A monopropellant engine with low thrust and high specific impulse

Abstract

The present application relates to a monopropellant engine with low thrust and high specific impulse, belonging to the technical field of spacecraft engines. The monopropellant engine with low thrust and high specific impulse includes an injector, the injector includes a capillary tube, the capillary tube includes a straight tube section and a curved tube section connected to the straight tube section, the curved tube section has a plurality of bending parts in three-dimensional space, two straight tube sections are relatively arranged, the curved tube section is located between the two straight tube sections, the curved tube section includes a spiral bending part and a transition bending part, the spiral bending part is located between the two straight tube sections, the transition bending part is connected between the straight tube section and the spiral bending part, and the spiral bending part is arranged in a circle. The present application has the effect of facilitating the precise control of the flow rate and pressure drop of the propellant, and realizing the precise control of the thrust of the special type of monopropellant engine.

Description

Technical Field

[0001] The present application relates to the technical field of spacecraft engines, and in particular to a single-propellant engine with low thrust and high specific impulse. Background Art

[0002] With the rapid development of the aerospace economy, the demand for microsatellites in the domestic and international markets has increased significantly. Compared with traditional large satellites, microsatellites have the advantages of light weight, small size, low cost, easy mass production and multiple satellites per launch. Microsatellites usually use monopropellant chemical propulsion systems to provide the forces and torques required for attitude control, orbit control and orbital maneuvers.

[0003] A monopropellant engine is a specific type of rocket engine. Due to its simple structure and reliable performance, it is widely used for attitude adjustment and orbit change of spacecraft such as satellites, spacecraft, and launch vehicles. It has the advantages of low thrust, pulse operation, long service life, large adjustment range, and high impulse control accuracy. A monopropellant engine mainly includes a solenoid valve, an injector, a thrust chamber, and a thermal control component. The injector has a capillary tube, which is a flow pressure drop control part of the injector to achieve flow pressure drop control of the propellant. At present, the capillary tube in the injector is mostly a straight tube capillary tube or a capillary tube with a C-shaped bend in the middle of the straight tube. When under a preset pressure drop, it is difficult for the relevant capillary tube to reach the preset flow coefficient, resulting in difficulty in accurately controlling the flow and pressure drop of the propellant. Summary of the invention

[0004] In order to help accurately control the flow rate and pressure drop of the propellant and to achieve accurate control of the thrust of a special type of monopropellant engine to a certain extent, the present application provides a monopropellant engine with low thrust and high specific impulse.

[0005] The present application provides a low-thrust, high-specific-impulse monopropellant engine that adopts the following technical solution:

[0006] A monopropellant engine with low thrust and high specific impulse, comprising an injector, the injector comprising a capillary tube, the capillary tube comprising a straight tube section and a curved tube section connected to the straight tube section, the curved tube section having a plurality of bending portions in three-dimensional space, two straight tube sections being arranged opposite to each other, the curved tube section being located between the two straight tube sections, the curved tube section comprising a spiral bending portion and a transition bending portion, the spiral bending portion being located between the two straight tube sections, the transition bending portion being connected between the straight tube section and the spiral bending portion, the spiral bending portion being arranged in a circle, the total length of the capillary tube being L, L=250D, wherein D is the inner diameter of the capillary tube;

[0007] The straight length of the capillary is L1, L1=0.65L;

[0008] The length of the straight pipe section is L2, L2=0.2333L;

[0009] The bending radius of the spiral bending portion is R1, R1=0.055L;

[0010] The bending radius of the transition bending portion is R2, R2=0.125L.

[0011] Preferably, the inner diameter D of the capillary is 0.12 mm-0.2 mm.

[0012] Preferably, the outer diameter of the capillary is 0.32 mm, and the inner diameter D is 0.12 mm.

[0013] Preferably, the injector also includes a flange, a frame, an antechamber and a splash screen, the frame is connected between the flange and the antechamber, the frame is a hollow cylindrical structure, the flange is used to provide a connecting and mounting surface for the installation of the frame, the capillary is located in the frame, the straight pipe section at one end of the capillary is connected to the center hole of the flange, and the straight pipe section at the other end is connected to the center hole of the antechamber, and the splash screen is arranged in the antechamber.

[0014] Preferably, the diameter of the interior of the front chamber gradually increases toward the side away from the frame, and the splash screen is arranged on the side of the interior of the front chamber away from the frame.

[0015] Preferably, the monopropellant engine further comprises a solenoid valve for realizing on-off control of the propellant, and the solenoid valve is connected to a flange on the frame.

[0016] Preferably, the monopropellant engine also includes a thrust chamber, which includes a catalyst bed and a nozzle arranged on the catalyst bed, the catalyst bed is connected to the end of the front chamber away from the frame, the nozzle is located on the side of the catalyst bed away from the front chamber, and the nozzle is a truncated ideal nozzle.

[0017] Preferably, the monopropellant engine also includes a thermal control component, which includes a heating section, a transition section and a lead-out section connected in sequence, the heating section is sleeved on the outer wall of the catalyst bed, and is used to control and monitor the temperature of the catalyst bed, the transition section is located on the side of the heating section close to the frame, the transition section adopts a thin-walled variable diameter armored tube, and the lead-out section contains a high-temperature epoxy resin material, which is used to seal the terminal lead.

[0018] In summary, this application includes the following beneficial technical effects:

[0019] The design operating conditions of the monopropellant engine are: inlet pressure 1.2Mpa; combustion chamber pressure 588Kpa; propellant flow rate 87.3mg / s; vacuum thrust 0.2N; according to the above monopropellant engine parameters, the design operating conditions of the capillary as a flow control element are: pressure drop of 0.612Mpa, flow coefficient of 0.2228; the capillary is designed to be a structure in which a straight tube section, a transition bend, a spiral bend, a transition bend and a straight tube section are connected in sequence, and the size parameters of the capillary are combined, so that the resistance can be adjusted by the effective length and twist of the capillary; and after many tests, the capillary of a specific shape in the present application is used in combination with the above specific size parameters of the capillary, so that the error of the flow coefficient compared with the design value is less than ±2%, which indicates that under the special shape of the capillary in the present application and the setting of the size parameters of the capillary, the flow rate and pressure drop of the special type of engine propellant under the design condition can be accurately controlled, thereby achieving accurate control of the thrust of the special type of monopropellant engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a schematic diagram of the overall structure of an embodiment of the present application.

[0021] Figure 2 It is an exploded diagram of the overall structure of an embodiment of the present application.

[0022] Figure 3 It is a sectional view of the overall structure of the injector in the embodiment of the present application.

[0023] Figure 4 It is a schematic diagram of the overall structure of the capillary in the embodiment of the present application.

[0024] Figure 5 It is a cross-sectional view of the overall structure of the thrust chamber in the embodiment of the present application.

[0025] Figure 6 It is a flow pressure drop characteristic diagram of a 25 mm unwound capillary drawn according to Table 1.

[0026] Figure 7 It is a flow pressure drop characteristic diagram of a 30 mm unwound capillary drawn based on Table 2.

[0027] Figure 8 It is a flow pressure drop characteristic diagram of a 35mm unwound capillary drawn based on Table 3.

[0028] Fig. 9 It is a 30mmπ-type capillary flow pressure drop characteristic diagram drawn according to Table 4.

[0029] Fig.10 It is a spiral capillary flow pressure drop characteristic diagram drawn according to Table 5.

[0030] Explanation of the reference numerals in the accompanying drawings: 1. Injector; 2. Capillary; 21. Straight pipe section; 22. Bend pipe section; 221. Spiral bend; 222. Transition bend; 3. Flange; 4. Frame; 5. Antechamber; 6. Splash screen; 7. Solenoid valve; 8. Thrust chamber; 81. Catalyst bed; 82. Nozzle; 9. Thermal control assembly; 91. Heating section; 92. Transition section; 93. Lead-out section; 10. Bolt; 11. Aluminum washer; 12. Expansion profile; 13. Groove; 14. Partition; 15. Partition screen; 16. First clamping notch; 17. Second clamping notch; 18. Retaining ring. DETAILED DESCRIPTION

[0031] The following is combined with Figure 1-10 This application is described in further detail.

[0032] The present application embodiment discloses a monopropellant engine with low thrust and high specific impulse. Figure 1 and Figure 2 A monopropellant engine with low thrust and high specific impulse comprises a solenoid valve 7, an injector 1, a thrust chamber 8 and a thermal control component 9, wherein the solenoid valve 7, the injector 1 and the thrust chamber 8 are connected in sequence, the thermal control component 9 is arranged on the thrust chamber 8, the solenoid valve 7 is used to realize the on-off control of the propellant, the injector 1 is used to realize the flow pressure drop control of the propellant and to block the heat of the thrust chamber 8 from being transferred to the solenoid valve 7.

[0033] Reference Figure 2 and Figure 3 Specifically, the injector 1 includes a flange 3, a frame 4, a capillary 2, a front chamber 5 and a splash screen 6. The frame 4 is fixed between the flange 3 and the front chamber 5. The frame 4 and the flange 3 are integrally formed. The front chamber 5 is fixedly connected to the frame 4 by brazing. The center positions of the flange 3 and the front chamber 5 are both provided with a center hole. The solenoid valve 7 is connected to the flange 3 of the frame 4 by bolts 10. The outlet of the solenoid valve 7 and the center hole of the flange 3 are sealed by an aluminum gasket 11. A spring gasket and a flat gasket are provided at the connection of the bolt 10 to ensure the connection stability. The flange 3 is used to provide a connection and installation surface for the connection between the injector 1 and the solenoid valve 7, and the engine and the satellite cabin board. When the engine is installed, the flange 3 is connected to the satellite cabin board by bolts and is heat-insulated by a heat-insulating gasket.

[0034] Reference Figure 2 and Figure 3, the frame 4 is a hollow cylindrical structure, which can effectively block the heat of the thrust chamber 8 from being transferred to the solenoid valve 7; the capillary 2 is a flow pressure drop control part of the injector 1. In the embodiment of the present application, the injector 1 is a single capillary 2 injection structure. Specifically, the capillary 2 is located in the frame 4, one end of the capillary 2 is connected to the center hole of the flange 3 by brazing, and the other end is connected to the center hole of the front chamber 5 by brazing. The capillary 2 includes a straight pipe section 21 and a curved pipe section 22 connected to the straight pipe section 21, and the curved pipe section 22 has multiple bending parts in three-dimensional space.

[0035] Reference Figure 2 and Figure 3 There are two straight pipe sections 21 arranged opposite to each other, and the curved pipe section 22 is located between the two straight pipe sections 21. The straight pipe section 21 at one end is connected to the central hole of the flange 3, and the straight pipe section 21 at the other end is connected to the central hole of the front chamber 5, wherein the curved pipe section 22 includes a spiral bending portion 221 and a transition bending portion 222, the spiral bending portion 221 is located between the two straight pipe sections 21, and the transition bending portion 222 is connected between the straight pipe section 21 and the spiral bending portion 221, and the spiral bending portion 221 is arranged in a circle. The spiral bending portion 221 and the transition bending portion 222 are both curved in three-dimensional space, which helps to ensure the effective length and twist of the capillary 2, thereby twisting the flow path of the propellant in three-dimensional space, adjusting the resistance, and helping to accurately control the flow rate and pressure drop of the propellant; and the two straight pipe sections 21 are connected to the flange 3 and the front chamber 5 respectively, which can ensure the smooth entry and output of the propellant.

[0036] Reference Figure 3 and Figure 4 , the total length of the capillary 2 is L, L=250D, where D is the inner diameter of the capillary 2; the straight length of the capillary 2 is L1, L1=0.65L; the length of the straight tube section 21 is L2, L2=0.2333L; the bending radius of the spiral bending portion 221 is R1, R1=0.055L; the bending radius of the transition bending portion 222 is R2, R2=0.125L. The design operating conditions of the monopropellant engine are: inlet pressure 1.2Mpa; combustion chamber pressure 588Kpa; propellant flow rate 87.3mg / s; vacuum thrust 0.2N. According to the above monopropellant engine parameters, the design operating conditions of the capillary 2 as a flow control element are: pressure drop of 0.612Mpa, flow coefficient of 0.2228.

[0037] During operation, the propellant enters the end of the capillary 2 near the flange 3 through the on-off control of the electromagnetic valve 7, and passes through the straight tube section 21, the transition bend 222, the spiral bend 221, the transition bend 222 and the straight tube section 21 of the capillary 2 in sequence, and then enters the front chamber 5; the propellant enters the thrust chamber 8 through the front chamber 5, and then the propellant is catalytically decomposed under the action of the catalyst in the thrust chamber 8, and then ejected to generate thrust. In the present application, the capillary 2 is designed as a structure including the straight tube section 21, the transition bend 222, the spiral bend 221, the transition bend 222 and the straight tube section 21 connected in sequence, which can adjust the resistance through the effective length and twist of the capillary 2, thereby facilitating the precise control of the flow rate and pressure drop of the propellant, and to a certain extent realizing the precise control of the thrust of the monopropellant engine. After many tests, the capillary 2 of the specific shape in the present application is used in combination with the above specific parameters of the capillary 2. When the pressure drop is 0.612 MPa, the flow coefficient is 0.219, so that the error of the flow coefficient compared with the design value is less than ±2%. This shows that through the setting of this specially formed ultrafine capillary 2 and the size parameters of the capillary 2, it is possible to achieve precise control of the propellant flow and pressure drop of this special type of engine, and to achieve precise control of the thrust of this special type of monopropellant engine, meeting the situation of small flow and small thrust of the monopropellant engine.

[0038] Reference Figure 3 and Figure 4 Specifically, the inner diameter D of the capillary 2 is 0.12mm-0.2mm. Preferably, the outer diameter of the capillary 2 is 0.32mm, the inner diameter D is 0.12mm, the total length L of the capillary 2 is 30mm, the straight length L1 of the capillary 2 is 19.5mm, the length L2 of the straight tube section 21 is 6.999mm, the bending radius R1 of the spiral bending portion 221 is 1.65mm, and the bending radius R2 of the transition bending portion 222 is 3.75mm.

[0039] Reference Figure 2 and Figure 3 The diameter of the front chamber 5 gradually increases toward the side away from the frame 4 to form an expansion surface 12. One end of the front chamber 5 is connected to the frame 4 by brazing, and the other end is connected to the thrust chamber 8 by electron beam welding. The outer wall of the front chamber 5 close to the frame 4 and the inner wall close to the thrust chamber 8 are both provided with grooves 13 to form a stepped end surface, which is convenient for connection with the frame 4 and the thrust chamber 8; the splash screen 6 is connected to the side of the front chamber 5 away from the frame 4 by resistance welding, that is, the splash screen 6 is arranged at the outlet of the expansion surface 12, and the splash screen 6 is located in the stepped groove of the front chamber 5 close to the thrust chamber 8.

[0040] The propellant entering the front chamber 5 through the capillary 2 is decelerated when passing through the expansion surface 12, and then is evenly sprayed into the thrust chamber 8 through the splash screen 6, which helps to ensure the uniformity of the propellant entering the thrust chamber 8.

[0041] Reference Figure 3 and Figure 5 , the thrust chamber 8 includes a catalyst bed 81 and a nozzle 82. One end of the catalyst bed 81 is connected to the stepped end face of the front chamber 5 away from the frame 4 by electron beam welding, and the welding depth is 0.8mm-1mm. The other end is integrally formed with the nozzle 82, so that the nozzle 82 is located on the side of the catalyst bed 81 away from the front chamber 5. The nozzle 82 is a truncated ideal nozzle 82. Specifically, the expansion section of the nozzle 82 is a truncated ideal nozzle aerodynamic profile designed using the Sivells characteristic line method. Among them, a partition 14 and a partition net 15 are fixedly arranged in the catalyst bed 81. The partition net 15 is located on the side of the partition 14 close to the front chamber 5. The partition net 15 and the partition 14 are mainly used to divide the catalyst bed 81. In other embodiments, the truncated ideal nozzle aerodynamic profile can be replaced by a conical nozzle aerodynamic profile.

[0042] When the propellant enters the catalyst bed 81 from the front chamber 5, the propellant is catalytically decomposed under the action of the catalyst in the catalyst bed 81, and then ejected through the nozzle 82 to generate thrust. The nozzle 82 adopts a shortened ideal nozzle aerodynamic profile. Compared with nozzles 82 with other aerodynamic profiles, there is no shock wave in the nozzle 82, and the decomposed propellant undergoes isentropic expansion in it before being ejected; the friction loss and shock wave loss in the nozzle 82 are significantly reduced, which can effectively reduce the thickness of the boundary layer of the propellant in the nozzle 82 and improve the specific impulse of the engine.

[0043] Reference Figure 2 The thermal control assembly 9 is used to control and monitor the temperature of the catalyst bed 81. Specifically, the thermal control assembly 9 includes a heating section 91, a transition section 92 and a lead section 93 which are fixedly connected in sequence. The heating section 91 is fixedly sleeved on the outer wall of the catalyst bed 81 by a welding retaining ring 18. The welding depth is 0.4mm-0.5mm. The retaining ring 18 is connected to the catalyst bed 81 by electron beam welding. The heating section 91 contains a high resistance density heating core to provide heating power for temperature control and monitoring of the catalyst bed 81. The transition section 92 is located on the side of the heating section 91 close to the frame 4. The transition section 92 uses a thin-walled variable diameter armored tube to help reduce heat transfer to the lead section 93 and prevent heat back-invasion. The lead section 93 contains a high-temperature epoxy resin material, which is used to seal the terminal lead in the thermal control assembly 9 and has high heat resistance and chemical stability. The thermal control assembly 9 integrates the two functions of heating and temperature measurement, thereby reducing the power consumption and weight of the monopropellant engine. The present application provides a small thrust monopropellant engine with high specific impulse, low power consumption, small size and light weight, which provides the force and torque required for attitude control, orbit control and orbit maneuvering for micro-satellites, and helps to expand the application space of micro-satellites. In other embodiments, the armored integrated thermal control component 9 can be replaced by an armored heater and a platinum resistor.

[0044] Reference Figure 2 A first snap-in notch 16 for the transition section 92 to pass through is provided on one side of the frame 4 connected to the front chamber 5; a second snap-in notch 17 snap-fitted with the lead-out section 93 is provided on the flange 3, making the engine structure compact and reducing the size.

[0045] The implementation principle of the embodiment of the present application is as follows: during operation, the propellant enters the capillary 2 through the on-off control of the solenoid valve 7, and enters the front chamber 5 through the straight tube section 21, the transition bend 222, the spiral bend 221, the transition bend 222 and the straight tube section 21 of the capillary 2 in sequence, decelerates when passing through the expansion profile 12 of the front chamber 5, and then is evenly sprayed into the catalyst bed 81 through the splash screen 6. The propellant is catalytically decomposed under the action of the catalyst in the catalyst bed 81, and then is ejected through the nozzle 82 to generate thrust; wherein, the thermal control component 9 controls and monitors the temperature of the catalyst bed 81. In the present application, the capillary 2 is designed as a structure in which the straight tube section 21, the transition bend 222, the spiral bend 221, the transition bend 222 and the straight tube section 21 are connected in sequence. In combination with the size parameters of the capillary 2, the resistance can be adjusted by changing the effective length and distortion of the capillary 2, thereby facilitating the precise control of the flow rate and pressure drop of the propellant. After many tests, the capillary 2 of a specific shape in the present application is used in combination with the above specific parameters of the capillary 2. When the pressure drop is 0.612 MPa, the flow coefficient is 0.219, so that the error of the flow coefficient compared with the design value is less than ±2%. This shows that by setting the size parameters of this specially formed ultrafine capillary 2 and the capillary 2, it is possible to achieve precise control of the propellant flow and pressure drop of this special type of engine, and to achieve precise control of the thrust of this special type of monopropellant engine.

[0046] The flow rate and pressure drop characteristics of capillaries of different shapes are compared below.

[0047] 1) Unwound capillary, i.e. straight capillary

[0048] Comparison of flow rate and pressure drop characteristics of unwound capillaries of different lengths:

[0049] (1) 25 mm unwound capillary

[0050]

[0051] It can be seen from Table 1 that when the pressure drop of a 25 mm unwound capillary is 0.612 MPa, the flow coefficient is between 0.3050 and 0.3308, which is 37% to 48% larger than the design value. Figure 6 The relationship diagram is drawn based on the flow pressure drop characteristics of the 25mm unwound capillary in Table 1 above. Figure 6It can be seen that the flow pressure drop characteristics of the 25mm unwound capillary are consistent, but cannot meet the precise control of the propellant flow and pressure drop. The winding process must be used to increase the pressure drop and reduce the flow coefficient. However, the length of the capillary after winding is reduced and cannot meet the engine structure size requirements.

[0052] (2) 30 mm unwound capillary

[0053]

[0054] It can be seen from Table 2 that when the pressure drop of a 30 mm unwound capillary is 0.612 MPa, the flow coefficient is between 0.2624 and 0.2750, which is 18% to 23% larger than the design value. Figure 7 is a relationship diagram drawn based on the flow pressure drop characteristics of the 30 mm unwound capillary in Table 2 above. Figure 7 It can be seen that the flow rate and pressure drop characteristics of the 30 mm unwound capillary have good consistency, but cannot meet the precise control of the propellant flow rate and pressure drop.

[0055] (3) 35 mm unwound capillary

[0056]

[0057] It can be seen from Table 3 that when the pressure drop of a 35mm unwound capillary is 0.612MPa, the flow coefficient is between 0.2202 and 0.2329, and the design value is within this flow coefficient range. However, due to the limitation of the engine structure size, the capillary of this length cannot meet the engine structure size requirements, and the length must be reduced by winding, which will lead to an increase in pressure drop and a decrease in flow coefficient, and cannot meet the precise control of propellant flow and pressure drop. Figure 8 This is a relationship diagram drawn based on the flow pressure drop characteristics of the 35mm unwound capillary in Table 3. Figure 8 It can be seen that the flow pressure drop characteristics of the 35 mm unwound capillary tube have good consistency, but due to the size limitation of the engine structure, the unwound capillary tube of this length cannot be used.

[0058] By comparing the flow and pressure drop characteristics of the above unwound capillaries of different lengths, a capillary with a length of 30 mm was selected. Through the special winding process of this application, precise control of the propellant flow and pressure drop can be achieved, and the engine structure size requirements can be met.

[0059] 2) 30mm wound capillary

[0060] The flow rate and pressure drop characteristics of capillaries with a length of 30 mm and different winding processes are compared:

[0061] (1) π-type capillary, that is, a capillary with a C-shaped bend in the middle of a straight tube

[0062] Two capillaries with a length of 30 mm were selected and wound in a π-type, that is, the middle part was wound into a semicircular arc. As shown in Table 4, when the pressure drop of the 30 mm π-type capillary was 0.612 MPa, the flow coefficient was 0.1940, which was 12.9% smaller than the design value, indicating that the π-type capillary could not meet the precise control of the propellant flow rate and pressure drop.

[0063]

[0064] Fig. 9 This is a relationship diagram drawn based on the flow pressure drop characteristics of the 30mm π-type capillary in Table 4 above. Fig. 9 It can be seen that the flow rate and pressure drop characteristics of the 30 mm π-type capillary have good consistency, but cannot meet the precise control of the propellant flow rate and pressure drop.

[0065] (2) Spiral capillary in this application

[0066] Six capillaries with a length of 30 mm were selected and wound according to the spiral type in the present application. As shown in Table 5, when the pressure drop of the spiral capillary in the present application is 0.612 MPa, the flow coefficient is between 0.2193 and 0.2277, and the error of the flow coefficient compared with the design value is less than ±2%, indicating that the use of this spiral capillary can achieve accurate control of the propellant flow rate and pressure drop, thereby accurately controlling the thrust of the monopropellant thruster.

[0067]

[0068] Fig.10 is a relationship diagram drawn based on the flow pressure drop characteristics of the spiral ultrafine capillary in Table 5 above, Fig.10 It can be seen that the flow rate and pressure drop characteristics of the 30mm spiral ultrafine capillary are consistent, which reduces the processing difficulty and manufacturing error, and improves the product qualification rate and production efficiency.

[0069] In summary, under the same operating parameters, when the pressure drop is 0.612MPa, the flow coefficient of the straight tube capillary and the π-type capillary has a large deviation from the design value, which cannot meet the precise control of the propellant flow and pressure drop. However, when the pressure drop is 0.612MPa, the flow coefficient of the specially formed spiral ultrafine capillary in this application is 0.219, and the error of the flow coefficient compared with the design value is less than ±2%, thereby achieving not only precise control of the propellant flow and pressure drop, but also precise control of the thrust of the monopropellant engine; it can also meet the requirements of the engine structure size.

[0070] The above are all preferred embodiments of the present application, and the protection scope of the present application is not limited thereto. Therefore, any equivalent changes made according to the structure, shape, and principle of the present application should be included in the protection scope of the present application.

Claims

1. A low-thrust, high-specific-impulse monopropellant engine, comprising an injector (1), characterized in that: The injector (1) comprises a capillary tube (2), the capillary tube (2) comprising a straight tube section (21) and a curved tube section (22) connected to the straight tube section (21), the curved tube section (22) having a plurality of bending portions in three-dimensional space, two straight tube sections (21) being arranged opposite to each other, the curved tube section (22) being located between the two straight tube sections (21), the curved tube section (22) comprising a spiral bending portion (221) and a transition bending portion (222), the spiral bending portion (221) being located between the two straight tube sections (21), the transition bending portion (222) being connected between the straight tube section (21) and the spiral bending portion (221), the spiral bending portion (221) being arranged in a circle, the total length of the capillary tube (2) being L, L=250D, wherein D is the inner diameter of the capillary tube (2); The straight length of the capillary (2) is L1, L1=0.65L; The length of the straight pipe section (21) is L2, L2=0.2333L; The bending radius of the spiral bending portion (221) is R1, R1=0.055L; The bending radius of the transition bending portion (222) is R2, where R2=0.125L.

2. A low thrust and high specific impulse monopropellant engine according to claim 1, characterized in that: The inner diameter D of the capillary (2) is 0.12 mm-0.2 mm.

3. A low thrust and high specific impulse monopropellant engine according to claim 2, characterized in that: The outer diameter of the capillary (2) is 0.32 mm, and the inner diameter D is 0.12 mm.

4. A low thrust and high specific impulse monopropellant engine according to any one of claims 1 to 3, characterized in that: The injector (1) further comprises a flange (3), a frame (4), an antechamber (5) and a splash screen (6); the frame (4) is connected between the flange (3) and the antechamber (5); the frame (4) is a hollow cylindrical structure; the flange (3) is used to provide a connection and mounting surface for mounting the frame (4); the capillary (2) is located in the frame (4); a straight pipe section (21) at one end of the capillary (2) is connected to the center hole of the flange (3); and a straight pipe section (21) at the other end is connected to the center hole of the antechamber (5); and the splash screen (6) is arranged in the antechamber (5).

5. A low thrust and high specific impulse monopropellant engine according to claim 4, characterized in that: The diameter of the interior of the front chamber (5) gradually increases toward the side away from the frame (4), and the splash screen (6) is arranged on the side of the interior of the front chamber (5) away from the frame (4).

6. A low thrust and high specific impulse monopropellant engine according to claim 4, characterized in that: The monopropellant engine also includes a solenoid valve (7) for realizing on-off control of the propellant, wherein the solenoid valve (7) is connected to the flange (3) on the frame (4).

7. A low thrust and high specific impulse monopropellant engine according to claim 4, characterized in that: The monopropellant engine further comprises a thrust chamber (8), wherein the thrust chamber (8) comprises a catalyst bed (81) and a nozzle (82) arranged on the catalyst bed (81), wherein the catalyst bed (81) is connected to an end of the front chamber (5) away from the frame (4), and the nozzle (82) is located on a side of the catalyst bed (81) away from the front chamber (5), and the nozzle (82) is a truncated ideal nozzle (82).

8. A low thrust and high specific impulse monopropellant engine according to claim 7, characterized in that: The monopropellant engine further comprises a thermal control assembly (9), wherein the thermal control assembly (9) comprises a heating section (91), a transition section (92) and a lead-out section (93) connected in sequence, wherein the heating section (91) is sleeved on the outer wall of the catalyst bed (81) and is used for temperature control and monitoring of the catalyst bed (81), wherein the transition section (92) is located on a side of the heating section (91) close to the frame (4), wherein the transition section (92) is made of a thin-walled variable diameter armored tube, and wherein the lead-out section (93) comprises a high-temperature epoxy resin material and is used for sealing the terminal lead wire.

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