A method for designing an engine thrust chamber and an engine thrust chamber

Abstract

The present invention discloses an engine thrust chamber design method and an engine thrust chamber, belonging to the field of aerospace technology, including determining the number and layout of engines according to the design requirements of the attitude control power system at the end of the repair, and determining the performance parameters of the engines; matching the control requirements of the attitude control power system at the end of the repair with the performance parameters of the engines, determining the flow range of the propellant entering the engine and the pressure range of the engine thrust chamber; determining the injection scheme and profile scheme of the engine thrust chamber; performing simulation calculation on the engine thrust chamber to verify whether the performance of the engine thrust chamber meets the requirements. The present invention can meet the overall requirements of the attitude control power system at the end of the repair, and provide sufficient power for the attitude control power system at the end of the repair.

Description

Technical Field

[0001] The invention relates to the field of aerospace technology, and in particular to an engine thrust chamber design method and an engine thrust chamber. Background Art

[0002] As a tool for transporting spacecraft, the rocket's final stage will separate from the spacecraft after sending the spacecraft into the predetermined orbit. Whether it is a satellite or other spacecraft, the interior is composed of very precise electronic components. The rocket's final stage not only sends them to the predetermined orbit, but also adjusts the satellite's posture.

[0003] The final attitude control power system provides power for the final stage of the rocket, ensuring that the final stage of the rocket and the payload can enter the preset orbit. The structure of the engine thrust chamber, as the direct power output structure of the final attitude control power system, determines the controllability of the final attitude control power system and ensures the accuracy of the payload entering the orbit.

[0004] The existing engine thrust chamber design does not take into account the overall requirements of the terminal attitude control power system, and is not fully matched with the overall terminal attitude control power system.

[0005] In view of this, it is necessary to provide a new technical solution to solve the above problems. Summary of the invention

[0006] In order to solve the above technical problems, the present application provides an engine thrust chamber design method and an engine thrust chamber, which can meet the overall requirements of the terminal attitude control power system and provide sufficient power for the terminal attitude control power system.

[0007] A method for designing an engine thrust chamber, comprising:

[0008] Determine the number and layout of engines according to the design requirements of the final attitude control power system, and determine the performance parameters of the engines;

[0009] Match the control requirements of the attitude control power system at the end of the repair with the performance parameters of the engine, and determine the flow range of the propellant entering the engine and the pressure range of the engine thrust chamber;

[0010] Determine the injection scheme and profile scheme of the engine thrust chamber;

[0011] Perform simulation calculations on the engine thrust chamber to verify whether the performance of the engine thrust chamber meets the requirements.

[0012] Preferably, the design requirements of the terminal attitude control power system include: standard design conditions of the terminal attitude control power system, type of propellant, total system mass, engine operating mode and engine performance parameters.

[0013] Preferably, the engine comprises a first thrust engine and a second thrust engine;

[0014] Among them, the design thrust of the first thrust engine is 50N; there are 4 first thrust engines;

[0015] The design thrust of the second thrust engine is 700N; there are 12 second thrust engines.

[0016] Preferably, the engine performance parameters include: vacuum specific impulse, design thrust, thrust deviation, paired working thrust deviation, mixture ratio deviation, aftereffect impulse deviation, longest continuous working time, shortest working time, cumulative working time, longest restart interval time, shortest interval time, cumulative number of working times, thrust line offset, thrust line deflection, start-up asynchrony, shutdown asynchrony, and the time from the engine receiving a start command to reaching 90% of the rated thrust and the time from the engine receiving a shutdown command to reaching 10% of the rated thrust.

[0017] Preferably, the engine thrust chamber pressure satisfies the following conditions:

[0018]

[0019] Where P c Δp is the steady-state working chamber pressure of the thrust chamber; v is the flow resistance of the propellant control valve group; ΔP h is the pressure drop of the pressure regulating orifice; P * is the inlet pressure of the propellant control valve group.

[0020] Preferably, matching the control requirements of the attitude control power system with the performance parameters of the engine includes:

[0021] Determine the number of engines to be started according to the control requirements of the attitude control power system at the end of the repair;

[0022] Calculate the pressure relief valve outlet pressure and engine thrust chamber inlet pressure when the current number of engines are turned on;

[0023] The thrust of the thrust chamber is determined based on the engine thrust chamber inlet pressure and the thrust deviation is calculated.

[0024] Preferably, in determining the number of engines to be turned on according to the control requirements of the attitude control power system, the attitude control power system has a first control state, a second control state, a third control state and a fourth control state;

[0025] Among them, in the first control state, the two 50N engines are working, and the inlet pressure of the engine thrust chamber is at the highest state;

[0026] In the second control state, four 700N engines and two 50N engines work simultaneously, and the engine thrust chamber inlet pressure fluctuates around the design state;

[0027] In the third control state, six 700N engines and two 50N engines work simultaneously, and the engine thrust chamber inlet pressure is lower than the design state;

[0028] In the fourth control state, eight 700N engines and two 50N engines work simultaneously, and the engine thrust chamber inlet pressure is at the lowest state.

[0029] According to another aspect of the present invention, there is also included an engine thrust chamber designed by the engine thrust chamber design method, comprising:

[0030] Combustion chamber;

[0031] Nozzle; the nozzle is fixedly arranged at the rear end of the combustion chamber and communicated with the combustion chamber; the nozzle includes an expansion section and an outlet; a throat is provided between the combustion chamber and the nozzle;

[0032] A control valve group; the control valve group is connected to the combustion chamber and is used to control the propellant to enter the combustion chamber;

[0033] A nozzle, the nozzle being connected to the control valve group through a pipeline and being used for atomizing the propellant;

[0034] A throttling orifice plate; the throttling orifice plate is connected to the outlet of the control valve group and is used to throttle the propellant.

[0035] Preferably, the injector of the 50N thrust chamber includes a head base, an oxidizer swirler and a fuel swirler; the head base includes a coaxially nested fuel nozzle and an oxidizer nozzle; the oxidizer nozzle surrounds the inside of the fuel nozzle; the fuel nozzle inlet end is connected to the fuel swirler; the oxidizer nozzle inlet end is connected to the oxidizer swirler.

[0036] Preferably, the 700N thrust chamber adopts a DC mutual impact injection form, including an injector head; the injector head includes an injection disk and a guide cavity; the injection disk includes 36 pairs of injection holes; the unit thrust of each pair of injection holes is 19.447N; the injection holes are arranged on inner and outer rings, wherein the inner ring has 12 pairs of injection holes and the outer ring has 24 pairs of injection holes.

[0037] Compared with the prior art, this application has at least the following beneficial effects:

[0038] 1. The present invention can meet the overall requirements of the terminal attitude control power system and provide sufficient power for the terminal attitude control power system.

[0039] 2. The present invention verifies the performance of the engine thrust chamber by performing simulation calculation on the engine thrust chamber, thereby ensuring the reliability of the engine thrust chamber performance.

[0040] 3. The present invention matches the control requirements of the terminal attitude control power system with the performance parameters of the engine, thereby determining the flow range of the propellant entering the engine and the pressure range of the engine thrust chamber, which can effectively achieve the matching of the engine thrust chamber and the terminal attitude control power system. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] Hereinafter, some specific embodiments of the present invention will be described in detail in an exemplary and non-limiting manner with reference to the accompanying drawings. The same reference numerals in the accompanying drawings indicate the same or similar components or parts. It should be understood by those skilled in the art that these drawings are not necessarily drawn to scale.

[0042] In the figure:

[0043] Figure 1 It is a schematic diagram of the overall process of the present invention;

[0044] Figure 2 This is a schematic diagram of the structural dimensions of the thrust chamber profile of a 50N engine;

[0045] Figure 3 This is a schematic diagram of the structural dimensions of the thrust chamber of a 50N engine;

[0046] Figure 4 This is a schematic diagram of the structural principle of the 50N engine;

[0047] Figure 5 This is a schematic diagram of the structural principle of the 700N engine;

[0048] Figure 6 This is the principle diagram of the 50N engine thrust chamber injector;

[0049] Figure 7 It is a schematic diagram of the structure of the thrust chamber injector of the 50N engine;

[0050] Figure 8 This is a schematic diagram of the structure of the 700N engine thrust chamber injector. DETAILED DESCRIPTION

[0051] In order to make the purpose, technical solution and advantages of the present application clearer, the technical solution of the present application will be clearly and completely described below in combination with the specific embodiments of the present application and the corresponding drawings. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present application.

[0052] like Figure 1 As shown, a method for designing an engine thrust chamber includes:

[0053] Step S1: determine the number and layout of engines according to the design requirements of the attitude control power system, and determine the performance parameters of the engines.

[0054] Among them, the design requirements of the terminal attitude control power system include: standard design conditions of the terminal attitude control power system, propellant type, total system mass, engine operating mode and engine performance parameters.

[0055] The standard design conditions for the final attitude control power system are: temperature of 20°C and vacuum environment.

[0056] The terminal attitude control power system uses liquid bipropellant, the oxidizer is green nitrogen tetroxide (MON-3), and the fuel is methyl hydrazine (MMH). The dry mass of the power system is no more than 100kg, the wet mass (including nitrogen and fuel) is no more than 355kg, and the effective working propellant is no less than 230kg.

[0057] The engine includes a first thrust engine and a second thrust engine, both of which have two working modes: steady-state continuous and pulse. Among them, the design thrust of the first thrust engine is 50N; there are 4 first thrust engines. The design thrust of the second thrust engine is 700N; there are 12 second thrust engines.

[0058] Among them, the engine performance parameters include: vacuum specific impulse, design thrust, thrust deviation, paired working thrust deviation, mixture ratio deviation, aftereffect impulse deviation, longest continuous working time, shortest working time, cumulative working time, longest restart interval, shortest interval, cumulative number of working times, thrust line offset, thrust line skew, start-up asynchrony, shutdown asynchrony, and the time from the engine receiving the start command to reaching 90% of the rated thrust and the time from the engine receiving the shutdown command to reaching 10% of the rated thrust.

[0059] In this embodiment, the engine performance parameters are shown in Table 1.

[0060] Table 1 Engine performance parameters

[0061]

[0062]

[0063] In the table, t90 is the time from when the engine receives the start command to when it reaches 90% of the rated thrust; t10 is the time from when the engine receives the shutdown command to when it reaches 10% of the rated thrust.

[0064] Step S2: Match the control requirements of the final attitude control power system with the performance parameters of the engine to determine the flow range of the propellant entering the engine and the pressure range of the engine thrust chamber.

[0065] Specifically, the control requirements of the final attitude control power system are matched with the performance parameters of the engine, including:

[0066] Step S21: Determine the number of engines to be started according to the control requirements of the attitude control power system.

[0067] Step S22, calculating the pressure reducing valve outlet pressure and the engine thrust chamber inlet pressure when the current number of engines are turned on.

[0068] Step S23, determining the thrust of the thrust chamber according to the inlet pressure of the thrust chamber of the engine and calculating the thrust deviation.

[0069] In this embodiment, the final attitude control power system has a first control state, a second control state, a third control state and a fourth control state.

[0070] The extruded gas and propellant required for the operation of the terminal attitude control power system are pre-filled in the gas cylinder and the tank, respectively. The downstream of the gas cylinder is a high-pressure electric explosion valve, and the front and rear of the tank are low-pressure electric explosion valves. At a certain moment, the control system energizes the electric explosion tube in the high-pressure electric explosion valve behind the gas cylinder, opens the gas cylinder, and then energizes to open the low-pressure electric explosion valves before and after the tank, so that the gas and liquid circuits are open.

[0071] The high-pressure nitrogen in the gas cylinder flows to the pressure reducing valve through the high-pressure electric explosion valve. After flowing out, it is in a low-pressure state with stable pressure. It is used to pressurize the storage tank and provide an air source for the electric gas valve to open.

[0072] The pressure reducing valve is located downstream of the gas cylinder and controls the pressure entering the fuel tank and the oxidizer tank, which in turn affects the pressure of the propellant supplied to the engine by the fuel tank and the oxidizer tank.

[0073] According to the design requirements, the nominal value of the pressure reducing valve outlet pressure should be 3.5MPa, and the volume flow rate range of nitrogen at the pressure reducing valve outlet is 0.026L / s~2.1L / s. As the flow rate increases, the pressure reducing valve outlet pressure decreases, and the pressure reducing valve outlet pressure control accuracy shall not be less than ±5%.

[0074] In the first control state, the two 50N engines are working, the pressure reducing valve outlet pressure is high, and the flow resistance of the pipeline system is small. The engine thrust chamber inlet pressure is at the highest state, which will cause the engine chamber pressure to increase. After calculation, it is found that the thrust in this state deviates from the design state by +12.62%.

[0075] Table 2 Parameters of attitude control power system in the first state

[0076]

[0077]

[0078] In the second control state, four 700N engines and two 50N engines work simultaneously, the pressure reducing valve outlet pressure is basically in the typical design state, the engine thrust chamber inlet pressure and thrust chamber pressure are close to the design state, and fluctuate within the design state. After calculation, it is found that in this state, the thrust of the 700N engine deviates from the design state by +4.60%, and the thrust of the 50N engine deviates from the design state by +4.33%.

[0079] Table 3 Parameters of attitude control power system in the second state

[0080]

[0081] In the third control state, six 700N engines and two 50N engines work simultaneously, the pressure reducing valve outlet pressure is lower than the typical design state, the engine thrust chamber inlet pressure and thrust chamber pressure are also lower than the design state. After calculation, it is found that in this state, the thrust of the 700N engine deviates from the design state by -3.54%, and the thrust of the 50N engine deviates from the design state by -4.09%.

[0082] Table 4 Parameters of the attitude control power system at the end of the third state

[0083]

[0084] In the fourth control state, eight 700N engines and two 50N engines work at the same time, the pressure reducing valve outlet pressure is at the lowest state, and the flow resistance of the pipeline system is the largest, which will cause the engine thrust chamber inlet pressure and thrust chamber pressure to be significantly lower than the design state and be at the lowest state. After calculation, it is found that in this state, the thrust of the 700N engine deviates from the design state by -12.99%, and the thrust of the 50N engine deviates from the design state by -13.05%.

[0085] Table 5 Parameters of the attitude control power system at the end of the fourth state

[0086]

[0087] Before the engine works, the oxidizer and fuel are filled in front of the control valve group through their respective pipelines under the action of tank pressure. When the working instruction is received, the control valve group opens synchronously, and the propellant enters the combustion chamber through the oxidizer nozzle and the fuel nozzle respectively. After atomization, mixing and combustion, high-temperature combustion gas is formed, which is then discharged from the nozzle at an accelerated speed to generate the thrust required for the attitude control of the upper stage.

[0088] Since the system flow resistance is basically proportional to the square of the flow rate, when the pressure reducing valve outlet pressure is constant, the engine inlet pressure varies greatly under high and low flow conditions, and the engine needs to adapt to a wide pressure range. It is necessary to reasonably match the pressure reducing valve outlet pressure, adjust the orifice pressure drop and other parameters, and reasonably allocate the system parameters in combination with the system cold test data, engine test data and the actual use needs of the control profession.

[0089] The engine pressure must satisfy the following relationship:

[0090] P c +ΔP inj +ΔP v +ΔP h ≤P * ;

[0091] Where P c ΔP is the steady-state working chamber pressure of the thrust chamber; inj is the injection pressure drop; Δp v is the flow resistance of the propellant control valve group; ΔP h is the pressure drop of the pressure regulating orifice; P * is the inlet pressure of the propellant control valve group.

[0092] Among them, the injection pressure drop and the steady-state working chamber pressure of the thrust chamber have the following relationship:

[0093] ΔP inj =0.35P c ;

[0094] In addition, according to the design requirements and the actual working conditions of the corresponding components, the parameters are set as follows:

[0095] ΔP v =0.2MPa;

[0096] ΔP h =0.1MPa;

[0097] P * =3MP a ;

[0098] Therefore, after derivation, the requirement for the steady-state working chamber pressure of the thrust chamber is as follows:

[0099]

[0100] Substituting the above parameters into the formula for calculation, the steady-state working chamber pressure of the thrust chamber is 2MPa.

[0101] Step S3, determining the injection scheme and profile scheme of the engine thrust chamber.

[0102] When determining the injection scheme, for the 50N thrust chamber, due to the small diameter of the thrust chamber combustion chamber, a DC mutual impact injection scheme is adopted. After the propellant is sputtered, a local oxygen-rich combustion zone is easily formed on the wall of the combustion chamber, which is not conducive to the structural thermal protection of the thrust chamber.

[0103] The fuel nozzle and the oxidizer nozzle are coaxially nested, with the fuel nozzle inside and the oxidizer nozzle outside. The propellants are accelerated by the swirl flow in their respective swirl chambers and then ejected, forming an atomized cone under the action of centrifugal force. After the oxidizer and fuel come into contact, they are quickly mixed and atomized under the action of the atomization kinetic energy, and then ignite and burn, and the engine starts to work.

[0104] For the 700N thrust chamber, a DC mutual impact injection design is adopted. By comparing and referring to the existing models, the diameter of the thrust chamber combustion chamber is determined to be 39mm and the characteristic length is about 310mm, achieving a combustion efficiency ≥0.92.

[0105] When determining the profile scheme, according to the gas residence time requirements (generally 1ms to 3ms) and the design experience of the combustion chamber of the same type of small thrust engine, the characteristic length L of the 50N engine combustion chamber is selected as 0.3m. The throat section diameter Dt is 4.24mm. The combustion chamber diameter Dc is calculated by the flow density method to obtain 12.5mm. The nozzle inlet arc radius Rct is selected as 20mm, the upstream throat section radius Rc1 is selected as 0.83Dt, and the downstream throat section Rc2 is selected as 0.41Dt.

[0106] In order to achieve the highest nozzle efficiency and maximum thrust, the 50N engine nozzle diffuser is designed using the characteristic line method. In this embodiment, the design point height is 120km. According to the design point height and the thrust chamber geometric installation size constraints, the nozzle expansion ratio ε=81, the nozzle diffuser expansion half angle βk=32°, and the outlet expansion angle βe=8° are selected. Based on the above calculation steps, the design obtains the 50N engine thrust chamber profile dimensions as follows: Figure 2 and Figure 3 shown.

[0107] The expansion section of the 700N thrust chamber combustion chamber adopts the maximum thrust nozzle shape design. The main design parameters of the nozzle are shown in Table 6.

[0108] Table 6 Main design parameters of nozzle

[0109]

[0110]

[0111] Step S4: Perform simulation calculation on the engine thrust chamber to verify whether the performance of the engine thrust chamber meets the requirements.

[0112] Relying on simulation software to build a model and simulate and verify the flow trajectory of the fuel in the nozzle cavity, it can be concluded from the flow field pressure cloud map of the fuel nozzle of the 50N engine that the nozzle inlet pressure is 0.653MPa, which is basically consistent with the design value of 0.65MPa.

[0113] At the same time, the overall, tangential and axial velocity cloud diagrams of the flow length in the fuel nozzle are used to verify that the velocity distribution conforms to the actual physical process of liquid propulsion flowing in the flow field inside the centrifugal nozzle.

[0114] Through simulation analysis of the liquid phase distribution cloud map in the inner cavity of the fuel nozzle, it was determined that under the action of swirl centrifugation, a conical "cavitation" was formed in the center of the nozzle outlet section.

[0115] For small thrust liquid rocket engines, common thermal protection schemes for the body include regenerative cooling, liquid film cooling, passive ablation cooling, etc. In the field of attitude control engines, the most widely used is the radiation cooling scheme. Referring to the current technical schemes of similar models, the thermal protection scheme for the thrust chamber body is selected as liquid film cooling. The body material is Nb521, and the inner wall is coated with r625 coating with a coating thickness of 80μm-120μm, which can play a role in anti-oxidation and ablation.

[0116] The body stress was simulated and calculated using structural strength simulation software. When the working pressure of the inner cavity of the thrust chamber was 2MPa, the stress distribution under the conditions of 5MPa and 10MPa pressure in the thrust chamber was verified. The maximum structural stress of the thrust chamber shell under the two working conditions was 29.57MPa and 59.17MPa, respectively, which is much lower than the tensile strength of Nb521 material at 1600℃ (σb-1600≥150MPa).

[0117] In addition, the stress conditions of the injector under various conditions were simulated, including a 4.1MPa hydraulic strength test, a 2.7MPa inner cavity starting hydraulic impact hydraulic strength test, and a stress and deformation analysis of the hydraulic chamber pressure of 2.7MPa in the inner cavity and 2.0MPa on the panel when the engine is working.

[0118] The maximum stress point stresses of the head in three cases are calculated to be 38.4MPa, 67.6MPa and 36.6MPa respectively, and the corresponding safety factors are 23.2, 13.16 (tensile strength 890MPa in cold state) and 15.57 (tensile strength 570MPa at 400℃). At the same time, the deformation under various working conditions is very small, which proves that the structural design of the 700N injector has sufficient safety margin.

[0119] After the research and study in the scheme stage, the technical indicators that the power system can currently achieve are shown in Table 7. Except for the engine thrust deviation and flight reliability indicators that do not meet the overall requirements, other indicators meet the overall requirements, and test verification is planned to be carried out in the prototype stage.

[0120] Table 7 Main indicators of the final attitude control power system are satisfied

[0121]

[0122]

[0123] An engine thrust chamber designed by the engine thrust chamber design method comprises:

[0124] Combustion chamber;

[0125] Nozzle; the nozzle is fixedly arranged at the rear end of the combustion chamber and communicated with the combustion chamber; the nozzle includes an expansion section and an outlet; a throat is provided between the combustion chamber and the nozzle;

[0126] A control valve group; the control valve group is connected to the combustion chamber and is used to control the propellant to enter the combustion chamber;

[0127] A nozzle, the nozzle being connected to the control valve group through a pipeline and being used for atomizing the propellant;

[0128] A throttling orifice plate; the throttling orifice plate is connected to the outlet of the control valve group and is used to throttle the propellant.

[0129] The engine in this embodiment uses a liquid bipropellant, the oxidizer is green nitrogen tetroxide (MON-3), the fuel is methyl hydrazine (MMH), and the theoretical mixing ratio is 1.65.

[0130] Among them, the structural principle of the 50N engine is as follows Figure 4 As shown, its control valve group adopts solenoid valve.

[0131] The structural principle of the 700N engine is as follows Figure 5 As shown, its control valve group adopts electric gas valve.

[0132] When working, the oxidant outlet and fuel outlet of the solenoid valve or electric gas valve are both equipped with throttle orifice plates. The two propellants enter the thrust chamber for atomization and mixing, and burn and decompose into high-temperature gas. The high-temperature gas is accelerated and ejected through the nozzle to generate thrust. The control system issues a power-off command for the solenoid valve or electric gas valve, and the solenoid valve or electric gas valve is closed, and the engine stops working.

[0133] like Figure 6 and Figure 7As shown, the injector of the thrust chamber of the 50N engine includes a head base, an oxidizer swirler and a fuel swirler; the head base includes a coaxially nested fuel nozzle and an oxidizer nozzle; the oxidizer nozzle surrounds the inside of the fuel nozzle; the fuel nozzle inlet end is connected to the fuel swirler; the oxidizer nozzle inlet end is connected to the oxidizer swirler. The main design parameters of the 50N engine injector are shown in Table 8.

[0134] Table 850N engine injector main design parameters

[0135]

[0136]

[0137] like Figure 8 As shown, the 700N thrust chamber adopts a DC mutual impact injection form, including an injector head; the injector head includes an injection disk and a guide cavity; the injection disk includes 36 pairs of injection holes; the unit thrust of each pair of injection holes is 19.447N; the injection holes are arranged on inner and outer rings, wherein the inner ring has 12 pairs of injection holes and the outer ring has 24 pairs of injection holes.

[0138] Specifically, the 700N thrust chamber adopts a DC mutual impact injection design. By comparing and referring to existing models, the diameter of the thrust chamber combustion chamber is determined to be 39mm, the characteristic length is about 310mm, and the combustion efficiency is ≥0.92. Based on the reference to the unit thrust of the existing thrust chamber, the number of injection pairs of this thrust chamber is determined to be 36 pairs (12+24), and the unit thrust is 19.447N / pair.

[0139] The total number of injection pairs in the 700N thrust chamber is 36, which are arranged in two circles, with 12 pairs in the inner circle and 24 pairs in the outer circle. In order to achieve uniform distribution of flow intensity and mixing ratio, the impact points of each circle should be distributed as evenly as possible along the diameter of the combustion chamber.

[0140] The head of the 700N thrust chamber injector is composed of an injector disc and a guide cavity. In view of the nearly 500 cumulative working times and the shortest 40ms switching pulse requirements of the thrust chamber, more stringent requirements are put forward for the waterproofing of the guide cavity at the injector inlet. The overall assembly of the injector is mainly connected by laser and electron beam welding, assisted by argon arc welding, which effectively reduces deformation and improves structural reliability.

[0141] The oxidant is in the central cavity, and the nozzles are evenly arranged in two circles, the inner and outer circles. The apertures and aspect ratios of the two circles of nozzles are set to be consistent to ensure uniform discharge of the oxidant; the fuel cavity is concentrated on the outside, and the diverter ring adopts a novel eccentric circle structure design to ensure uniform discharge of the nozzles at the proximal and distal ends of the liquid collecting cavity.

[0142] It should be noted that the terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments according to the present application. As used herein, unless the context clearly indicates otherwise, the singular form is also intended to include the plural form. In addition, it should be understood that when the terms "comprise" and / or "include" are used in this specification, it indicates the presence of features, steps, operations, devices, components and / or combinations thereof.

[0143] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application and the above-mentioned drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the numbers used in this way can be interchanged where appropriate, so that the embodiments of the present application described herein can be implemented in an order other than those illustrated or described herein.

[0144] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for designing an engine thrust chamber, characterized in that: include: Determine the number and layout of engines according to the design requirements of the final attitude control power system, and determine the performance parameters of the engines; Match the control requirements of the attitude control power system at the end of the repair with the performance parameters of the engine, and determine the flow range of the propellant entering the engine and the pressure range of the engine thrust chamber; Determine the injection scheme and profile scheme of the engine thrust chamber; Perform simulation calculations on the engine thrust chamber to verify whether the performance of the engine thrust chamber meets the requirements; The engine includes a first thrust engine and a second thrust engine; Among them, the design thrust of the first thrust engine is 50N; there are 4 first thrust engines; The design thrust of the second thrust engine is 700N; there are 12 second thrust engines; The engine performance parameters include: vacuum specific impulse, design thrust, thrust deviation, paired working thrust deviation, mixture ratio deviation, aftereffect impulse deviation, one-time longest continuous working time, shortest working time, cumulative working time, longest restart interval time, shortest interval time, cumulative number of working times, thrust line offset, thrust line deflection, start asynchronicity, shutdown asynchronicity, and the time from when the engine receives a start command to when it reaches 90% of the rated thrust and the time from when the engine receives a shutdown command to when it reaches 10% of the rated thrust; According to the control requirements of the final attitude control power system, the number of engines to be turned on is determined, wherein the final attitude control power system has a first control state, a second control state, a third control state and a fourth control state; Among them, in the first control state, the two 50N engines are working, and the inlet pressure of the engine thrust chamber is at the highest state; In the second control state, four 700N engines and two 50N engines work simultaneously, and the engine thrust chamber inlet pressure fluctuates around the design state; In the third control state, six 700N engines and two 50N engines work simultaneously, and the engine thrust chamber inlet pressure is lower than the design state; In the fourth control state, eight 700N engines and two 50N engines work simultaneously, and the engine thrust chamber inlet pressure is at the lowest state; The matching of the control requirements of the final attitude control power system with the performance parameters of the engine includes: Determine the number of engines to be started according to the control requirements of the attitude control power system at the end of the repair; Calculate the pressure relief valve outlet pressure and engine thrust chamber inlet pressure when the current number of engines are turned on; The thrust of the thrust chamber is determined based on the engine thrust chamber inlet pressure and the thrust deviation is calculated.

2. The engine thrust chamber design method according to claim 1, characterized in that: The design requirements of the terminal attitude control power system include: standard design conditions for the terminal attitude control power system, propellant type, total system mass, engine operating mode and engine performance parameters.

3. The engine thrust chamber design method according to claim 2, characterized in that: The thrust chamber pressure of the engine meets the following conditions: ； In the formula, is the steady-state working chamber pressure of the thrust chamber; is the flow resistance of the propellant control valve group; is the pressure drop of the pressure regulating orifice; is the inlet pressure of the propellant control valve group.

4. An engine thrust chamber designed by the engine thrust chamber design method according to any one of claims 1 to 3, characterized in that: include: Combustion chamber; Nozzle; The nozzle is fixedly arranged at the rear end of the combustion chamber and communicated with the combustion chamber; the nozzle includes an expansion section and an outlet; A throat is provided between the combustion chamber and the nozzle; A control valve group; the control valve group is connected to the combustion chamber and is used to control the propellant to enter the combustion chamber; A nozzle, the nozzle being connected to the control valve group through a pipeline and being used for atomizing the propellant; A throttling orifice plate; the throttling orifice plate is connected to the outlet of the control valve group and is used to throttle the propellant.

5. The engine thrust chamber design method according to claim 4, characterized in that: The injector of the 50N thrust chamber includes a head base, an oxidizer swirler and a fuel swirler; the head base includes a coaxially nested fuel nozzle and an oxidizer nozzle; the oxidizer nozzle surrounds the inside of the fuel nozzle; the fuel nozzle inlet end is connected to the fuel swirler; the oxidizer nozzle inlet end is connected to the oxidizer swirler.

6. The engine thrust chamber design method according to claim 4, characterized in that: The 700N thrust chamber adopts a DC mutual impact injection form, including an injector head; the injector head includes an injection disk and a guide cavity; the injection disk includes 36 pairs of injection holes; the unit thrust of each pair of injection holes is 19.447N; the injection holes are arranged on inner and outer rings, wherein the inner ring has 12 pairs of injection holes and the outer ring has 24 pairs of injection holes.

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