 1 Place the bottom plate as an assembly reference.
  • Figure 4 is a schematic diagram of the bottom plate.
  • Step 2 Install the support arm on the bottom plate.
  • Each support arm is fixed by 4 M8 ⁇ 16 hexagon socket titanium screws, and the screws are pre-tightened according to the standard torque.
  • Figure 5 which is a schematic diagram of the support arm.
  • Figure 6 which is an assembly drawing of the support arm and the bottom plate.
  • Step 3 Put the 4 ejection springs into the round holes of the 4 support arms respectively. Place the spring top rod concentrically above the spring to form Figure 7.
  • Step 4 Fasten the spring sleeve on the top of the spring mandrel, and fix each spring sleeve with 4 M4 ⁇ 10 inner hexagonal titanium screws. Pay attention to the direction of the spring sleeve to form Figure 8.
  • Step 5 Install 4 pretensioner brackets on the 4 support arms, and each part is fixed by 4 M4 ⁇ 10 hexagon socket titanium screws.
  • the theoretical requirement of a single screw pre-tightening force is 2475N, which is converted into a torque wrench with a torque of about 1.88Nm.
  • the installation parallelism is required to reach 0.1mm, forming Figure 9.
  • Step 6 Place the pretensioner.
  • Each pretensioner is respectively connected to a pretensioner bracket by two M6 ⁇ 12 hexagon socket titanium alloy screws.
  • a 0.6mm thick washer is pasted on the upper surface of the pre-tensioning part of the first and third quadrants to form Figure 12. Refer to Figure 10 for the placement of the pretensioner tooling; refer to Figure 11 for the placement of the washer.
  • Step 7 Install the limit base of the travel switch and the electromagnetic lock release base.
  • the installation position of the travel switch base is the 1, 3 quadrants
  • the electromagnetic lock release base is the 2, 4 quadrants.
  • Each base is connected to the support arm by 4 horizontal and 4 vertical M4 ⁇ 10 hexagon socket titanium screws. Note that the fixing screw of the electromagnetic lock release base needs to be tightened and installed according to the standard torque; the screw of the limit switch base does not need to be tightened, and the subsequent process needs to be adjusted, forming Figure 13.
  • Step 8 Install two electromagnets, and pay attention to routing the wires through the holes of the support arm.
  • Each electro-permanent magnet is connected to the base by 4 M4 ⁇ 10 hexagon socket titanium screws. Note that the surface of the electromagnet is clean. When installing, the electromagnet should be as close as possible to the inside of the separation mechanism, as shown in Figure 14.
  • Step 9 Install the spring compression tooling.
  • Each mounting block is connected to the support arm with M4 ⁇ 16 hexagon socket screws, forming Figure 15.
  • Step 10 Compress the ejection springs and operate them one by one. First press down the spring with a tool, and then insert the spring pressure bar from the spring sleeve. Both ends of each spring pressure bar pass through the M5 ⁇ 30 hexagon socket screws and screw them into the spring mounting block installed in the previous step, and screw in the screws so that the highest point of the spring mandrel is lower than the upper surface of the pretensioner, forming Figure 16.
  • Step 11 Place the backup center bearing ring first. If there is no problem, remove the center bearing ring and place the satellite.
  • the bottom travel switch of the satellite is in contact with the bottom pressing surface of the separation mechanism.
  • the satellite's panoramic camera and the ground hole on the bottom plate of the separation mechanism are in the opposite quadrants, forming Figure 17.
  • Step 12 Place the locking pin. Apply grease before installation. Note that the plane of the locking pin fits the plane of the base, and the pin can only be put in from one side of the base, forming Figure 18.
  • Step 13 Install the locking sliders on both sides. Grease the locking slider before installation. First press the moving core of the electromagnet to make room for the locking slider, and then insert the unlocking spring into the middle of the locking slider and the locking release base (each has a positioning rod, the spring is not shown in the figure), and then press the lock Hold the slider tightly, rotate the moving core of the electromagnet so that the moving core enters the waist hole of the locking slider, confirm that the limit nut of the electromagnet fits the electromagnet shell, and then complete the installation of this step. In this step, the distance between the moving core baffle and the electromagnet housing needs to be measured, which is required to be 3.6 to 3.8 mm.
  • This size can be controlled by adding or subtracting a 0.1mm spacer between the limit nut at the end of the moving core and the electromagnet housing.
  • the measurement position is the symmetrical diameter of the electromagnet disc in the horizontal direction.
  • the two measured values are averaged to form Figure 19.
  • Figure 20 is the spacing control diagram.
  • Step 14 Install 2 KX-1 travel switches, and pay attention to routing the wires through the holes of the support arm.
  • Each limit switch is connected with the limit base of the limit switch through two M4 ⁇ 10 hexagonal titanium screws. After tightening the screws between the limit switch and the limit switch base, adjust the position of the limit switch base, test the limit switch signal, confirm that the compression amount of the limit switch is correct, and then tighten the 8 M4 screws connecting the limit switch base and the support arm to form Figure 21.
  • Step 15 Adjust the satellite and the locking pin, and apply pre-tightening force. Take out the 0.5-thick tooling between the pretensioner on both sides of the lock release base and the pretensioner bracket, and tighten the 8 M6 screws of the 4 pretensioners. The symmetrical screws need to be synchronized to apply the pretension force as much as possible. During the tightening process, the positions of the locking pin and the satellite need to be adjusted. The locking pin should be placed in the center as far as possible. The unilateral distance between the side of the pin and the inner wall of the base groove should be controlled to 6 ⁇ 0.5mm, and the distance between the outermost side of the satellite and the innermost surface of the base should be controlled to 0.9 ⁇ 0.2mm, forming Figure 22.
  • Figure 23 is a schematic diagram of the distance between the outermost side of the satellite and the innermost side of the base opposite to the surface of 0.9mm.
  • Fig. 24 is a schematic diagram when the distance between the side surface of the pin and the inner wall of the base groove is 6 mm.
  • Step 16 Install 4 stiffeners, and each stiffener is connected to two adjacent support arms.
  • Step 17 Remove the spring compression tooling installed in steps 9 and 10, make the spring ejector pin against the satellite, and check whether the ejector pin is tilted. At this time, the ground resistance between the satellite and the separation mechanism is measured.
  • Step 18 Apply pre-tightening force to the pre-tightening parts of the first and third quadrants, and tighten the M6 screws.
  • Step 19 Install 2 limit brackets, each bracket is connected to the support arm by 4 M4 ⁇ 10 inner hexagonal titanium screws.
  • Step 20 The assembly is completed, and Figure 3 is formed.
  • the invention provides an electromagnetic lock release mechanism and an electromagnetic lock release method for separating stars and arrows, which has the following advantages:
  • test items are overload, sinusoidal vibration, random vibration, shock, etc.
  • the satellite and the rocket can be reliably separated, and the separation speed can be effectively controlled by adjusting the design parameters of the spring.
  • the separation speed can be measured by shooting with a high-speed camera.
  • the separation test can be repeated on the ground without causing damage to itself, that is, the original components can be used to rebuild the assembly after the separation is completed.
  • the separation action is a continuous mechanism action, which has no impact on the satellite, and the impact response of the satellite can be tested through the sensor.

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Abstract

一种用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法，电磁锁紧释放机构包括：机架(1)、弹射单元、卫星单元、锁紧释放单元以及锁紧释放驱动单元；弹射单元包括弹射弹簧(2)、弹射顶杆(3)和弹簧套筒(4)；锁紧释放单元包括锁紧销(10)、锁紧滑块(11)、解锁弹簧(12)和基座(13)；锁紧释放驱动单元包括电磁铁限位螺母(14)、电磁铁动芯(15)和电磁铁(16)。用于星箭分离的电磁锁紧释放机构是一种可重复试验的微小卫星与火箭分离使用的点式定位锁紧释放机构，能够在复杂的力学环境下，实现卫星和火箭的可靠锁紧和可靠分离；能够在地面重复进行分离试验；分离完全为机构动作，无污染；分离动作为连续机构动作，对卫星无冲击；降低了卫星研制的成本。

Description

用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法 技术领域
本发明属于航天技术领域，具体涉及一种用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法。
背景技术
卫星锁紧释放机构主要应用于星箭分离机构，其主要工作原理为：火箭在到达预定轨道之前，实现卫星与火箭的可靠连接；在火箭入轨后，给出分离指令，从而使卫星安全释放，即：使卫星以设计的速度与火箭分离。
卫星锁紧释放机构的关键设计要素是机构形式和主动器件。现有技术中，机构形式主要采用对接环、包带式的分离机构，存在的主要问题为：对接环、包带式的分离机构要求卫星必须有一个对接面与分离机构连接安装，无法应用于不规则形状的需要点式定位锁紧的卫星。主动器件主要采用火攻品、热刀、记忆合金、电机等，存在的主要问题为：火工品成本昂贵，管理要求复杂，而且是破坏性的分离，无法在地面进行重复试验；热刀能够承受的载荷很小，难以用于整个卫星的固定，而且动作时间不确定，不能保证同步性；记忆合金动作时间不确定，不能保证同步性；电机需要持续供电，微小卫星一般是搭载火箭发射，火箭一般不能提供持续的电流，因此难以应用在微小卫星上。
因此，现有技术中急需一种能够灵活应用于各类构型的微小卫星的可重复试验的点式定位锁紧释放机构。
发明内容
针对现有技术存在的缺陷，本发明提供一种用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法，可有效解决上述问题。
本发明采用的技术方案如下：
本发明提供一种用于星箭分离的电磁锁紧释放机构，包括：机架(1)、弹射单元、卫星单元、锁紧释放单元以及锁紧释放驱动单元；设置水平方向为X方向；垂直方向为Y方向；
所述卫星单元包括卫星(7)、定位销(8)和卫星凸台(9)；所述卫星(7) 通过所述定位销(8)与所述卫星凸台(9)固定连接；
所述弹射单元设置于所述卫星单元的下方，用于向所述卫星单元提供Y正方向的弹射力；所述弹射单元包括弹射弹簧(2)、弹射顶杆(3)和弹簧套筒(4)；所述弹簧套筒(4)与所述机架(1)固定连接；所述弹簧套筒(4)内部设置所述弹射弹簧(2)；所述弹射顶杆(3)的底部位于所述弹簧套筒(4)内，并压于所述弹射弹簧(2)的上面；所述弹射顶杆(3)的顶部穿过所述弹簧套筒(4)，而延伸到所述卫星单元的下面，并与所述卫星单元的下表面接触；在锁紧状态下，所述弹射弹簧(2)被所述弹射顶杆(3)压紧，所述弹射弹簧(2)通过所述弹射顶杆(3)向所述卫星单元提供Y正方向的弹射力；在释放状态下，在所述弹射弹簧(2)的弹射力作用下，所述卫星单元向Y正方向运动，从而与所述机架(1)分离；
所述锁紧释放单元包括锁紧销(10)、锁紧滑块(11)、解锁弹簧(12)和基座(13)；
所述基座(13)位于所述卫星单元的左侧，并与所述机架(1)固定连接；所述基座(13)的右侧开设第一空腔，所述锁紧销(10)为D形状；所述卫星凸台(9)的上表面设计为斜面；在锁紧状态下，所述锁紧销(10)部分位于所述第一空腔内，所述锁紧销(10)的圆柱面压于所述卫星凸台(9)的斜面，进而限制所述卫星凸台(9)向Y正方向运动；此外，所述卫星凸台(9)的斜面向所述锁紧销(10)提供沿斜面法向的接触弹力；所述接触弹力分解为Y正方向接触弹力和X负方向接触弹力；Y正方向接触弹力使所述锁紧销(10)压紧所述基座(13)；X负方向接触弹力使所述锁紧销(10)压紧位于其后面的锁紧滑块(11)；
所述基座(13)的左侧开设沿Y方向的导向槽，所述锁紧滑块(11)设置于所述导向槽内，使所述锁紧滑块(11)仅能进行Y方向运动；所述锁紧滑块(11)和所述基座(13)之间设置所述解锁弹簧(12)，所述解锁弹簧(12)向所述锁紧滑块(11)提供Y正方向的推力；
所述锁紧滑块(11)的下部开设沿X方向的空腔；所述锁紧滑块(11)位于所述锁紧销(10)的后面，在锁紧状态下，所述锁紧滑块(11)上部平面部分压于所述锁紧销(10)的后面；在释放状态下，所述锁紧滑块(11)在所述解 锁弹簧(12)的作用下向Y正方向运动，从而使所述锁紧滑块(11)的空腔上升，所述锁紧销(10)在所述卫星凸台(9)提供的X负方向接触弹力的作用下，沿X负方向运动，进入所述锁紧滑块(11)的空腔内，从而使所述锁紧销(10)释放对所述卫星单元的运动阻碍；
所述锁紧释放驱动单元包括电磁铁限位螺母(14)、电磁铁动芯(15)和电磁铁(16)；所述电磁铁(16)与所述机架(1)固定；所述电磁铁(16)内置线圈；所述电磁铁动芯(15)沿X方向设置，所述电磁铁(16)的内置线圈在没有通电时，所述电磁铁动芯(15)的右端压于所述锁紧滑块(11)的顶端，限制所述锁紧滑块(11)沿Y正方向运动；当所述电磁铁(16)的内置线圈通电时，所述电磁铁动芯(15)向X负方向运动，释放对所述锁紧滑块(11)的限位，使所述锁紧滑块(11)在解锁弹簧(12)的作用下向Y正方向运动。
优选的，还包括限位支架(17)；所述限位支架(17)固定安装于所述锁紧滑块(11)的上方，用于限制所述锁紧滑块(11)向Y正方向运动的距离。
优选的，还包括预紧单元；
所述预紧单元包括预紧件支架(5)和预紧件(6)；
所述预紧件支架(5)与所述机架(1)固定连接；所述预紧件(6)通过预紧螺钉与所述预紧件支架(5)连接；并且，所述预紧件(6)位于所述卫星凸台(9)的下面；通过旋转所述预紧螺钉，使所述预紧件(6)沿Y正方向运动，直到所述预紧件(6)的上表面与所述卫星凸台(9)的下表面紧密接触；同时，在所述预紧螺钉拧紧过程中，所述预紧件(6)推动所述卫星凸台(9)沿Y正方向运动，使所述卫星凸台(9)和所述锁紧销(10)之间的接触产生弹性形变，消除接触间隙。
本发明提供一种用于星箭分离的电磁锁紧释放机构的电磁锁紧释放方法，包括以下步骤：
步骤1，火箭在到达预定轨道之前，卫星与火箭通过以下方式实现锁紧：
步骤1.1，电磁铁(16)的内置线圈通电，从而使电磁铁动芯(15)的右端压于锁紧滑块(11)的顶端，使锁紧滑块(11)位于导向槽的下方；
步骤1.2，当锁紧滑块(11)位于导向槽的下方时，锁紧滑块(11)的上部 平面部分压于锁紧销(10)的后面，限制锁紧销(10)沿X负方向运动，从而使锁紧销(10)的右端凸出于基座(13)的第一空腔；
步骤1.3，因此，锁紧销(10)的圆柱面压于卫星凸台(9)的斜面，进而限制卫星凸台(9)向Y正方向运动；同时，弹射顶杆(3)在弹射弹簧(2)的作用下向卫星单元提供Y正方向的弹射力，但该弹射力小于锁紧销(10)向卫星凸台(9)施加的力；通过锁紧销(10)和弹射顶杆(3)的共同作用，使卫星单元锁紧在机架(1)上面；
步骤2，在火箭入轨后，给出分离指令，从而使卫星安全释放，具体包括以下步骤：
步骤2.1，在火箭入轨后，给出分离指令，从而使电磁铁(16)的内置线圈断电；
当电磁铁(16)的内置线圈断电时，电磁铁动芯(15)沿X负方向运动，从而使电磁铁动芯(15)的右端离开锁紧滑块(11)的顶端，释放对锁紧滑块(11)在Y方向的限位作用；
步骤2.2，当电磁铁动芯(15)释放对锁紧滑块(11)在Y方向的限位作用时，锁紧滑块(11)在解锁弹簧(12)的作用下向Y正方向运动，并通过限位支架(17)对锁紧滑块(11)的运动距离进行限制，从而使锁紧滑块(11)的空腔上升一定高度；
步骤2.3，当锁紧滑块(11)的空腔上升一定高度时，释放对锁紧销(10)在X负方向运动的阻碍；
因此，锁紧销(10)在卫星凸台(9)的斜面提供的X负方向接触弹力的作用下，沿X负方向运动，并部分进入到锁紧滑块(11)的空腔中，从而使锁紧销(10)脱离对卫星凸台(9)沿Y正方向的限位；
步骤2.4，当锁紧销(10)脱离对卫星凸台(9)沿Y正方向的限位时，在弹射弹簧(2)的弹射力作用下，使弹射顶杆(3)沿Y正方向运动，从而使包含卫星凸台(9)的卫星单元沿Y正方向弹射发射，实现卫星安全释放。
本发明提供的用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法具有以下优点：
本发明是一种可重复试验的微小卫星与火箭分离使用的点式定位锁紧释放机构，能够在复杂的力学环境下，实现卫星和火箭的可靠锁紧和可靠分离；能够在地面重复进行分离试验；分离完全为机构动作，无污染；分离动作为连续机构动作，对卫星无冲击；降低了卫星研制的成本。
附图说明
图1为本发明提供的用于星箭分离的电磁锁紧释放机构在锁紧状态下的剖面图；
图2为本发明提供的用于星箭分离的电磁锁紧释放机构在释放状态下的剖面图；
图3为本发明提供的用于星箭分离的电磁锁紧释放机构的装配图；
图4为底板示意图；
图5为支撑臂示意图；
图6为支撑臂和底板的装配图；
图7为步骤3装配后形成的结构图；
图8为步骤4装配后形成的结构图；
图9为步骤5装配后形成的结构图；
图10为预紧件工装放置位置图；
图11为垫圈放置位置图；
图12为步骤6装配后形成的结构图；
图13为步骤7装配后形成的结构图；
图14为步骤8装配后形成的结构图；
图15为步骤9装配后形成的结构图；
图16为步骤10装配后形成的结构图；
图17为步骤11装配后形成的结构图；
图18为步骤12装配后形成的结构图；
图19为步骤13装配后形成的结构图；
图20为间距控制图；
图21为步骤14装配后形成的结构图；
图22为步骤15装配后形成的结构图；
图23为卫星最外侧与基座最内侧相对面距离0.9mm的示意图；
图24为销侧面与基座槽内壁单侧距离6mm时的示意图。
具体实施方式
为了使本发明所解决的技术问题、技术方案及有益效果更加清楚明白，以下结合附图及实施例，对本发明进行进一步详细说明。应当理解，此处所描述的具体实施例仅用以解释本发明，并不用于限定本发明。
本发明提供一种用于星箭分离的电磁锁紧释放机构，用于卫星与火箭的锁紧释放，是一种可重复试验的微小卫星与火箭分离使用的点式定位锁紧释放机构，即：在一般状态下为锁紧状态，给出指令后变成分离释放状态。
参考图1和图2，包括机架1、弹射单元、卫星单元、锁紧释放单元以及锁紧释放驱动单元；在图1中，设置水平方向为X方向，水平向右方向为X正方向；垂直方向为Y方向，垂直向上方向为Y正方向。下面对各单元分别详细介绍：
(1)机架
机架，静止固定部件，与火箭固定连接，起到支撑作用，随着卫星7形状的变化而变化。
(2)卫星单元
卫星单元包括卫星7、定位销8和卫星凸台9；卫星7通过定位销8与卫星凸台9固定连接，通过卫星凸台9实现与机构的其余部分相互作用，即：使用本机构需要在卫星7上找到合适位置安装定位销8和卫星凸台9。一般需要在单个卫星上以对称方式安装多个定位销8和卫星凸台9的组合。
因此，卫星7、定位销8和卫星凸台9整体连接形成一个整体，当进行星箭分离时，卫星单元整体被弹射发射。其中，卫星7为运动部件，本锁紧释放机构的作用对象，本锁紧释放机构并不限定卫星的具体形式。
(3)弹射单元
弹射单元设置于卫星单元的下方，用于向卫星单元提供Y正方向的弹射力；弹射单元包括弹射弹簧2、弹射顶杆3和弹簧套筒4；弹簧套筒4与机架1固定连接；弹簧套筒4内部设置弹射弹簧2；弹射顶杆3的底部位于弹簧套筒4内，并压于弹 射弹簧2的上面；弹射顶杆3的顶部穿过弹簧套筒4，而延伸到卫星单元的下面，并与卫星单元的下表面接触，具体是与卫星凸台9的下表面接触；
在锁紧状态下，弹射弹簧2被弹射顶杆3压紧，弹射弹簧为压缩状态，压缩后存储的弹性势能即为卫星分离的能量来源，因此，弹射弹簧2通过弹射顶杆3向卫星单元提供Y正方向的弹射力，但在锁紧状态时，锁紧释放单元的锁紧功能会向卫星单元提供Y负方向的作用力，从而克服弹射弹簧2的弹射力，使卫星单元与机架1锁紧；在释放状态下，锁紧释放单元释放对卫星单元的作用力，因此，在弹射弹簧2的弹射力作用下，卫星单元向Y正方向运动，从而与机架1分离。
(4)锁紧释放单元
锁紧释放单元主要包括锁紧功能和释放功能。
锁紧功能是指：阻止卫星单元在弹射弹簧2和弹射顶杆3的作用下与机架1分离。大致原理为：
为实现锁紧功能，卫星凸台9的上表面设计为斜面，斜面角度α为设计值。卫星凸台9的斜面与锁紧销10接触，锁紧销10限制卫星凸台9向Y正方向运动。锁紧销10为“D”形状，与卫星凸台9的斜面接触的部分为圆柱面，从而锁紧销10受到卫星凸台9给予的沿斜面法向的接触弹力，该接触弹力可以分解为+Y和-X两个方向的分力，+Y的分力使得锁紧销10压紧基座13，-X的分力使得锁紧销10压紧锁紧滑块11，基座13与机架1固连，锁紧滑块11安装在基座13的导向槽内只能沿Y方向运动，因此锁紧销10在卫星凸台9、锁紧滑块11、基座13的作用下平衡静止，卫星凸台9在锁紧销10、基座13的作用下静止，从而实现对卫星的锁紧功能。
锁紧释放机构的释放功能，是指解除锁紧销10对卫星凸台9运动的阻止，原理：
当锁紧滑块11沿+Y运动时，锁紧滑块11的下方的空腔会上升，锁紧销10失去与锁紧滑块11的接触。此时锁紧销10在卫星凸台9给予的-X方向的接触力作用下会沿-X方向运动，进入锁紧滑块11的空腔内，卫星凸台9不再存在运动的阻碍，从而在弹射弹簧2、弹射顶杆3的作用下向+Y方向运动，实现分离。
对锁紧释放运动动作控制的方法：
具体结构如下：锁紧释放单元包括锁紧销10、锁紧滑块11、解锁弹簧12和 基座13；
基座13位于卫星单元的左侧，并与机架1固定连接；基座13的右侧开设第一空腔，锁紧销10为D形状；卫星凸台9的上表面设计为斜面；在锁紧状态下，锁紧销10部分位于第一空腔内，锁紧销10的圆柱面压于卫星凸台9的斜面，进而限制卫星凸台9向Y正方向运动；此外，卫星凸台9的斜面向锁紧销10提供沿斜面法向的接触弹力；接触弹力分解为Y正方向接触弹力和X负方向接触弹力；Y正方向接触弹力使锁紧销10压紧基座13；X负方向接触弹力使锁紧销10压紧位于其后面的锁紧滑块11；
基座13的左侧开设沿Y方向的导向槽，锁紧滑块11设置于导向槽内，使锁紧滑块11仅能进行Y方向运动；锁紧滑块11和基座13之间设置解锁弹簧12，解锁弹簧12向锁紧滑块11提供Y正方向的推力；在锁紧状态时，解锁弹簧为压缩状态。
还包括限位支架17；限位支架17固定安装于锁紧滑块11的上方，用于限制锁紧滑块11向Y正方向运动的距离。
锁紧滑块11的下部开设沿X方向的空腔；锁紧滑块11位于锁紧销10的后面，在锁紧状态下，锁紧滑块11上部平面部分压于锁紧销10的后面；在释放状态下，锁紧滑块11在解锁弹簧12的作用下向Y正方向运动，从而使锁紧滑块11的空腔上升，锁紧销10在卫星凸台9提供的X负方向接触弹力的作用下，沿X负方向运动，进入锁紧滑块11的空腔内，从而使锁紧销10释放对卫星单元的运动阻碍；
(5)锁紧释放驱动单元
锁紧过程和释放过程的关键是锁紧滑块11沿+Y方向的运动。锁紧滑块11与基座13之间安装有解锁弹簧12，解锁弹簧12为压紧状态，会推动锁紧滑块11向+Y运动。但锁紧滑块11的上表面与电磁铁动芯15接触，电磁铁动芯15只能沿X方向运动，因此仅电磁铁动芯15向-X运动一段行程后与锁紧滑块11不再接触时，锁紧滑块11才能运动。电磁铁动芯15与电磁铁16为轴孔配合，且电磁铁动芯15与电磁铁16之间安装有复位弹簧，复位弹簧为压缩弹簧，给予电磁铁动芯15+X方向的弹力。当给出分离信号后，对电磁铁16进行通电，电磁铁16产生的电磁力会吸附电磁铁动芯15，从而使得电磁铁动芯15克服复位弹簧的弹力向-X运动，从而解除对锁紧滑块11的约束。锁紧滑块11在解锁弹簧12的作用下向+Y运动， 解除对锁紧销10的约束。锁紧销10在卫星凸台9的作用下沿-X运动，当卫星凸台9完全推开锁紧销10后，即可以完成分离动作。为了确保锁紧滑块11运动后锁紧销10一定能够进入锁紧滑块11的空腔内，限位支架17会限制锁紧滑块11的运动行程。
具体结构如下：
锁紧释放驱动单元包括电磁铁限位螺母14、电磁铁动芯15和电磁铁16；
电磁铁16与机架1固定；电磁铁16内置线圈；电磁铁动芯15沿X方向设置，电磁铁16的内置线圈在没有通电时，电磁铁动芯15的右端压于锁紧滑块11的顶端，限制锁紧滑块11沿Y正方向运动；当电磁铁16的内置线圈通电时，吸引电磁铁动芯15向X负方向运动，释放对锁紧滑块11的限位，使锁紧滑块11在解锁弹簧12的作用下向Y正方向运动。
(6)预紧单元
预紧方法：在火箭发射过程中，整个机构面临着严苛的力学环境，包括振动、冲击、过载等。为了防止在发射中机构出现间隙，消除装配误差、振动带来的偏差，设计预紧单元。
预紧单元包括预紧件支架5和预紧件6；
预紧件支架5与机架1固定连接；预紧件6通过预紧螺钉与预紧件支架5连接；并且，预紧件6位于卫星凸台9的下面；通过旋转预紧螺钉，使预紧件6沿Y正方向运动，直到预紧件6的上表面与卫星凸台9的下表面紧密接触；同时，在预紧螺钉拧紧过程中，预紧件6推动卫星凸台9沿Y正方向运动，使卫星凸台9和锁紧销10之间的接触产生弹性形变，消除接触间隙，并与弹射顶杆3一同在机构解锁后提供初始的分离力。
本发明还提供一种用于星箭分离的电磁锁紧释放方法，包括以下步骤：
步骤1，火箭在到达预定轨道之前，卫星与火箭通过以下方式实现锁紧：
步骤1.1，电磁铁16的内置线圈通电，从而使电磁铁动芯15的右端压于锁紧滑块11的顶端，使锁紧滑块11位于导向槽的下方；
步骤1.2，当锁紧滑块11位于导向槽的下方时，锁紧滑块11的上部平面部分压于锁紧销10的后面，限制锁紧销10沿X负方向运动，从而使锁紧销10的右端凸 出于基座13的第一空腔；
步骤1.3，因此，锁紧销10的圆柱面压于卫星凸台9的斜面，进而限制卫星凸台9向Y正方向运动；同时，弹射顶杆3在弹射弹簧2的作用下向卫星单元提供Y正方向的弹射力，但该弹射力小于锁紧销10向卫星凸台9施加的力；通过锁紧销10和弹射顶杆3的共同作用，使卫星单元锁紧在机架1上面；
步骤2，在火箭入轨后，给出分离指令，从而使卫星安全释放，具体包括以下步骤：
步骤2.1，在火箭入轨后，给出分离指令，从而使电磁铁16的内置线圈断电；
当电磁铁16的内置线圈断电时，电磁铁动芯15沿X负方向运动，从而使电磁铁动芯15的右端离开锁紧滑块11的顶端，释放对锁紧滑块11在Y方向的限位作用；
步骤2.2，当电磁铁动芯15释放对锁紧滑块11在Y方向的限位作用时，锁紧滑块11在解锁弹簧12的作用下向Y正方向运动，并通过限位支架17对锁紧滑块11的运动距离进行限制，从而使锁紧滑块11的空腔上升一定高度；
步骤2.3，当锁紧滑块11的空腔上升一定高度时，释放对锁紧销10在X负方向运动的阻碍；
因此，锁紧销10在卫星凸台9的斜面提供的X负方向接触弹力的作用下，沿X负方向运动，并部分进入到锁紧滑块11的空腔中，从而使锁紧销10脱离对卫星凸台9沿Y正方向的限位；
步骤2.4，当锁紧销10脱离对卫星凸台9沿Y正方向的限位时，在弹射弹簧2的弹射力作用下，使弹射顶杆3沿Y正方向运动，从而使包含卫星凸台9的卫星单元沿Y正方向弹射发射，实现卫星安全释放。
对各个工艺要求：
本机构对于各部件材料均提出严格要求。一方面，存在相互运动的材料一般有硬度要求，同时由于航天真空环境，一般相互运动的材料不宜采用同种金属防止冷焊。
机架1：一般采用铝合金，必须进行导电氧化处理；
弹射顶杆3：与弹簧套筒4、卫星凸台9不能采用同种材料，硬度要大于1；
预紧件支架5、预紧件6：选用强度较大的材料，如不锈钢、钛合金，不能与卫星凸台9采用同种材料；
卫星凸台9：表面需抛光处理；
锁紧销10：硬度要大于其他所有部件，表面必须抛光处理，可镀二硫化钼，与卫星凸台9、基座13不能是同种材料；
锁紧滑块11：与锁紧销10、基座13不能是同种材料；
电磁铁动芯15：必须为铁质材料，承力部位为不锈钢，装配时与电磁铁16的间隙必须严格控制在设计值±0.3mm内。
电磁铁16：内部线圈参数需根据载荷进行核算；
图3即为本机构的一种实施例。图中，1为与火箭连接固定安装的机架，18为本发明的锁紧释放机构，19为球形卫星。球形卫星通过2组对称的锁紧释放机构与机架1连接，从而实现星箭连接和星箭分离。当指令给出后，2组锁紧释放机构同时动作，卫星与火箭分离。
本实施例中，卫星质量为22kg，球体直径为550mm，机架和锁紧释放机构全重为15kg。
关键的设计参数：
卫星凸台9的斜面角度α为30°，选用的弹射弹簧2作用力为200N，解锁弹簧12作用力为100N，电磁铁复位弹簧作用力为40N。选用的电磁铁动芯15功率为100W，选用的机架1为2A12铝合金，弹射顶杆3为1Cr18不锈钢，弹簧套筒4为2A12铝合金，预紧件支架5、预紧件6为1Cr18不锈钢，定位销8、卫星凸台9为TC4，锁紧销10为9Cr18不锈钢(淬火)，锁紧滑块11为TC4，基座13为1Cr18不锈钢，限位支架17为2A12铝合金。锁紧销10直径为16mm，锁紧滑块11的动作行程为2mm，电磁铁16的动作行程为4mm。
本实施例的卫星分离速度为0.8m/s，力学环境适应能力见表1-表3，分离动作时间约为40ms，能够反复进行地面试验。
表1低频正弦扫描试验量级
Figure PCTCN2020136506-appb-000001
Figure PCTCN2020136506-appb-000002
表2冲击环境条件
频率范围(Hz) 加速度冲击响应谱(Q＝10)
100-1500 +9dB/oct
1500-5000 2800g
表3随机振动试验量级
Figure PCTCN2020136506-appb-000003
装配步骤：
步骤1、放置底板，作为装配基准。参考图4，为底板示意图。
步骤2、将支撑臂安装在底板上，每个支撑臂通过4个M8×16内六角钛螺钉固定，按照标准力矩预紧螺钉。参考图5，为支撑臂示意图。参考图6，为支撑臂和底板的装配图。
步骤3、将4只弹射弹簧分别放入4只支撑臂的圆孔内。在弹簧上方同心放置弹簧顶杆，形成图7。
步骤4、将弹簧套筒扣在弹簧顶杆上方，每个弹簧套筒用4个M4×10内六角钛螺钉拧紧固定，注意弹簧套筒的方向，形成图8。
步骤5、在4个支撑臂上安装4个预紧件支架，每个零件通过4只M4×10内六角钛螺钉固定。单个螺钉预紧力理论需求为2475N，转换为力矩扳手力矩约为1.88Nm。安装时注意和上表面和支撑臂上表面的平行度。要求安装平行度达到0.1mm，形成图9。
步骤6、安放预紧件。每个预紧件通过2只M6×12内六角钛合金螺钉分别连接一个预紧件支架。安装中，使用零点五厚工装保持2处螺钉连接处相对面距离均为0.5mm，此处螺钉后续需要调整，本工序不需要施加预紧力矩。第1、3象限 的预紧件上表面粘贴0.6mm厚垫圈，形成图12。参考图10，为预紧件工装放置位置图；参考图11，为垫圈放置位置图。
步骤7、安装行程开关限位基座和电磁锁紧释放基座，注意行程开关基座安装位置为1、3象限，电磁锁紧释放基座为2、4象限。每个基座通过4只水平、4只竖直的M4×10内六角钛螺钉与支撑臂连接。注意，电磁锁紧释放基座的固定螺钉需要拧紧，按照标准力矩安装；行程开关限位基座的螺钉不需要拧紧，后续工序需要调整，形成图13。
步骤8、安装2个电磁铁，注意从支撑臂的孔中走线。每个电永磁体通过4只M4×10内六角钛螺钉与基座相连。注意电磁铁表面清洁。安装时，电磁铁尽量靠近分离机构内侧，形成图14。
步骤9、安装弹簧压缩工装。在每个支撑臂两侧各安装一个弹簧安装块，每个安装块使用M4×16内六角螺钉与支撑臂相连，形成图15。
步骤10、压缩弹射弹簧，逐个弹簧进行操作。先用工具将弹簧压下，再从弹簧套筒处穿入弹簧压条。每根弹簧压条两端均穿过M5×30内六角螺钉旋入上一步骤安装的弹簧安装块，拧入螺钉，使得弹簧顶杆最高处低于预紧件上表面，形成图16。
步骤11、先放置备份中心承力环，如无问题，则取下中心承力环，放置卫星。卫星的底部行程开关与分离机构底部压紧面接触。卫星的全景相机与分离机构底板的接地孔在相反象限，形成图17。
步骤12、放置锁紧销。安装前先涂抹润滑脂。注意锁紧销的平面与基座平面贴合，销仅能从基座单侧斜面放入，形成图18。
步骤13、安装两侧的锁紧滑块。安装前应先对锁紧滑块涂抹润滑脂。先按紧电磁铁的动芯为锁紧滑块腾出空间，再将解锁弹簧穿入锁紧滑块和锁紧释放基座中间(各有定位杆，图中未体现弹簧)，按下锁紧滑块保持，旋转电磁铁的动芯，使得其动芯进入锁紧滑块的腰孔，确认电磁铁的限位螺母贴合电磁铁外壳后，即完成本步骤安装。本步骤需测量动芯挡板与电磁铁外壳间距，要求为3.6～3.8mm。该尺寸可以在动芯末端限位螺母与电磁铁外壳之间增减0.1mm垫片控制，测量位置为电磁铁圆盘水平方向直径对称处，两处测量值取平均，形成 图19。其中，图20为间距控制图。
步骤14、安装2个KX-1行程开关，注意从支撑臂的孔中走线。每个行程开关通过2只M4×10内六角钛螺钉与行程开关限位基座相连。拧紧行程开关与行程开关基座之间的螺钉后，调整行程开关基座位置，测试行程开关信号，确认行程开关压缩量正确，再拧紧行程开关基座与支撑臂连接的8只M4螺钉，形成图21。
步骤15、调整卫星与锁紧销，施加预紧力。取出锁紧释放基座两侧的预紧件与预紧件支架之间的零点五厚工装，对4个预紧件的8个M6螺钉拧紧，需保持对称螺钉尽量同步施加预紧力。在拧紧过程中，需调整锁紧销和卫星的位置。锁紧销应该尽量居中放置，销侧面与基座槽内壁单侧距离应控制为6±0.5mm，卫星最外侧与基座最内侧相对面距离应控制为0.9±0.2mm，，形成图22。其中，图23为卫星最外侧与基座最内侧相对面距离0.9mm的示意图。图24为销侧面与基座槽内壁单侧距离6mm时的示意图。
步骤16、安装4个加强筋，每个加强筋与相邻2个支撑臂连接。
步骤17、卸下步骤9、10安装的弹簧压紧工装，使得弹簧顶杆顶住卫星，检查顶杆是否倾斜。此时测量卫星与分离机构之间接地电阻。
步骤18、对第1、3象限的预紧件施加预紧力，将M6螺钉拧紧。
步骤19、安装2个限位支架，每个支架通过4个M4×10内六角钛螺钉连接在支撑臂上。
步骤20、装配完成，形成图3。
本发明提供一种用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法，具有以下优点：
(1)可靠锁紧：
在合理的参数设计下，在复杂的力学环境下，能够实现卫星和火箭的可靠锁紧，并已通过力学试验进行检验，检验项目为过载、正弦振动、随机振动、冲击等。
(2)在释放指令给出后，能够实现卫星和火箭的可靠分离，能够有效地通过调整弹簧的设计参数控制分离速度。分离速度可以通过高速相机拍摄测量得到。
(3)能够在地面重复进行分离试验，不会对自身产生破坏，即：完成分离后可以用原有部件重复复原装配。
(4)分离完全为机构动作，无污染，即：分离后没有除卫星外其他任何物体分离。
(5)分离动作为连续机构动作，对卫星无冲击，通过传感器可以测试卫星受到的冲击响应。
(6)极大地降低了卫星研制的成本，相对于传统的火工品式的锁紧释放机构，单组机构可节约经济成本约20万元。
以上所述仅是本发明的优选实施方式，应当指出，对于本技术领域的普通技术人员来说，在不脱离本发明原理的前提下，还可以做出若干改进和润饰，这些改进和润饰也应视本发明的保护范围。

Claims (4)

  1. 一种用于星箭分离的电磁锁紧释放机构，其特征在于，包括：机架(1)、弹射单元、卫星单元、锁紧释放单元以及锁紧释放驱动单元；设置水平方向为X方向；垂直方向为Y方向；
    所述卫星单元包括卫星(7)、定位销(8)和卫星凸台(9)；所述卫星(7)通过所述定位销(8)与所述卫星凸台(9)固定连接；
    所述弹射单元设置于所述卫星单元的下方，用于向所述卫星单元提供Y正方向的弹射力；所述弹射单元包括弹射弹簧(2)、弹射顶杆(3)和弹簧套筒(4)；所述弹簧套筒(4)与所述机架(1)固定连接；所述弹簧套筒(4)内部设置所述弹射弹簧(2)；所述弹射顶杆(3)的底部位于所述弹簧套筒(4)内，并压于所述弹射弹簧(2)的上面；所述弹射顶杆(3)的顶部穿过所述弹簧套筒(4)，而延伸到所述卫星单元的下面，并与所述卫星单元的下表面接触；在锁紧状态下，所述弹射弹簧(2)被所述弹射顶杆(3)压紧，所述弹射弹簧(2)通过所述弹射顶杆(3)向所述卫星单元提供Y正方向的弹射力；在释放状态下，在所述弹射弹簧(2)的弹射力作用下，所述卫星单元向Y正方向运动，从而与所述机架(1)分离；
    所述锁紧释放单元包括锁紧销(10)、锁紧滑块(11)、解锁弹簧(12)和基座(13)；
    所述基座(13)位于所述卫星单元的左侧，并与所述机架(1)固定连接；所述基座(13)的右侧开设第一空腔，所述锁紧销(10)为D形状；所述卫星凸台(9)的上表面设计为斜面；在锁紧状态下，所述锁紧销(10)部分位于所述第一空腔内，所述锁紧销(10)的圆柱面压于所述卫星凸台(9)的斜面，进而限制所述卫星凸台(9)向Y正方向运动；此外，所述卫星凸台(9)的斜面向所述锁紧销(10)提供沿斜面法向的接触弹力；所述接触弹力分解为Y正方向接触弹力和X负方向接触弹力；Y正方向接触弹力使所述锁紧销(10)压紧所述基座(13)；X负方向接触弹力使所述锁紧销(10)压紧位于其后面的锁紧滑块(11)；
    所述基座(13)的左侧开设沿Y方向的导向槽，所述锁紧滑块(11)设置于所述导向槽内，使所述锁紧滑块(11)仅能进行Y方向运动；所述锁紧滑块(11)和所述基座(13)之间设置所述解锁弹簧(12)，所述解锁弹簧(12)向所述锁 紧滑块(11)提供Y正方向的推力；
    所述锁紧滑块(11)的下部开设沿X方向的空腔；所述锁紧滑块(11)位于所述锁紧销(10)的后面，在锁紧状态下，所述锁紧滑块(11)上部平面部分压于所述锁紧销(10)的后面；在释放状态下，所述锁紧滑块(11)在所述解锁弹簧(12)的作用下向Y正方向运动，从而使所述锁紧滑块(11)的空腔上升，所述锁紧销(10)在所述卫星凸台(9)提供的X负方向接触弹力的作用下，沿X负方向运动，进入所述锁紧滑块(11)的空腔内，从而使所述锁紧销(10)释放对所述卫星单元的运动阻碍；
    所述锁紧释放驱动单元包括电磁铁限位螺母(14)、电磁铁动芯(15)和电磁铁(16)；所述电磁铁(16)与所述机架(1)固定；所述电磁铁(16)内置线圈；所述电磁铁动芯(15)沿X方向设置，所述电磁铁(16)的内置线圈在没有通电时，所述电磁铁动芯(15)的右端压于所述锁紧滑块(11)的顶端，限制所述锁紧滑块(11)沿Y正方向运动；当所述电磁铁(16)的内置线圈通电时，所述电磁铁动芯(15)向X负方向运动，释放对所述锁紧滑块(11)的限位，使所述锁紧滑块(11)在解锁弹簧(12)的作用下向Y正方向运动。
  2. 根据权利要求1所述的用于星箭分离的电磁锁紧释放机构，其特征在于，还包括限位支架(17)；所述限位支架(17)固定安装于所述锁紧滑块(11)的上方，用于限制所述锁紧滑块(11)向Y正方向运动的距离。
  3. 根据权利要求1所述的用于星箭分离的电磁锁紧释放机构，其特征在于，还包括预紧单元；
    所述预紧单元包括预紧件支架(5)和预紧件(6)；
    所述预紧件支架(5)与所述机架(1)固定连接；所述预紧件(6)通过预紧螺钉与所述预紧件支架(5)连接；并且，所述预紧件(6)位于所述卫星凸台(9)的下面；通过旋转所述预紧螺钉，使所述预紧件(6)沿Y正方向运动，直到所述预紧件(6)的上表面与所述卫星凸台(9)的下表面紧密接触；同时，在所述预紧螺钉拧紧过程中，所述预紧件(6)推动所述卫星凸台(9)沿Y正方向运动，使所述卫星凸台(9)和所述锁紧销(10)之间的接触产生弹性形变，消除接触间隙。
  4. 一种基于权利要求1-3任一项所述的用于星箭分离的电磁锁紧释放机构的电磁锁紧释放方法，其特征在于，包括以下步骤：
    步骤1，火箭在到达预定轨道之前，卫星与火箭通过以下方式实现锁紧：
    步骤1.1，电磁铁(16)的内置线圈通电，从而使电磁铁动芯(15)的右端压于锁紧滑块(11)的顶端，使锁紧滑块(11)位于导向槽的下方；
    步骤1.2，当锁紧滑块(11)位于导向槽的下方时，锁紧滑块(11)的上部平面部分压于锁紧销(10)的后面，限制锁紧销(10)沿X负方向运动，从而使锁紧销(10)的右端凸出于基座(13)的第一空腔；
    步骤1.3，因此，锁紧销(10)的圆柱面压于卫星凸台(9)的斜面，进而限制卫星凸台(9)向Y正方向运动；同时，弹射顶杆(3)在弹射弹簧(2)的作用下向卫星单元提供Y正方向的弹射力，但该弹射力小于锁紧销(10)向卫星凸台(9)施加的力；通过锁紧销(10)和弹射顶杆(3)的共同作用，使卫星单元锁紧在机架(1)上面；
    步骤2，在火箭入轨后，给出分离指令，从而使卫星安全释放，具体包括以下步骤：
    步骤2.1，在火箭入轨后，给出分离指令，从而使电磁铁(16)的内置线圈断电；
    当电磁铁(16)的内置线圈断电时，电磁铁动芯(15)沿X负方向运动，从而使电磁铁动芯(15)的右端离开锁紧滑块(11)的顶端，释放对锁紧滑块(11)在Y方向的限位作用；
    步骤2.2，当电磁铁动芯(15)释放对锁紧滑块(11)在Y方向的限位作用时，锁紧滑块(11)在解锁弹簧(12)的作用下向Y正方向运动，并通过限位支架(17)对锁紧滑块(11)的运动距离进行限制，从而使锁紧滑块(11)的空腔上升一定高度；
    步骤2.3，当锁紧滑块(11)的空腔上升一定高度时，释放对锁紧销(10)在X负方向运动的阻碍；
    因此，锁紧销(10)在卫星凸台(9)的斜面提供的X负方向接触弹力的作用下，沿X负方向运动，并部分进入到锁紧滑块(11)的空腔中，从而使锁紧销 (10)脱离对卫星凸台(9)沿Y正方向的限位；
    步骤2.4，当锁紧销(10)脱离对卫星凸台(9)沿Y正方向的限位时，在弹射弹簧(2)的弹射力作用下，使弹射顶杆(3)沿Y正方向运动，从而使包含卫星凸台(9)的卫星单元沿Y正方向弹射发射，实现卫星安全释放。
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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113665844A (zh) * 2021-09-10 2021-11-19 上海卫星工程研究所 用于堆叠卫星与运载分离的星箭解锁装置
CN113955157A (zh) * 2021-11-26 2022-01-21 深圳力合精密装备科技有限公司 航天重复锁紧系统

Families Citing this family (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111284731B (zh) 2020-01-16 2020-11-10 清华大学 用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法
CN111924131B (zh) * 2020-07-01 2022-01-04 北京卫星制造厂有限公司 一种在轨可重复使用的多功能电磁接口装置
CN111806733B (zh) * 2020-07-13 2021-11-09 中北大学 一种卫星对接的锁紧与解锁装置
CN112061420B (zh) * 2020-09-15 2021-11-09 哈尔滨工业大学 一种连杆触发式锁释装置及其工作方法
CN112208803B (zh) * 2020-09-30 2021-11-30 哈尔滨工业大学 一种可实现多星顺序释放的锁紧分离机构及其工作方法
CN112697184B (zh) * 2020-11-30 2023-03-24 上海航天控制技术研究所 一种火星探测器的分离伴星的底座
CN113879569A (zh) * 2021-10-09 2022-01-04 航天科工火箭技术有限公司 低冲击解锁星箭分离装置及星箭分离系统
CN114132534B (zh) * 2021-11-02 2024-03-29 南京航空航天大学 一种低冲击大承载堆栈式多星锁紧释放机构及其工作方法
CN113911404B (zh) * 2021-11-26 2024-01-23 深圳力合精密装备科技有限公司 航天重复锁紧系统
CN113998154B (zh) * 2021-11-26 2024-01-23 深圳力合精密装备科技有限公司 卫星重复锁紧装置
CN114353603B (zh) * 2022-01-13 2023-11-14 北京中科宇航技术有限公司 一种非火工品星箭分离解锁装置
CN115610703B (zh) * 2022-10-26 2024-05-10 深圳市魔方卫星科技有限公司 一种抽拉式压紧释放装置及使用方法
CN116086678B (zh) * 2023-04-10 2023-06-30 东方空间技术(山东)有限公司 一种气体推冲装置能量测试系统

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4679752A (en) * 1985-01-28 1987-07-14 Hughes Aircraft Company Payload depolyment from shuttle employing an ejection restraint device
GB2385310A (en) * 2002-02-14 2003-08-20 Insys Ltd Satellite launch assembly including means for connection and release
CN107054700A (zh) * 2017-03-30 2017-08-18 哈尔滨工业大学（威海） 一种星箭分离用锁紧释放装置
CN108116697A (zh) * 2017-12-08 2018-06-05 浙江大学 一种球形卫星分离解锁装置
CN108516111A (zh) * 2018-04-26 2018-09-11 孝感呈睿知识产权服务有限公司 一种电磁约束式分离作动器及其制造方法
CN108583940A (zh) * 2018-06-04 2018-09-28 上海微小卫星工程中心 一种四点式星箭分离机构
CN110329552A (zh) * 2019-08-08 2019-10-15 北京微分航宇科技有限公司 一种多点压紧同步解锁的星箭适配器
CN111284731A (zh) * 2020-01-16 2020-06-16 清华大学 用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4324374A (en) * 1978-05-30 1982-04-13 Hughes Aircraft Company Integrated spacecraft and cradle structure
US4506852A (en) * 1980-03-18 1985-03-26 Hughes Aircraft Company Payload deployment from launch vehicle using canted springs for imparting angular and linear velocities
FR2735099B1 (fr) * 1995-06-06 1997-08-29 Aerospatiale Dispositif adaptateur basculant pour l'emport de plusieurs charges utiles sur un meme lanceur.
US5695306A (en) * 1996-05-08 1997-12-09 Lockheed Martin Corp. Fusible member connection apparatus and method
CN2286723Y (zh) * 1996-10-23 1998-07-29 中国科学院上海技术物理研究所 一种抛罩机构
US6126115A (en) * 1997-01-21 2000-10-03 Lockheed Martin Corporation Apparatus for retaining and releasing a payload
ES2131476B1 (es) * 1997-09-26 2000-03-01 Const Aeronauticas Sa Sistema de fijacion y separacion de satelites.
JP4211131B2 (ja) * 1999-04-06 2009-01-21 株式会社Ｉｈｉ ペイロード制振機構
US6896441B1 (en) * 2002-08-14 2005-05-24 Lockheed Martin Corporation Automated latching device with active damping
CN100546876C (zh) * 2007-11-21 2009-10-07 航天东方红卫星有限公司 星星或星箭分离机构
CN107089349A (zh) * 2017-04-24 2017-08-25 上海航天控制技术研究所 一种对接后可旋转的微纳卫星电磁对接装置及对接方法
CN107826271B (zh) * 2017-12-07 2020-10-16 上海宇航系统工程研究所 卡爪式星箭连接分离机构

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4679752A (en) * 1985-01-28 1987-07-14 Hughes Aircraft Company Payload depolyment from shuttle employing an ejection restraint device
GB2385310A (en) * 2002-02-14 2003-08-20 Insys Ltd Satellite launch assembly including means for connection and release
CN107054700A (zh) * 2017-03-30 2017-08-18 哈尔滨工业大学（威海） 一种星箭分离用锁紧释放装置
CN108116697A (zh) * 2017-12-08 2018-06-05 浙江大学 一种球形卫星分离解锁装置
CN108516111A (zh) * 2018-04-26 2018-09-11 孝感呈睿知识产权服务有限公司 一种电磁约束式分离作动器及其制造方法
CN108583940A (zh) * 2018-06-04 2018-09-28 上海微小卫星工程中心 一种四点式星箭分离机构
CN110329552A (zh) * 2019-08-08 2019-10-15 北京微分航宇科技有限公司 一种多点压紧同步解锁的星箭适配器
CN111284731A (zh) * 2020-01-16 2020-06-16 清华大学 用于星箭分离的电磁锁紧释放机构以及电磁锁紧释放方法

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113665844A (zh) * 2021-09-10 2021-11-19 上海卫星工程研究所 用于堆叠卫星与运载分离的星箭解锁装置
CN113665844B (zh) * 2021-09-10 2023-09-19 上海卫星工程研究所 用于堆叠卫星与运载分离的星箭解锁装置
CN113955157A (zh) * 2021-11-26 2022-01-21 深圳力合精密装备科技有限公司 航天重复锁紧系统
CN113955157B (zh) * 2021-11-26 2023-08-29 深圳力合精密装备科技有限公司 航天重复锁紧系统

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