CN117932779B - A method for designing an engine thrust chamber and an engine thrust chamber - Google Patents
A method for designing an engine thrust chamber and an engine thrust chamber Download PDFInfo
- Publication number
- CN117932779B CN117932779B CN202410001633.8A CN202410001633A CN117932779B CN 117932779 B CN117932779 B CN 117932779B CN 202410001633 A CN202410001633 A CN 202410001633A CN 117932779 B CN117932779 B CN 117932779B
- Authority
- CN
- China
- Prior art keywords
- engine
- thrust
- chamber
- thrust chamber
- engines
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000000034 method Methods 0.000 title claims abstract description 18
- 238000013461 design Methods 0.000 claims abstract description 63
- 238000002347 injection Methods 0.000 claims abstract description 37
- 239000007924 injection Substances 0.000 claims abstract description 37
- 239000003380 propellant Substances 0.000 claims abstract description 32
- 238000004364 calculation method Methods 0.000 claims abstract description 9
- 238000004088 simulation Methods 0.000 claims abstract description 7
- 238000002485 combustion reaction Methods 0.000 claims description 35
- 239000000446 fuel Substances 0.000 claims description 31
- 239000007800 oxidant agent Substances 0.000 claims description 29
- 230000001105 regulatory effect Effects 0.000 claims description 3
- 230000001186 cumulative effect Effects 0.000 claims 2
- 239000000203 mixture Substances 0.000 claims 1
- 230000001590 oxidative effect Effects 0.000 description 23
- 239000007789 gas Substances 0.000 description 10
- 239000007788 liquid Substances 0.000 description 8
- 239000007921 spray Substances 0.000 description 7
- 238000004880 explosion Methods 0.000 description 6
- 238000002156 mixing Methods 0.000 description 6
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 description 5
- 230000009471 action Effects 0.000 description 5
- 238000001816 cooling Methods 0.000 description 5
- 238000003860 storage Methods 0.000 description 5
- 238000012360 testing method Methods 0.000 description 5
- 238000000889 atomisation Methods 0.000 description 4
- 230000001276 controlling effect Effects 0.000 description 4
- 238000012937 correction Methods 0.000 description 4
- 238000010586 diagram Methods 0.000 description 4
- WFPZPJSADLPSON-UHFFFAOYSA-N dinitrogen tetraoxide Chemical compound [O-][N+](=O)[N+]([O-])=O WFPZPJSADLPSON-UHFFFAOYSA-N 0.000 description 4
- 238000009826 distribution Methods 0.000 description 4
- HDZGCSFEDULWCS-UHFFFAOYSA-N monomethylhydrazine Chemical compound CNN HDZGCSFEDULWCS-UHFFFAOYSA-N 0.000 description 4
- 239000000243 solution Substances 0.000 description 3
- XKRFYHLGVUSROY-UHFFFAOYSA-N Argon Chemical compound [Ar] XKRFYHLGVUSROY-UHFFFAOYSA-N 0.000 description 2
- 239000011248 coating agent Substances 0.000 description 2
- 238000000576 coating method Methods 0.000 description 2
- 239000002737 fuel gas Substances 0.000 description 2
- 238000012423 maintenance Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 229910052757 nitrogen Inorganic materials 0.000 description 2
- 238000012795 verification Methods 0.000 description 2
- 238000003466 welding Methods 0.000 description 2
- 238000002679 ablation Methods 0.000 description 1
- 230000003064 anti-oxidating effect Effects 0.000 description 1
- 229910052786 argon Inorganic materials 0.000 description 1
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000005119 centrifugation Methods 0.000 description 1
- 238000009792 diffusion process Methods 0.000 description 1
- 229910001873 dinitrogen Inorganic materials 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000010894 electron beam technology Methods 0.000 description 1
- 238000001125 extrusion Methods 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 239000007791 liquid phase Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 239000001301 oxygen Substances 0.000 description 1
- 229910052760 oxygen Inorganic materials 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 230000005855 radiation Effects 0.000 description 1
- 230000001172 regenerating effect Effects 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000009827 uniform distribution Methods 0.000 description 1
- 238000011144 upstream manufacturing Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/10—Geometric CAD
- G06F30/15—Vehicle, aircraft or watercraft design
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/96—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by specially adapted arrangements for testing or measuring
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
- G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Theoretical Computer Science (AREA)
- Geometry (AREA)
- General Engineering & Computer Science (AREA)
- Mathematical Analysis (AREA)
- Evolutionary Computation (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Mathematical Optimization (AREA)
- Computer Hardware Design (AREA)
- Pure & Applied Mathematics (AREA)
- Mathematical Physics (AREA)
- Computing Systems (AREA)
- Algebra (AREA)
- Fluid Mechanics (AREA)
- Automation & Control Theory (AREA)
- Aviation & Aerospace Engineering (AREA)
- Computational Mathematics (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
- Fuel-Injection Apparatus (AREA)
Abstract
本发明公开了一种发动机推力室设计方法及发动机推力室，属于航天技术领域，包括根据末修姿控动力系统的设计要求确定发动机的数量和布局方式，并确定发动机的性能参数；将末修姿控动力系统的控制要求与发动机的性能参数进行匹配，确定进入发动机推进剂的流量范围以及发动机推力室室压范围；确定发动机推力室的喷注方案和型面方案；对发动机推力室进行仿真计算，验证发动机推力室的性能是否满足要求。本发明能够满足末修姿控动力系统的整体需求，为末修姿控动力系统提供足够的动力。
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
The invention relates to the technical field of aerospace, in particular to an engine thrust chamber design method and an engine thrust chamber.
Background
The rocket is used as a tool for transporting the spacecraft, and after the spacecraft is sent into a preset orbit, the rocket final stage is separated from the spacecraft. The rocket final stage is composed of very precise electronic components, whether satellites or other spacecrafts, and is used for not only delivering the satellites to a preset orbit, but also adjusting the satellites to a proper posture.
The final-repair attitude control power system provides power for the rocket final stage, and ensures that the rocket final stage and load can enter a preset orbit. The structure of the engine thrust chamber is used as a direct power output structure of the final-repair attitude control power system, so that the action controllability of the final-repair attitude control power system is determined, and the accuracy of load track entering is ensured.
The existing engine thrust chamber design scheme does not start from the overall requirement of the final-repair attitude control power system, and is not completely matched with the overall requirement of the final-repair attitude control power system.
In view of the foregoing, it is necessary to provide a new solution to the above-mentioned problems.
Disclosure of Invention
In order to solve the technical problems, the application provides an engine thrust chamber design method and an engine thrust chamber, which can meet the overall requirement of a final-repair gesture control power system and provide sufficient power for the final-repair gesture control power system.
An engine thrust chamber design method comprising:
Determining the number and layout modes of the engines according to the design requirements of the final-repair attitude control power system, and determining the performance parameters of the engines;
matching the control requirement of the final-repair attitude control power system with the performance parameters of the engine, and determining the flow range of the propellant entering the engine and the chamber pressure range of the thrust chamber of the engine;
determining an injection scheme and a profile scheme of an engine thrust chamber;
and (3) performing simulation calculation on the engine thrust chamber, and verifying whether the performance of the engine thrust chamber meets the requirement.
Preferably, the design requirements of the final-repair attitude control power system comprise design standard conditions, propellant types, total mass of the system, working mode of the engine and performance parameters of the engine.
Preferably, the engine comprises a first thrust engine and a second thrust engine;
The design thrust of the first thrust engine is 50N, and the first thrust engine is provided with 4 engines;
the second thrust engine was designed with a thrust of 700N and had 12 engines.
Preferably, the engine performance parameters include vacuum specific impulse, design thrust, thrust deviation, paired working thrust deviation, mixing ratio deviation, post-effect impulse deviation, one continuous longest working time, shortest working time, accumulated working time, longest restarting interval time, shortest interval time, accumulated working times, thrust line deviation, starting asynchronism, shutdown asynchronism, and time from when the engine receives a starting command to 90% of rated thrust and time from when the engine receives a shutdown command to 10% of rated thrust.
Preferably, the engine thrust chamber pressure satisfies the following condition:
Wherein P c is the steady-state working chamber pressure of the thrust chamber, deltap v is the flow resistance of the propellant control valve group, deltaP h is the pressure drop of the pressure regulating orifice plate, and P * is the inlet pressure of the propellant control valve group.
Preferably, the matching the control requirement of the final-repair attitude control power system with the performance parameter of the engine includes:
determining the starting quantity of the engine according to the control requirement of the final-repair attitude control power system;
Calculating the outlet pressure of the pressure reducing valve and the inlet pressure of the thrust chamber of the engine when the current number of engines are started;
and determining the thrust of the thrust chamber according to the inlet pressure of the thrust chamber of the engine and calculating the thrust deviation.
Preferably, the number of the started engines is determined according to the control requirement of the final-repair attitude control power system, wherein the final-repair attitude control power system is provided with a first control state, a second control state, a third control state and a fourth control state;
In the first control state, two 50N engines work, and the inlet pressure of the engine thrust chamber is in the highest state;
In a second control state, four 700N engines and two 50N engines work simultaneously, and the inlet pressure of an engine thrust chamber floats up and down in a design state;
in a third control state, six 700N engines and two 50N engines work simultaneously, and the inlet pressure of an engine thrust chamber is lower than the design state;
In the fourth control state, eight 700N engines and two 50N engines are operated simultaneously, with engine thrust chamber inlet pressure at a minimum.
According to another aspect of the present invention, there is also provided an engine thrust chamber designed by the engine thrust chamber design method, comprising:
A combustion chamber;
The combustion device comprises a combustion chamber, a spray pipe, a nozzle and a nozzle, wherein the spray pipe is fixedly arranged at the tail end of the combustion chamber and is communicated with the combustion chamber;
the control valve group is communicated with the combustion chamber and is used for controlling the propellant to enter the combustion chamber;
the nozzle is communicated with the control valve group through a pipeline and is used for atomizing the propellant;
and the throttle orifice plate is communicated with an outlet of the control valve group and used for throttling the propellant.
Preferably, the injector of the 50N thrust chamber comprises a head base body, an oxidant swirler and a fuel swirler, wherein the head base body comprises a fuel nozzle and an oxidant nozzle which are coaxially nested, the oxidant nozzle surrounds the inside of the fuel nozzle, the inlet end of the fuel nozzle is communicated with the fuel swirler, and the inlet end of the oxidant nozzle is communicated with the oxidant swirler.
Preferably, the 700N thrust chamber adopts a direct current mutual impact injection mode, and comprises an injector head, wherein the injector head comprises an injection disc and a flow guide cavity, the injection disc comprises 36 pairs of injection holes, the unit thrust of each pair of injection holes is 19.447N, the injection holes are arranged on an inner ring and an outer ring, 12 pairs of injection holes are formed in the inner ring, and 24 pairs of injection holes are formed in the outer ring.
Compared with the prior art, the application has at least the following beneficial effects:
1. The invention can meet the overall requirement of the power system for controlling the final maintenance posture and provides enough power for the power system for controlling the final maintenance posture.
2. According to the invention, the performance of the engine thrust chamber is verified by performing simulation calculation on the engine thrust chamber, so that the reliability of the performance of the engine thrust chamber is ensured.
3. According to the invention, the control requirement of the final-repair attitude control power system is matched with the performance parameters of the engine, so that the flow range of the propellant entering the engine and the chamber pressure range of the engine thrust chamber are determined, and the matching property of the engine thrust chamber and the final-repair attitude control power system can be effectively realized.
Drawings
Some specific embodiments of the invention will be described in detail hereinafter by way of example and not by way of limitation with reference to the accompanying drawings. The same reference numbers will be used throughout the drawings to refer to the same or like parts or portions. It will be appreciated by those skilled in the art that the drawings are not necessarily drawn to scale. Attached with
In the figure:
FIG. 1 is a schematic overall flow chart of the present invention;
FIG. 2 is a schematic illustration of the structural dimensions of a 50N engine thrust chamber profile;
FIG. 3 is a schematic view of the structural dimensions of a 50N engine thrust chamber body;
FIG. 4 is a schematic diagram of the structural principle of a 50N engine;
FIG. 5 is a schematic diagram of the structural principle of a 700N engine;
FIG. 6 is a schematic diagram of a 50N engine thrust chamber injector;
FIG. 7 is a schematic diagram of a 50N engine thrust chamber injector;
fig. 8 is a schematic structural view of a 700N engine thrust chamber injector.
Detailed Description
In order to make the objects, technical solutions and advantages of the present application more apparent, the technical solutions of the present application will be clearly and completely described below with reference to specific embodiments of the present application and corresponding drawings. It will be apparent that the described embodiments are only some, but not all, embodiments of the application. All other embodiments, which can be made by those skilled in the art based on the embodiments of the application without making any inventive effort, are intended to be within the scope of the application.
As shown in fig. 1, a method for designing an engine thrust chamber includes:
and S1, determining the number and layout modes of the engines according to the design requirements of the final-repair attitude control power system, and determining the performance parameters of the engines.
The design requirements of the final-repair attitude control power system comprise design standard conditions, propellant types, total mass of the system, working mode of the engine and performance parameters of the engine.
The design standard condition of the final-repair gesture control power system is that the temperature is 20 ℃ and the vacuum environment is adopted.
The final-repair gesture control power system adopts a liquid double-component propellant, the oxidant is green dinitrogen tetroxide (MON-3), and the fuel is methyl hydrazine (MMH). The dry mass of the power system is not more than 100kg, the wet mass (nitrogen gas and fuel) is not more than 355kg, and the effective working propellant quantity is not less than 230kg.
The engine comprises a first thrust engine and a second thrust engine, and the first thrust engine and the second thrust engine are in steady-state continuous and pulse operation modes. The design thrust of the first thrust engine is 50N, and the first thrust engine is provided with 4 engines. The second thrust engine was designed with a thrust of 700N and had 12 engines.
The engine performance parameters comprise vacuum specific impulse, design thrust, thrust deviation, paired working thrust deviation, mixing ratio deviation, aftereffect impulse deviation, continuous longest working time, shortest working time, accumulated working time, longest restarting interval time, shortest interval time, accumulated working times, thrust line deviation, starting asynchronism, shutdown asynchronism, time from when the engine receives a starting instruction to reach 90% of rated thrust and time from when the engine receives a shutdown instruction to reach 10% of rated thrust.
In this embodiment, the engine performance parameters are shown in table 1.
Table 1 engine performance parameter table
In the table, t90 is the time from when the engine receives the start command to when the engine reaches 90% of the rated thrust, and t10 is the time from when the engine receives the shutdown command to when the engine reaches 10% of the rated thrust.
And S2, matching the control requirement of the final-repair attitude control power system with the performance parameters of the engine, and determining the flow range of the propellant entering the engine and the chamber pressure range of the thrust chamber of the engine.
Specifically, matching the control requirement of the final-repair attitude control power system with the performance parameters of the engine comprises the following steps:
And S21, determining the starting quantity of the engine according to the control requirement of the final-repair attitude control power system.
Step S22, calculating the outlet pressure of the relief valve and the inlet pressure of the thrust chamber of the engine when the current number of engines are started.
And 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.
In this embodiment, the last-repair gesture control power system has a first control state, a second control state, a third control state, and a fourth control state.
The extrusion gas and propellant needed by the work of the final-repair gesture control power system are respectively filled in a gas cylinder and a storage tank in advance, a high-pressure explosion valve is arranged at the downstream of the gas cylinder, and the front and the rear of the storage tank are respectively provided with a low-pressure explosion valve. The control system electrifies the electric explosion tube in the high-pressure electric explosion valve behind the gas cylinder at a certain moment, opens the gas cylinder, and then electrifies the electric explosion valve in front of and behind the storage tank, and the gas circuit and the liquid circuit are respectively opened for flowing.
The high-pressure nitrogen in the gas cylinder flows to the pressure reducing valve through the high-pressure explosion valve, and is in a low-pressure state with stable pressure after flowing out, so as to pressurize the storage tank and provide an open air source for the electric air valve.
The pressure relief valve is downstream of the cylinder and controls the pressure into the fuel and oxidant reservoirs, thereby affecting the pressure at which the fuel and oxidant reservoirs supply the propellant to the engine.
According to design requirements, the nominal value of the outlet pressure of the pressure reducing valve is 3.5MPa, the volume flow of nitrogen at the outlet of the pressure reducing valve ranges from 0.026L/s to 2.1L/s, the outlet pressure of the pressure reducing valve is reduced along with the increase of the flow, and the control precision of the outlet pressure of the pressure reducing valve is not lower than +/-5%.
In the first control state, two 50N engines work, the outlet pressure of the pressure reducing valve is high, the flow resistance of the pipeline system is small, the inlet pressure of the engine thrust chamber is in the highest state, the pressure of the engine chamber can be increased, and the calculation shows that the thrust in the state deviates from the design state by +12.62%.
TABLE 2 non-correction gesture control Power System parameter Table in first State
In the second control state, four 700N engines and two 50N engines work simultaneously, the outlet pressure of the pressure reducing valve is basically in a typical design state, the inlet pressure of the engine thrust chamber and the pressure of the thrust chamber are close to the design state, the engine thrust chamber floats up and down in the design state, and the calculation shows that the deviation between the 700N engine thrust and the design state is +4.60% and the deviation between the 50N engine thrust and the design state is +4.33%.
TABLE 3 non-correction power system parameter table in second state
In a third control state, six 700N engines and two 50N engines are simultaneously operated, the outlet pressure of the pressure reducing valve is lower than that of a typical design state, the inlet pressure of an engine thrust chamber and the chamber pressure of the thrust chamber are also lower than that of the design state, and the calculation shows that the deviation between the 700N engine thrust and the design state in the state is-3.54%, and the deviation between the 50N engine thrust and the design state is-4.09%.
TABLE 4 non-correction gesture control power system parameter table in third state
In the fourth control state, the eight 700N engines and the two 50N engines work simultaneously, the outlet pressure of the pressure reducing valve is in the lowest state, and the flow resistance of the pipeline system is maximum, so that the inlet pressure of the engine thrust chamber and the chamber pressure of the thrust chamber are greatly lower than the design state, and in the lowest state, the 700N engine thrust and the design state deviate by-12.99% and the 50N engine thrust and the design state deviate by-13.05% in the state.
TABLE 5 non-correction power system parameter table in fourth state
Before the engine works, the oxidant and the fuel are respectively filled into the control valve group through the pipelines under the action of the pressure of the storage tank, after a working instruction is received, the control valve group is synchronously opened, the propellant respectively enters the combustion chamber through the oxidant nozzle and the fuel nozzle, high-temperature fuel gas is formed after atomization, blending and combustion, and then the fuel gas is discharged from the spray pipe in an accelerating way, so that the thrust required by the upper-stage attitude control is generated.
Because the flow resistance of the system is basically proportional to the square of the flow, under the condition that the outlet pressure of the pressure reducing valve is certain, the variation range of the inlet pressure of the engine is large under the working condition of high and low flow, the pressure range which the engine needs to adapt to is wide, the parameters such as the outlet pressure of the pressure reducing valve, the pressure drop of an adjusting orifice plate and the like need to be reasonably matched, and the system parameters are reasonably distributed by combining the system cold test data, the engine test data and the actual use requirements of control professions.
The engine pressure needs to satisfy the following relationship:
Pc+ΔPinj+ΔPv+ΔPh≤P*;
Wherein P c is the steady-state working chamber pressure of the thrust chamber, deltaP inj is the injection pressure drop, deltaP v is the flow resistance of the propellant control valve bank, deltaP h is the pressure drop of the pressure regulating orifice plate, and P * is the inlet pressure of the propellant control valve bank.
The injection pressure drop and the steady-state working chamber pressure of the thrust chamber have the following relation:
ΔPinj＝0.35Pc;
In addition, according to the design requirement and the actual working conditions of the corresponding components, the parameters are set as follows:
ΔPv=0.2MPa;
ΔPh=0.1MPa;
P*＝3MPa;
therefore, the derived steady-state working chamber pressure requirement of the thrust chamber is as follows:
substituting the parameters into a formula to calculate, wherein the steady-state working chamber pressure of the thrust chamber is 2MPa.
And S3, determining an injection scheme and a profile scheme of an engine thrust chamber.
When determining the injection scheme, for the 50N thrust chamber, the direct current mutual impact injection scheme is adopted because the diameter of the combustion chamber of the thrust chamber is smaller, and the propellant is easy to form a local oxygen-enriched combustion zone on the wall surface of the combustion chamber after being sputtered, so that the structural heat protection of the thrust chamber is not facilitated.
The fuel nozzle and the oxidant nozzle are coaxially nested, the fuel nozzle is arranged inside, the oxidant nozzle is arranged outside, propellant is accelerated by the rotational flow of each rotational flow cavity and then is sprayed out, and an atomization conical surface is formed under the action of centrifugal force. After the oxidant contacts with the fuel, the oxidant is rapidly mixed and atomized under the action of atomization kinetic energy, so that the oxidant is ignited for combustion, and the engine is started to work.
For a 700N thrust chamber, a direct current mutual impact injection design form is adopted, and compared with the existing model, the diameter of the combustion chamber of the thrust chamber is 39mm, the characteristic length is about 310mm, and the combustion efficiency is more than or equal to 0.92.
In determining the profile scheme, the characteristic length L=0.3m of the 50N engine combustion chamber is selected according to the gas residence time requirement (generally 1 ms-3 ms) and the design experience of the combustion chamber of the same type of low-thrust engine. Throat cross-sectional diameter dt=4.24 mm. Calculated by flow density method, the combustion chamber diameter dc=12.5 mm is obtained. The arc radius Rct of the nozzle inlet is selected to be 20mm, the radius Rc1=0.83 Dt of the upstream section of the throat, and the radius Rc2=0.41 Dt of the downstream section of the throat.
In order to achieve the highest nozzle efficiency and the highest thrust, a 50N engine nozzle diffusion section is designed by adopting a characteristic line method. In this embodiment, the design point height is 120km, and the nozzle expansion ratio epsilon=81, the nozzle diffuser expansion half angle βk=32°, and the outlet expansion angle βe=8° are selected according to the design point height and the geometric installation size constraint of the thrust chamber. From the above calculation steps, 50N engine thrust chamber profile dimensions are designed as shown in FIGS. 2 and 3.
The expansion section of the 700N thrust chamber combustion chamber adopts the design of the maximum thrust jet pipe, and the main design parameters of the jet pipe are shown in Table 6.
TABLE 6 Main design parameter table of spray pipe
And S4, performing simulation calculation on the engine thrust chamber, and verifying whether the performance of the engine thrust chamber meets the requirement.
Modeling is carried out by means of simulation software, simulation verification is carried out on the flow trace of the fuel in the inner cavity of the nozzle, and the flow field pressure cloud picture of the fuel nozzle of the 50N engine can be obtained, wherein the inlet pressure of the nozzle is 0.653MPa and basically coincides with the design value of 0.65 MPa.
And meanwhile, verifying the integral, tangential and axial velocity cloud patterns of the flow length in the fuel nozzle, wherein the velocity distribution accords with the actual physical process of the flow of the liquid propelling in the flow field in the centrifugal nozzle.
Simulation analysis is performed through a liquid phase distribution cloud picture of the inner cavity of the fuel nozzle, and it is determined that a conical air cavity is formed at the center of the outlet section of the nozzle under the action of cyclone centrifugation.
For a low-thrust liquid rocket engine, common body heat protection schemes include regenerative cooling, liquid film cooling, passive ablative cooling and the like, and the most applied is a radiation cooling scheme in the field of attitude control engines. Referring to the technical scheme of the same type at present, the thrust chamber body heat protection scheme is selected to be liquid film cooling. The body is made of Nb521, the inner wall is coated with r625 coating, and the thickness of the coating is 80-120 mu m, so that the anti-oxidation and ablation effects can be achieved.
The structural strength simulation software is adopted to simulate and calculate the stress of the body, under the condition that the working pressure of the inner cavity of the thrust chamber is 2MPa, the stress distribution of the inner cavity of the thrust chamber is verified under the conditions that the pressure of the inner cavity of the thrust chamber is 5MPa and the pressure of the inner cavity of the thrust chamber is 10MPa, the structural maximum stress of the shell of the thrust chamber under two working conditions is 29.57MPa and 59.17MPa respectively, and the structural maximum stress is far lower than the tensile strength (sigma b-1600 is more than or equal to 150 MPa) of the Nb521 material at 1600 ℃.
In addition, stress conditions of the injector under various conditions are simulated, and a 4.1MPa hydraulic pressure intensity test, a 2.7MPa inner cavity starting hydraulic impact hydraulic pressure intensity test and stress and deformation analysis of hydraulic chamber pressures of an inner cavity 2.7MPa and a panel 2.0MPa when an engine works are respectively carried out.
The maximum stress point stress of the head under three conditions is calculated to be 38.4MPa, 67.6MPa and 36.6MPa respectively, and the corresponding safety coefficients are 23.2, 13.16 (tensile strength 890MPa in cold state) and 15.57 (tensile strength 570MPa in 400 ℃) respectively, and meanwhile, the deformation under various working conditions is very small, so that the 700N injector structure design is proved to have enough safety margin.
Through the scheme stage of attack and research, the technical indexes which can be achieved by the power system at present are shown in table 7. Besides the engine thrust deviation and the flight reliability index do not meet the overall requirements, other indexes meet the overall requirements, and the initial sample stage is to be subjected to test verification.
TABLE 7 Main index satisfaction of final-repair gesture control power system
An engine thrust chamber designed by an engine thrust chamber design method, comprising:
A combustion chamber;
The combustion device comprises a combustion chamber, a spray pipe, a nozzle and a nozzle, wherein the spray pipe is fixedly arranged at the tail end of the combustion chamber and is communicated with the combustion chamber;
the control valve group is communicated with the combustion chamber and is used for controlling the propellant to enter the combustion chamber;
the nozzle is communicated with the control valve group through a pipeline and is used for atomizing the propellant;
and the throttle orifice plate is communicated with an outlet of the control valve group and used for throttling the propellant.
The engine in this example uses a liquid two-component propellant, the oxidizer is green dinitrogen tetroxide (MON-3), the fuel is methyl hydrazine (MMH), and the theoretical mixing ratio is 1.65.
The structure principle of the 50N engine is shown in fig. 4, and the control valve group of the 50N engine adopts electromagnetic valves.
The structure principle of the 700N engine is shown in fig. 5, and the control valve group adopts an electric air valve.
When the device works, the oxidant outlet and the fuel outlet of the electromagnetic valve or the electric air valve are both provided with the throttle plate, and the two propellants enter the thrust chamber for atomization and mixing, are combusted and decomposed into high-temperature gas, and the high-temperature gas is accelerated to be sprayed out through the spray pipe to generate thrust. The control system sends out a solenoid valve or an electric air valve power-off instruction, the solenoid valve or the electric air valve is closed, and the engine stops working.
As shown in FIGS. 6 and 7, the injector of the 50N engine thrust chamber comprises a head base body, an oxidant swirler and a fuel swirler, wherein the head base body comprises a fuel nozzle and an oxidant nozzle which are coaxially nested, the oxidant nozzle surrounds the inside of the fuel nozzle, the inlet end of the fuel nozzle is communicated with the fuel swirler, and the inlet end of the oxidant nozzle is communicated with the oxidant swirler. The main design parameters of the 50N engine injector are shown in table 8.
Table 850N engine injector main design parameter table
As shown in FIG. 8, the 700N thrust chamber adopts a direct current mutual impact injection mode, and comprises an injector head, wherein the injector head comprises an injection disc and a flow guide cavity, the injection disc comprises 36 pairs of injection holes, the unit thrust of each pair of injection holes is 19.447N, the injection holes are arranged on an inner ring and an outer ring, 12 pairs of injection holes are formed on the inner ring, and 24 pairs of injection holes are formed on the outer ring.
Specifically, the 700N thrust chamber adopts a direct current mutual impact injection design form, and compared with the existing model, the diameter of the combustion chamber of the thrust chamber is 39mm, the characteristic length is about 310mm, and the combustion efficiency is more than or equal to 0.92. The number of injection pairs of the thrust chamber is determined to be 36 pairs (12+24) on the basis of reference to the existing thrust chamber unit thrust, and the unit thrust is 19.447N/pair.
The total injection logarithm of the 700N thrust chamber is 36 pairs, the 700N thrust chamber is distributed in two circles, and each circle logarithm is respectively an inner circle 12 pair and an outer circle 24 pair. In order to achieve a uniform distribution of the flow strength and mixing ratio, the impingement points of the individual rings should be distributed as uniformly as possible along the combustion chamber distribution diameter.
The 700N thrust chamber injector head consists of an injection disk and a diversion cavity. Aiming at the requirements of nearly 500 accumulated working times and the shortest switching pulse of 40ms of the thrust chamber, more severe requirements are provided for the water-proof impact of the diversion cavity at the inlet of the injector, the whole assembly of the injector is mainly connected by laser and electron beam welding, argon arc welding is assisted, the deformation is effectively reduced, and the structural reliability is improved.
The oxidant is in the central cavity, the nozzles are uniformly arranged into inner and outer circles, the aperture and the length-diameter ratio of the two circles of nozzles are consistent, the oxidant is ensured to be uniformly discharged, the fuel cavity is concentrated at the outer side, and the diverter ring adopts a novel eccentric circular structure design, so that the nozzles at the near end and the far end of the liquid collecting cavity are ensured to be uniformly discharged.
It is noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of exemplary embodiments according to the present application. As used herein, the singular is also intended to include the plural unless the context clearly indicates otherwise, and furthermore, it is to be understood that the terms "comprises" and/or "comprising" when used in this specification are taken to specify the presence of stated features, steps, operations, devices, components, and/or combinations thereof.
It should be noted that the terms "first," "second," and the like in the description and the claims of the present application and the above figures are used for distinguishing between similar objects and not necessarily for describing a particular sequential or chronological order. It is to be understood that the data so used may be interchanged where appropriate such that embodiments of the application described herein may be implemented in sequences other than those illustrated or otherwise described herein.
The above description is only of the preferred embodiments of the present invention and is not intended to limit the present invention, but various modifications and variations can be made to the present invention by those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (6)
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202410001633.8A CN117932779B (en) | 2024-01-02 | 2024-01-02 | A method for designing an engine thrust chamber and an engine thrust chamber |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202410001633.8A CN117932779B (en) | 2024-01-02 | 2024-01-02 | A method for designing an engine thrust chamber and an engine thrust chamber |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN117932779A CN117932779A (en) | 2024-04-26 |
| CN117932779B true CN117932779B (en) | 2025-02-11 |
Family
ID=90755029
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202410001633.8A Active CN117932779B (en) | 2024-01-02 | 2024-01-02 | A method for designing an engine thrust chamber and an engine thrust chamber |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117932779B (en) |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106134479B (en) * | 2011-06-28 | 2014-04-23 | 上海空间推进研究所 | The method of a kind of Accurate Assessment space rail control engine vacuum specific impulse performance |
| CN117108413A (en) * | 2023-10-23 | 2023-11-24 | 东方空间技术(山东)有限公司 | Last appearance accuse driving system of repairing |
Family Cites Families (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP6569371B2 (en) * | 2015-08-07 | 2019-09-04 | 京三電機株式会社 | Liquid fuel catcher |
| CN105160051B (en) * | 2015-06-30 | 2018-01-05 | 中国空间技术研究院 | Truss antenna reflector based on multisomic analysis experiment deploys dynamic modeling method |
| CN106672272A (en) * | 2015-11-10 | 2017-05-17 | 上海新跃仪表厂 | Flexibility parameter in-orbit identification ground testing system for spacecraft with flexible structure |
| CN107607325B (en) * | 2017-09-04 | 2019-04-23 | 西安航天动力研究所 | Needle fastens formula ejector filler Variable Thrust Engine System Operating condition adjustment method |
| WO2021110244A1 (en) * | 2019-12-03 | 2021-06-10 | Volvo Truck Corporation | A method for controlling a hybrid powertrain system of a vehicle |
| CN117109372A (en) * | 2023-10-25 | 2023-11-24 | 东方空间技术(山东)有限公司 | Control method of final-repair gesture control power system |
-
2024
- 2024-01-02 CN CN202410001633.8A patent/CN117932779B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106134479B (en) * | 2011-06-28 | 2014-04-23 | 上海空间推进研究所 | The method of a kind of Accurate Assessment space rail control engine vacuum specific impulse performance |
| CN117108413A (en) * | 2023-10-23 | 2023-11-24 | 东方空间技术(山东)有限公司 | Last appearance accuse driving system of repairing |
Also Published As
| Publication number | Publication date |
|---|---|
| CN117932779A (en) | 2024-04-26 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| RU2158839C2 (en) | Liquid-propellant rocket reheat engine | |
| CN111720238B (en) | Liquid oxygen expansion cycle-based depth variable-thrust multi-start liquid rocket engine | |
| CN104329187B (en) | A kind of rocket system of rocket based combined cycle engine variable working condition | |
| US7980056B2 (en) | Methods and apparatus for controlling air flow within a pulse detonation engine | |
| CN109538379B (en) | Device and method for testing thrust chamber of high-thrust liquid rocket engine | |
| EP1741917A2 (en) | Booster rocket engine using gaseous hydrocarbon in catalytically enhanced gas generator cycle | |
| JP2001355515A5 (en) | ||
| CN109139296B (en) | Rocket-based combined cycle engine | |
| CN108612598B (en) | Method for changing combustion mode of bi-mode ramjet engine and engine combustion chamber | |
| CN110907123A (en) | A high-enthalpy propulsion wind tunnel air heating method and device | |
| CN101393098A (en) | Material ablation detection device and method in high-temperature gas | |
| CN113154391A (en) | Oxygen-methane torch ignition device and torch generation method thereof | |
| Thomas et al. | Buildup and operation of a rotating detonation engine | |
| CN115875673A (en) | Pulse combustion wind tunnel hydrogen-oxygen combustion heater applied to high-temperature air incoming flow | |
| CN107642436A (en) | A kind of hybrid rocket engine thrust gas vector controlled structure and method | |
| CN117932779B (en) | A method for designing an engine thrust chamber and an engine thrust chamber | |
| CN112628018B (en) | High-thrust liquid rocket engine and low-power-consumption semi-self-starting method thereof | |
| US6904750B2 (en) | Integral pulse detonation system for a gas turbine engine | |
| CN115014778B (en) | Large-scale high-enthalpy wind tunnel inflow simulated combustion device | |
| CN114962008B (en) | A wide range combined engine ignition and flame stabilization method | |
| CN113339844A (en) | Air hydrogen injection unit and combustion organization method thereof | |
| CN110700963B (en) | Compact layout type solid rocket gas scramjet engine based on axial symmetry | |
| CN117109372A (en) | Control method of final-repair gesture control power system | |
| CN106838902B (en) | A kind of supersonic speed combustion gas injector | |
| CN114109654B (en) | Solid-liquid mixed engine and aircraft |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| CB03 | Change of inventor or designer information |
Inventor after: Wei Kai Inventor after: Huang Shuai Inventor after: Peng Haomin Inventor after: Bu Xiangwei Inventor after: Yao Song Inventor after: Han Yawei Inventor after: Tong Xianyi Inventor before: Wei Kai Inventor before: Peng Haomin Inventor before: Bu Xiangwei Inventor before: Yao Song Inventor before: Han Yawei Inventor before: Tong Xianyi |
|
| CB03 | Change of inventor or designer information | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |