CN113404621B - Solid-liquid mixed engine and method for Mars ascending aircraft - Google Patents
Solid-liquid mixed engine and method for Mars ascending aircraft Download PDFInfo
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- CN113404621B CN113404621B CN202110681883.7A CN202110681883A CN113404621B CN 113404621 B CN113404621 B CN 113404621B CN 202110681883 A CN202110681883 A CN 202110681883A CN 113404621 B CN113404621 B CN 113404621B
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- 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
- F02K9/72—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid and solid propellants, i.e. hybrid rocket-engine plants
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- 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
- F02K9/80—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control
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- 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
- F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
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- 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
- F02K9/97—Rocket nozzles
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Abstract
The invention relates to a solid-liquid mixed engine and a method for a Mars rising aircraft, belonging to the technical field of aerospace propulsion; the device comprises a thrust chamber, an oxidant supply system and a thrust vector control system; the oxidant supply system comprises an oxidant storage tank, an extrusion type oxidant supply subsystem and a pumping pressure type expansion circulation oxidant supply subsystem; the extrusion type oxidant supply subsystem inputs oxidant into the thrust chamber in an extrusion mode, and the pumping pressure type expansion circulation oxidant supply subsystem inputs oxidant into the thrust chamber through expansion circulation; the thrust vector control system regulates the flow of nitrous oxide entering the spray pipe by opening and closing control valves on different secondary flow pipelines, so that thrust vector control and riser posture adjustment are realized. The invention adopts paraffin fuel with excellent ultralow temperature mechanical properties and nitrous oxide with low freezing point and high saturated vapor pressure as propellant, and solves the problems that solid and liquid engines cannot adapt to extreme temperature difference of sparks and are difficult to store for a long time and the like.
Description
Technical Field
The invention belongs to the technical field of aerospace propulsion, and particularly relates to a solid-liquid mixed engine and a method for a mars ascending aircraft.
Background
Mars is the celestial body closest to the earth in the solar system and is the focus of the current planet detection. Since the 60's of the 20 th century, mankind has successfully launched multiple detectors to the surface of a spark, but the scientific instruments carried by the detectors have failed to meet the research requirements of the current scientific community. Therefore, the sample return work on the surface of the mars needs to be carried out urgently. The sample return task can be divided into the following steps: 1) transmitting a detector to the surface of the spark for sampling; 2) the Mars ascending aircraft sends the sample to the Mars orbit; 3) the Mars orbit vehicle carries the sample back to the Earth.
The Mars ascent aircraft is the key of the sample return task, but the Mars surface environment is severe, the average temperature is-63 ℃, and the minimum temperature in winter can reach-133 ℃. At present, most solid and liquid rocket engines cannot adapt to the temperature condition, the glass transition temperature of a propellant adopted by the solid rocket engine is about-70 ℃, the propellant can crack, deform and the like at extremely low temperature, so that the engine cannot work normally, the integral heat preservation of the riser is needed to solve the problem of propellant cracking, and the cost of returning a sample to a task is greatly increased; the propellant commonly used by the liquid rocket engine has the characteristics of high freezing point, low saturated vapor pressure and the like, cannot be stored for a long time, and cannot be used as a spark-lift aircraft engine.
The oxidant conveying mode of the solid-liquid mixed engine is divided into an extrusion type and a pumping pressure type, and an extrusion type oxidant conveying system is not suitable for being applied to engines working at high room pressure for a long time; and the performance of the turbine pump is reduced under the condition of large flow regulation ratio of the pumping type oxidant conveying system. Based on this, a patent entitled "electric pump delivery system for solid-liquid rocket engine", published by yunan jia, zhuan jun, etc. of beijing aerospace university, the invention of which patent application No. 201710618855.4, proposes that a pump is driven by a motor to increase pressure, the flow rate is controlled by a flow rate adjusting valve, and the oxidant can be supplied to a combustion chamber at stable pressure and flow rate by adjusting the rotation speed of the pump and the flow rate adjusting valve. In an article published in the journal of aviation dynamics by the university of Beijing aerospace, named as 'simulation and optimization design of pump pressure type solid-liquid rocket engine system', a pump pressure type oxidant conveying system for driving a turbine pump by high-temperature working medium generated by catalytic decomposition of hydrogen peroxide is provided, and compared with an extrusion type conveying system, the volume of the system is reduced by 20%.
Disclosure of Invention
The technical problem to be solved is as follows:
in order to avoid the defects of the prior art, the invention provides a solid-liquid mixed engine for a Mars lift aircraft, wherein the engine takes paraffin fuel/nitrous oxide as a propellant. The glass transition temperature of the paraffin fuel is extremely low, and the paraffin fuel can be kept relatively low at the temperature of a Mars environmentGood mechanical property, and nitrous oxide has the characteristics of high saturated vapor pressure and low freezing point, and can be stored for a long time at low temperature. Meanwhile, the theoretical specific impulse of the paraffin fuel/nitrous oxide solid-liquid mixed engine can reach 3100 N.s.kg -1 The average thrust can reach 3400N, and the task requirement of a Mars ascending aircraft is met.
The technical scheme of the invention is as follows: a solid-liquid mixed engine for a Mars rising aircraft comprises a thrust chamber; the method is characterized in that: the thrust chamber is connected with the oxidant supply system and the thrust vector control system;
the oxidant supply system comprises an oxidant storage tank 1, an extrusion type oxidant supply subsystem and a pumping pressure type expansion circulation oxidant supply subsystem; the oxidant storage tank 1 is arranged above the thrust chamber, the extrusion type oxidant supply subsystem inputs oxidant into the thrust chamber in an extrusion mode, and the pumping type expansion circulation oxidant supply subsystem inputs oxidant into the thrust chamber through expansion circulation;
the extrusion type oxidant supply subsystem comprises a high-pressure helium storage tank 5, a pressurization pipeline 28 and an extrusion supply pipeline 29; a plurality of high-pressure helium storage tanks 5 are arranged at the periphery of the thrust chamber, the outlets of the high-pressure helium storage tanks are communicated with each other through pipelines, and the high-pressure helium storage tanks are communicated with the oxidizer storage tank 1 through a pressurization pipeline 28; the outlet of the pressurization pipeline 28 extends into the oxidizer storage tank 1 and extends to the upper part of the inside of the oxidizer storage tank 1, and is used for introducing helium in the high-pressure helium storage tank 5 into the upper part of the inside of the oxidizer storage tank 1 to supply downward pressure to the oxidizer; the output port of the oxidant storage tank 1 is communicated with the thrust chamber through an extrusion supply pipeline 29, and the oxidant output by extrusion is input into the thrust chamber and is combusted with the paraffin fuel 11 in the combustion chamber;
the pumping type expansion circulation oxidant supply subsystem comprises a turbo pump 2, a first cooling circulation pipeline 31, a turbo pump exhaust pipeline 33, a first pumping pressure supply pipeline 34, a second cooling circulation pipeline 41 and a second pumping pressure supply pipeline 35; the upper end of the first cooling circulation pipeline 31 is communicated with the oxidant storage tank 1, the lower end of the first cooling circulation pipeline is communicated with a spray pipe interlayer of the thrust chamber, and the oxidant is input into the spray pipe interlayer; the upper end of the second cooling circulation pipeline 41 is communicated with the turbo pump 2, the lower end of the second cooling circulation pipeline is communicated with a spray pipe interlayer of the thrust chamber, an oxidant heated by the spray pipe interlayer is input into the turbo pump 2, and the turbine expands to do work to drive the turbo pump 2 to work; the turbo pump 2 is arranged above the thrust chamber, the inlet of the turbo pump is communicated with the oxidant storage tank 1 through a first pump pressure supply pipeline 34, and the outlet of the turbo pump is communicated with the thrust chamber through a second pump pressure supply pipeline 35; after the turbine pump 2 is started, an oxidant is input into the thrust chamber and is combusted with the paraffin fuel 11 in the combustion chamber, and meanwhile, the oxidant driving the turbine pump to work enters the combustion chamber through a turbine pump exhaust pipeline 33 to participate in combustion;
the thrust vector control system comprises a vector control flow regulator 27, a thrust vector control pipeline 30 and a secondary flow pipeline 32; the thrust vector control pipeline 30 is communicated with the oxidant storage tank 1 at the upper end and communicated with the secondary flow pipeline 32 at the lower end, the secondary flow pipelines 32 are circumferentially arranged at the periphery of the expansion section of the thrust chamber spray pipe and are communicated with a plurality of spray ports formed in the circumferential direction of the expansion section of the thrust chamber spray pipe, and the flow of nitrous oxide entering the expansion section of the spray pipe is adjusted by opening and closing control valves on different secondary flow pipelines 32, so that the thrust vector control and the posture adjustment of the riser are realized.
The further technical scheme of the invention is as follows: the thrust chamber comprises a boron/potassium nitrate igniter 3, a thrust chamber shell 6, a regenerative cooling spray pipe 7, an injector 8, a front combustion chamber 9, a front combustion chamber heat-insulating layer 10, a paraffin grain 11, a combustion chamber 12, a combustion chamber heat-insulating layer 13, a rear combustion chamber heat-insulating layer 14 and a rear combustion chamber 15; 2 boron/potassium nitrate igniters 3 are arranged at the front combustion chamber; the injector 8 adopts a porous straight-flow injector with small hydraulic loss; the front combustion chamber 9 further atomizes and disperses the oxidant, and a front combustion chamber heat-insulating layer 10 is arranged on the wall surface of the front combustion chamber 9 to carry out heat protection on the thrust chamber shell 6; the outer peripheral surface of the combustion chamber 12 is coated with a combustion chamber heat-insulating layer 13 and a thrust chamber shell 6, and the paraffin fuel 11 is arranged in the combustion chamber, so that the low-temperature mechanical property of the paraffin fuel can adapt to the surface temperature environment of a spark; in order to improve the combustion efficiency, a rear combustion chamber 15 is designed, and a rear combustion chamber heat insulation layer 14 is arranged on the wall surface of the rear combustion chamber; the spray pipe adopts a regenerative cooling spray pipe 7, the outer wall of the spray pipe is of a sandwich structure formed by double walls, and the thermal protection of the spray pipe is realized by utilizing the convection cooling of liquid nitrous oxide in the sandwich layer.
The further technical scheme of the invention is as follows: the front end cover 4 of the thrust chamber is arranged at the upper end of the thrust chamber, and 2 boron/potassium nitrate igniters 3 are arranged on the front end cover and are used for realizing ignition of an engine.
The further technical scheme of the invention is as follows: the extrusion type oxidant supply subsystem further comprises a filling discharge valve 16, a pressure reducing valve 17, a pressure increasing control valve 19, an extrusion system control valve 23 and an extrusion system flow regulator 25; the filling and discharging valve 16 is arranged at the outer end of the mutually communicated pipelines of the high-pressure helium storage tanks 5 and is used for filling helium into the high-pressure helium storage tanks 5; the pressure reducing valve 17 and the pressure increasing control valve 19 are arranged on the pressure increasing pipeline 28 and are used for controlling the pressure of the gas output by the pressure increasing pipeline 28; the extrusion system control valve 23 and the extrusion system flow regulator 25 are arranged on the extrusion supply pipeline 29 and are used for controlling the output of the oxidant and regulating the flow;
the further technical scheme of the invention is as follows: the pumping type expansion circulation oxidant supply subsystem further comprises a check valve 18, a cooling circulation pipeline control valve 20, a circulation loop pipeline control valve 22, a pumping system control valve 24 and a cooling circulation flow regulator 26; the check valve 18 is arranged on the exhaust pipeline 33 of the turbopump to prevent the oxidant from flowing back; the cooling circulation line control valve 20 and the cooling circulation flow regulator 26 are provided on the first cooling circulation line 31 for output and flow regulation of the oxidizing agent; a recycle loop line control valve 22 is provided on the second cooling recycle line 41 and a pump system control valve 24 is provided on the first pump supply line 34, both for controlling oxidant input.
The further technical scheme of the invention is as follows: the number of the high-pressure helium storage tanks 5 is 4, the high-pressure helium storage tanks are uniformly distributed on the periphery of the thrust chamber shell 6 along the circumferential direction and are used for pressurizing the oxidant storage tank.
The further technical scheme of the invention is as follows: the secondary flow pipeline 32 is an L-shaped pipe, the long arm end of the secondary flow pipeline is arranged on the outer wall surface of the spray pipe along the axial direction of the expansion section of the spray pipe, and the short arm end of the secondary flow pipeline is inserted into the nozzle of the spray pipe; the upper ends of the plurality of long arms are communicated with the thrust vector control line 30 after being communicated with each other through an annular pipe.
The invention further adopts the technical scheme that: the number of the secondary flow pipelines 32 is 4, and the secondary flow pipelines are uniformly distributed along the circumferential direction.
The further technical scheme of the invention is as follows: the thrust vector control system further includes a secondary flow line control valve 21, and the on/off of each secondary flow line 32 is notified by the secondary flow line control valve 21.
An oxidant supply method of a solid-liquid mixed engine for a Mars aircraft is characterized by comprising the following specific steps:
the method comprises the following steps: after an engine standby instruction is given, opening a pressure increasing control valve 19, increasing pressure in the oxidant storage tank 1, and enabling the system to enter a standby state;
step two: after a formal working instruction is issued, the boron/potassium nitrate igniter 3 is started to generate high-temperature fuel gas, the extrusion system control valve 23 is opened after 0.5s, the paraffin wax fuel 11 is ignited under the action of the high-temperature fuel gas and nitrous oxide, the ignition and starting of the engine are finished, and the engine adopts an extrusion type oxidant supply mode at the moment;
step three: after ignition is started, a cooling circulation pipeline control valve 20 and a circulation loop pipeline control valve 22 are opened, nitrous oxide flows through an interlayer in the regeneration cooling spray pipe 7 to absorb heat and then enters a turbine, a turbine pump 2 is started, a pump pressure system control valve 24 is opened, and the turbine pump 2 absorbs the nitrous oxide from the oxidant storage tank 1 and supplies the nitrous oxide to a thrust chamber; the turbo pump works stably, after the pressure in the combustion chamber is monitored to be stable, the control valve 23 of the extrusion system is closed, and the engine adopts an expansion cycle type oxidant supply mode at the moment;
step four: if the attitude of the riser needs to be adjusted in a certain direction in the flying process, one or two secondary flow pipeline control valves 21 on the four groups of secondary flow pipelines 32 are opened, the flow of the oxidant entering the expansion section of the spray pipe is controlled by the vector control flow regulator 27, nitrous oxide enters the expansion section of the spray pipe, is rapidly atomized and evaporated and is mixed with fuel gas to interfere the supersonic fuel gas flow to generate oblique shock waves, so that the supersonic fuel gas flow is deflected in the direction to generate lateral force, and the attitude adjustment of the riser is realized.
Advantageous effects
The invention has the beneficial effects that:
1) the solid-liquid mixed engine designed by the invention adopts the paraffin fuel with excellent ultralow temperature mechanical property and the nitrous oxide with low freezing point and high saturated vapor pressure as the propellant, thereby solving the problems that solid and liquid engines cannot adapt to extreme temperature difference of sparks and are difficult to store for a long time and the like;
2) the oxidant supply system of the solid-liquid mixed engine designed by the invention combines an extrusion type oxidant supply system and a pumping pressure type oxidant supply system, and solves the problems that because the environmental temperature of a spark is lower, the vapor pressure of nitrous oxide is lower, and the expansion cycle supply system cannot be stably started by only depending on the vapor pressure of nitrous oxide;
3) the solid-liquid mixed engine oxidant supply system provided by the invention utilizes the advantage of quick response of an extrusion type oxidant supply mode system, can quickly build pressure in a thrust chamber of an engine to achieve a stable working state, then converts the pressure into a pumping type oxidant supply mode, utilizes the stable working state of the pumping type oxidant supply mode, can effectively avoid the problem of overlarge pressure drop of the thrust chamber in the working process of the engine in the extrusion type oxidant supply mode, and can be used for realizing the highest boron/potassium nitrate fuel gas concentration in the thrust chamber at 0.5s through the calculation of the ignition gas forming rate of a boron/potassium nitrate igniter adopted by the invention, thereby being most beneficial to the starting of the engine;
4) the solid-liquid mixed engine oxidant supply system designed by the invention utilizes the oxidant to absorb the heat transferred from high-temperature fuel gas to the wall surface of the spray pipe, the generated high temperature drives the turbine pump to work, the energy utilization rate of the engine is improved, and meanwhile, the oxidant evaporates and absorbs heat to cool the wall surface of the spray pipe, so that the spray pipe can be prevented from being ablated;
5) the thrust vector control system adopted by the invention injects nitrous oxide with certain pressure into the expansion section of the spray pipe through the control valve and the flow regulator to interfere supersonic speed gas flow, thereby achieving the purpose of thrust vector control. The thrust vector control system only needs to control the valve and the flow regulator to work coordinately, has good time responsiveness, and is used for fixing the spray pipe, so that the structure of the spray pipe is simplified. Specifically, during the flight of the riser, the attitude of the riser needs to be adjusted in a certain direction, one or two secondary flow pipeline control valves 21 on four groups of secondary flow pipelines 32 are opened, the flow of an oxidant entering the expansion section of the spray pipe is controlled by a vector control flow regulator 27, nitrous oxide enters the expansion section of the spray pipe, is rapidly atomized and evaporated and is mixed with gas to interfere with supersonic gas flow to generate oblique shock waves, so that the supersonic gas flow is deflected in the direction to generate lateral force, and the attitude adjustment of the riser is realized. It is calculated that the maximum controllable riser vector deflection is 3 ° in each direction.
Drawings
FIG. 1 is a diagram of a solid-liquid hybrid engine system according to the present invention;
FIG. 2 is a structural diagram of a solid-liquid hybrid engine according to the present invention; (1) is a front view, (2) a front sectional view, (3) a right sectional view;
FIG. 3 is a partial pipeline structure diagram of the solid-liquid mixing engine according to the present invention; (1) is a structure diagram of a supply pipeline; (2) thrust vector control pipeline structure diagram.
Description of reference numerals: 1. oxidant storage tank, 2 turbo pump, 3 boron/potassium nitrate igniter, 4 thrust chamber front end cover, 5 high-pressure helium storage tank, 6 thrust chamber shell, 7 regenerative cooling spray pipe, 8 injector, 9 front combustion chamber, 10 front combustion chamber heat insulation layer, 11 paraffin grain, 12 combustion chamber, 13 combustion chamber heat insulation layer, 14 rear combustion chamber heat insulation layer, 15 rear combustion chamber, 16 filling discharge valve, 17 pressure reducing valve, 18 check valve, 19 pressure increasing control valve, 20 cooling circulation pipeline control valve, 21 secondary flow pipeline control valve, 22 circulation loop pipeline control valve, 23 extrusion system control valve, 24 pumping system control valve, 25 extrusion system flow regulator, 26 cooling circulation flow regulator, 27 vector flow regulator, 28 pressure increasing pipeline, 29 extrusion supply pipeline, 30. Thrust vector control circuit, 31, first cooling circulation circuit, 32, secondary flow circuit, 33, turbo pump exhaust circuit, 34, first pump pressure supply circuit, 35, second pump pressure supply circuit, 41, second cooling circulation circuit.
Detailed Description
The embodiments described below with reference to the drawings are illustrative and intended to be illustrative of the invention and are not to be construed as limiting the invention.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", and the like, indicate orientations and positional relationships based on those shown in the drawings, and are used only for convenience of description and simplicity of description, and do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be considered as limiting the present invention.
Referring to fig. 1-3, the solid-liquid mixed engine for the Mars aircraft comprises a thrust chamber, an oxidant supply system and a thrust vector control system;
the oxidant supply system comprises an oxidant storage tank 1, an extrusion type oxidant supply subsystem and a pumping pressure type expansion circulation oxidant supply subsystem; the oxidant storage tank 1 is arranged above the thrust chamber, the extrusion type oxidant supply subsystem inputs oxidant into the thrust chamber in an extrusion mode, and the pumping type expansion circulation oxidant supply subsystem inputs oxidant into the thrust chamber through expansion circulation;
the extrusion type oxidant supply subsystem comprises a high-pressure helium storage tank 5, a pressurization pipeline 28 and an extrusion supply pipeline 29; a plurality of high-pressure helium storage tanks 5 are arranged at the periphery of the thrust chamber, the outlets of the high-pressure helium storage tanks are communicated with each other through pipelines, and the high-pressure helium storage tanks are communicated with the oxidizer storage tank 1 through a pressurization pipeline 28; the outlet of the pressurization pipeline 28 extends into the oxidizer storage tank 1 and extends to the upper part of the inside of the oxidizer storage tank 1, and is used for introducing helium in the high-pressure helium storage tank 5 into the upper part of the inside of the oxidizer storage tank 1 to supply downward pressure to the oxidizer; the output port of the oxidant storage tank 1 is communicated with the thrust chamber through an extrusion supply pipeline 29, and the oxidant output by extrusion is input into the thrust chamber to be combusted with the paraffin fuel 11 in the combustion chamber;
the extrusion type oxidant supply subsystem further comprises a filling discharge valve 16, a pressure reducing valve 17, a pressure increasing control valve 19, an extrusion system control valve 23 and an extrusion system flow regulator 25; the filling and discharging valve 16 is arranged at the outer end of the mutually communicated pipelines of the high-pressure helium storage tanks 5 and is used for filling helium into the high-pressure helium storage tanks 5; the pressure reducing valve 17 and the pressure increasing control valve 19 are arranged on the pressure increasing pipeline 28 and are used for controlling the pressure of the gas output by the pressure increasing pipeline 28; an extrusion system control valve 23 and an extrusion system flow regulator 25 are provided on the extrusion supply line 29 for controlling the oxidant output and flow regulation.
The pumping pressure type expansion circulation oxidant supply subsystem comprises a turbo pump 2, a first cooling circulation pipeline 31, a turbo pump exhaust pipeline 33, a first pumping pressure supply pipeline 34, a second pumping pressure supply pipeline 35 and a second cooling circulation pipeline 41; the upper end of the first cooling circulation pipeline 31 is communicated with the oxidant storage tank 1, and the lower end of the first cooling circulation pipeline is communicated with a spray pipe interlayer of the thrust chamber and is input into the spray pipe interlayer; the upper end of the second cooling circulation pipeline 41 is communicated with the turbo pump 2, the lower end of the second cooling circulation pipeline is communicated with a spray pipe interlayer of the thrust chamber, an oxidant heated by the spray pipe interlayer is input into the turbo pump 2, and the turbine expands to do work to drive the turbo pump 2 to work; the turbo pump 2 is arranged above the thrust chamber, the inlet of the turbo pump is communicated with the oxidant storage tank 1 through a first pump pressure supply pipeline 34, and the outlet of the turbo pump is communicated with the thrust chamber through a second pump pressure supply pipeline 35; after the turbine pump 2 is started, an oxidant is input into the thrust chamber and is combusted with the paraffin fuel 11 in the combustion chamber, and meanwhile, the oxidant driving the turbine pump to work enters the combustion chamber through a turbine pump exhaust pipeline 33 to participate in combustion;
the pumping type expansion circulation oxidant supply subsystem further comprises a check valve 18, a cooling circulation pipeline control valve 20, a circulation loop pipeline control valve 22, a pumping system control valve 24 and a cooling circulation flow regulator 26; the check valve 18 is arranged on the exhaust pipeline 33 of the turbopump to prevent the oxidant from flowing back; the cooling circulation line control valve 20 and the cooling circulation flow regulator 26 are disposed on the first cooling circulation line 31 for output and flow regulation of the oxidizing agent; a recycle loop line control valve 22 is provided on the second cooling recycle line 41 and a pump system control valve 24 is provided on the first pump supply line 34, both for controlling oxidant input.
The thrust vector control system comprises a vector control flow regulator 27, a thrust vector control pipeline 30 and a secondary flow pipeline 32; the thrust vector control pipeline 30 is communicated with the oxidant storage tank 1 at the upper end and communicated with the secondary flow pipeline 32 at the lower end, the secondary flow pipelines 32 are circumferentially arranged at the periphery of the expansion section of the thrust chamber spray pipe and are communicated with a plurality of spray ports formed in the circumferential direction of the expansion section of the thrust chamber spray pipe, and the flow of nitrous oxide entering the expansion section of the spray pipe is adjusted by opening and closing control valves on different secondary flow pipelines 32, so that the thrust vector control and the posture adjustment of the riser are realized.
The thrust chamber comprises a boron/potassium nitrate igniter 3, a thrust chamber shell 6, a regenerative cooling spray pipe 7, an injector 8, a front combustion chamber 9, a front combustion chamber heat-insulating layer 10, a paraffin grain 11, a combustion chamber 12, a combustion chamber heat-insulating layer 13, a rear combustion chamber heat-insulating layer 14 and a rear combustion chamber 15; 2 boron/potassium nitrate igniters 3 are arranged at the front combustion chamber; the injector 8 adopts a porous straight-flow injector with small hydraulic loss; the front combustion chamber 9 further atomizes and disperses the oxidant, and a front combustion chamber heat-insulating layer 10 is arranged on the wall surface of the front combustion chamber 9 to carry out heat protection on the thrust chamber shell 6; the outer peripheral surface of the combustion chamber 12 is coated with a combustion chamber heat-insulating layer 13 and a thrust chamber shell 6, and the paraffin fuel 11 is arranged in the combustion chamber, so that the low-temperature mechanical property of the paraffin fuel can adapt to the surface temperature environment of a spark; in order to improve the combustion efficiency, a rear combustion chamber 15 is designed, and a rear combustion chamber heat insulation layer 14 is arranged on the wall surface of the rear combustion chamber; the spray pipe adopts a regenerative cooling spray pipe 7, the outer wall of the spray pipe is of a sandwich structure formed by double walls, and the thermal protection of the spray pipe is realized by utilizing the convection cooling of liquid nitrous oxide in the sandwich layer. The front end cover 4 of the thrust chamber is arranged at the upper end of the thrust chamber, and 2 boron/potassium nitrate igniters 4 are arranged on the front end cover and are used for realizing ignition of an engine.
Example (b):
the invention relates to a solid-liquid mixed engine for a Mars rising aircraft, which comprises: an oxidizer storage tank 1, a turbopump 2, a boron/potassium nitrate igniter 3, a thrust chamber front end cover 4, a high-pressure helium storage tank 5, a thrust chamber housing 6, a regenerative cooling nozzle 7, an injector 8, a front combustion chamber 9, a front combustion chamber heat-insulating layer 10, a paraffin grain 11, a combustion chamber 12, a combustion chamber heat-insulating layer 13, a rear combustion chamber heat-insulating layer 14, a rear combustion chamber 15, a filling and discharging valve 16, a pressure reducing valve 17, a check valve 18, a pressure increasing control valve 19, a cooling circulation line control valve 20, a secondary flow line control valve 21, a circulation circuit line control valve 22, an extrusion system control valve 23, a pumping system control valve 24, an extrusion system flow regulator 25, a cooling circulation flow regulator 26, a vector control flow regulator 27, a pressure increasing line 28, an extrusion supply line 29, a thrust vector control line 30, a first cooling circulation line 31, a secondary flow line 32, a second cooling circulation line 32, a pressure increasing line 17, a pressure increasing line, a pressure increasing, A turbo pump exhaust line 33, a first pump pressure supply line 34, a second pump pressure supply line 35, and a second cooling circulation line 41.
Extrusion type oxidant supply subsystem assembly: the four high-pressure helium storage tanks 5 are communicated through a pressurization pipeline 28, a filling and discharging valve 16 is installed at one end of the pipeline 28 and used for filling helium into the high-pressure helium storage tanks 5, the other end of the pressurization pipeline 28 is communicated into the oxidant storage tank 1, and a pressurization control valve 19 and a pressure reducing valve 17 are installed on the pressurization pipeline 28. The extrusion supply pipeline 29 is used for communicating the oxidant storage tank 1 with the thrust chamber and conveying the oxidant into the thrust chamber, and an extrusion system control valve 23 and an extrusion system flow regulator 25 are installed on the extrusion supply pipeline 29 to realize the flow management of the oxidant entering the thrust chamber. And a filling discharge valve is additionally connected to the oxidant storage tank to realize the filling of the oxidant.
The pumping type expansion circulation epoxidizing agent supply subsystem is assembled: the first cooling circulation pipeline 31 is communicated with the oxidant storage tank 1, the regeneration cooling spray pipe 7 and the turbine pump 2 and is used for introducing low-temperature oxidant into an interlayer of the regeneration cooling spray pipe 7 to realize active thermal protection of the spray pipe, the high-temperature oxidant after heat absorption is introduced into the turbine to expand and do work to push the turbine pump 2 to work, the turbine exhausts the expanded nitrous oxide gas into a thrust chamber through a turbine pump exhaust pipeline 33, and the turbine pump exhaust pipeline 33 is provided with a one-way valve 18 to prevent the oxidant from flowing back. The first pump pressure supply line 34 and the second pump pressure supply line 35 respectively communicate the oxidizer tank 1, the turbo pump 2, and the thrust chamber, and the turbo pump sucks the oxidizer to supply the oxidizer to the thrust chamber at a certain pressure.
Assembling a thrust vector control system: one end of a thrust vector control pipeline 30 is connected with the oxidant storage tank 1, the other end of the thrust vector control pipeline is connected with a flow divider, the flow divider leads out four secondary flow pipelines 32 which are connected with four jet ports of an expansion section of the regenerative cooling spray pipe 7 and lead the oxidant to the expansion section of the spray pipe, each secondary flow pipeline is provided with a vector control flow regulator 27 and a secondary flow pipeline control valve 21, and thrust vector control and riser posture adjustment are realized by opening and closing different secondary flow pipelines and regulating the flow of the oxidant.
After the engine is assembled, the high-pressure helium tank 5 is filled with high-pressure helium through the filling and discharging valve 16 on the pressurization pipeline 28, and the pressure in the high-pressure helium tank 5 is required to be not lower than 10 MPa. The tank is filled with nitrous oxide liquid through a filling discharge valve on the oxidant tank 1, and the mass of the nitrous oxide liquid is not less than 238 kg.
The specific operation of this embodiment:
after the engine standby instruction is given, the pressure increasing control valve 19 is opened to increase the pressure in the oxidant storage tank 1, and the system enters a standby state. After a formal working instruction is issued, the boron/potassium nitrate igniter 3 is started to generate high-temperature fuel gas, the extrusion system control valve 23 is opened after 0.5s, the paraffin fuel 11 is ignited under the action of the high-temperature fuel gas and nitrous oxide, the ignition and starting of the engine are completed, and at the moment, the engine adopts an extrusion type oxidant supply mode. After ignition is started, the cooling circulation pipeline control valve 20 and the circulation loop pipeline control valve 22 are opened, nitrous oxide flows through an interlayer in the regeneration cooling spray pipe 7 to absorb heat and then enters a turbine, the turbine pump 2 is started, the pump pressure system control valve 24 is opened, and the turbine pump 2 sucks the nitrous oxide from the oxidant storage tank 1 and supplies the nitrous oxide to the thrust chamber. The turbo pump works smoothly, after the pressure in the combustion chamber is monitored to be stable, the control valve 23 of the extrusion system is closed, and at the moment, the engine adopts an expansion circulation type oxidant supply mode. By adjusting the cooling circulation flow regulator 26, the flow of nitrous oxide entering the interlayer of the regenerative cooling nozzle 7 is managed, the rotation speed of the turbopump 2 is controlled, and the oxidant is stably supplied to the thrust chamber.
If the attitude of the riser needs to be adjusted in a certain direction in the flying process, one or two secondary flow pipeline control valves 21 on the four groups of secondary flow pipelines 32 are opened, the flow of the oxidant entering the expansion section of the spray pipe is controlled by the vector control flow regulator 27, nitrous oxide enters the expansion section of the spray pipe, is rapidly atomized and evaporated and is mixed with fuel gas to interfere the supersonic fuel gas flow to generate oblique shock waves, so that the supersonic fuel gas flow is deflected in the direction to generate lateral force, and the attitude adjustment of the riser is realized. It is calculated that the maximum controllable riser vector deflection is 3 ° in each direction.
Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that variations, modifications, substitutions and alterations can be made in the above embodiments by those of ordinary skill in the art without departing from the principle and spirit of the present invention.
Claims (10)
1. A solid-liquid mixed engine for a Mars rising aircraft comprises a thrust chamber; the method is characterized in that: the thrust chamber is connected with the oxidant supply system and the thrust vector control system;
the oxidant supply system comprises an oxidant storage tank (1), an extrusion type oxidant supply subsystem and a pumping type expansion circulation oxidant supply subsystem; the oxidant storage tank (1) is arranged above the thrust chamber, the extrusion type oxidant supply subsystem inputs oxidant into the thrust chamber in an extrusion mode, the extrusion type oxidant supply mode is firstly utilized to quickly build pressure in the thrust chamber of the engine to achieve a stable working state, then the extrusion type oxidant supply mode is converted into a pumping type oxidant supply mode, and the pumping type expansion circulation oxidant supply subsystem inputs oxidant into the thrust chamber through expansion circulation;
the extrusion type oxidant supply subsystem comprises a high-pressure helium storage tank (5), a pressurization pipeline (28) and an extrusion supply pipeline (29); a plurality of high-pressure helium storage tanks (5) are arranged at the periphery of the thrust chamber, and the outlets of the high-pressure helium storage tanks are communicated with each other through pipelines and are communicated with the oxidant storage tank (1) through a pressurization pipeline (28); the outlet of the pressurization pipeline (28) extends into the oxidant storage tank (1) and extends to the upper part of the inside of the oxidant storage tank (1) for introducing helium in the high-pressure helium storage tank (5) into the upper part of the inside of the oxidant storage tank (1) to supply downward pressure to the oxidant; the output port of the oxidant storage tank (1) is communicated with the thrust chamber through an extrusion supply pipeline (29), and the oxidant output by extrusion is input into the thrust chamber and is combusted with the paraffin fuel (11) in the combustion chamber;
the pumping pressure type expansion circulation oxidant supply subsystem comprises a turbo pump (2), a first cooling circulation pipeline (31), a turbo pump exhaust pipeline (33), a first pumping pressure supply pipeline (34), a second cooling circulation pipeline (41) and a second pumping pressure supply pipeline (35); the upper end of the first cooling circulation pipeline (31) is communicated with the oxidant storage tank (1), the lower end of the first cooling circulation pipeline is communicated with a spray pipe interlayer of the thrust chamber, and the oxidant is input into the spray pipe interlayer; the upper end of the second cooling circulation pipeline (41) is communicated with the turbo pump (2), the lower end of the second cooling circulation pipeline is communicated with a spray pipe interlayer of the thrust chamber, an oxidant heated by the spray pipe interlayer is input into the turbo pump (2), and the turbo pump (2) is driven to work by expansion; the turbine pump (2) is arranged above the thrust chamber, the inlet of the turbine pump is communicated with the oxidant storage tank (1) through a first pump pressure supply pipeline (34), and the outlet of the turbine pump is communicated with the thrust chamber through a second pump pressure supply pipeline (35); after the turbine pump (2) is started, an oxidant is input into the thrust chamber and is combusted with the paraffin fuel (11) in the combustion chamber, and meanwhile, the oxidant driving the turbine pump to work enters the combustion chamber through a turbine pump exhaust pipeline (33) to participate in combustion;
the thrust vector control system comprises a vector control flow regulator (27), a thrust vector control pipeline (30) and a secondary flow pipeline (32); the thrust vector control pipeline (30) is communicated with the oxidant storage tank (1) at the upper end, is communicated with the secondary flow pipeline (32) at the lower end, is arranged at the periphery of the expansion section of the thrust chamber spray pipe along the circumferential direction, is communicated with a plurality of spray ports formed in the circumferential direction of the expansion section of the thrust chamber spray pipe, and is used for adjusting the flow of nitrous oxide entering the expansion section of the spray pipe by opening and closing control valves on different secondary flow pipelines (32) so as to realize thrust vector control and riser posture adjustment.
2. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the thrust chamber comprises a boron/potassium nitrate igniter (3), a thrust chamber shell (6), a regenerative cooling spray pipe (7), an injector (8), a front combustion chamber (9), a front combustion chamber heat-insulating layer (10), paraffin fuel (11), a combustion chamber (12), a combustion chamber heat-insulating layer (13), a rear combustion chamber heat-insulating layer (14) and a rear combustion chamber (15); 2 boron/potassium nitrate igniters (3) are arranged at the front combustion chamber; the injector (8) adopts a porous straight-flow injector with small hydraulic loss; the front combustion chamber (9) further atomizes and disperses the oxidant, and a front combustion chamber heat-insulating layer (10) is arranged on the wall surface of the front combustion chamber (9) to carry out heat protection on the thrust chamber shell (6); the outer peripheral surface of the combustion chamber (12) is coated with a combustion chamber heat-insulating layer (13) and a thrust chamber shell (6), and the paraffin fuel (11) is arranged in the combustion chamber, so that the low-temperature mechanical property of the paraffin fuel can adapt to the surface temperature environment of a spark; in order to improve the combustion efficiency, a rear combustion chamber (15) is designed, and a rear combustion chamber heat insulation layer (14) is arranged on the wall surface of the rear combustion chamber; the spray pipe adopts a regenerative cooling spray pipe (7), the outer wall of the spray pipe is of a sandwich structure formed by double walls, and the thermal protection of the spray pipe is realized by utilizing the convection cooling of liquid nitrous oxide in the sandwich layer.
3. The Mars aircraft solid-liquid hybrid engine of claim 2, characterized in that: the front end cover (4) of the thrust chamber is arranged at the upper end of the thrust chamber, and 2 boron/potassium nitrate igniters (3) are arranged on the front end cover and are used for realizing ignition of an engine.
4. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the extrusion type oxidant supply subsystem further comprises a filling discharge valve (16), a pressure reducing valve (17), a pressure increasing control valve (19), an extrusion system control valve (23) and an extrusion system flow regulator (25); the filling and discharging valve (16) is arranged at the outer end of the mutually communicated pipelines of the high-pressure helium storage tanks (5) and is used for filling helium into the high-pressure helium storage tanks (5); the pressure reducing valve (17) and the pressure increasing control valve (19) are arranged on the pressure increasing pipeline (28) and are used for controlling the pressure of the output gas of the pressure increasing pipeline (28); the extrusion system control valve (23) and the extrusion system flow regulator (25) are arranged on the extrusion supply pipeline (29) and are used for controlling the output of the oxidant and regulating the flow.
5. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the pumping type expansion circulation oxidant supply subsystem further comprises a check valve (18), a cooling circulation pipeline control valve (20), a circulation loop pipeline control valve (22), a pumping system control valve (24) and a cooling circulation flow regulator (26); the one-way valve (18) is arranged on an exhaust pipeline (33) of the turbopump and used for preventing the oxidant from flowing back; the cooling circulation pipeline control valve (20) and the cooling circulation flow regulator (26) are arranged on the first cooling circulation pipeline (31) and are used for outputting and regulating the flow of the oxidant; a recycle loop line control valve (22) is provided on the second cooling recycle line (41), and a pumping system control valve (24) is provided on the first pumping supply line (34), both for controlling oxidant input.
6. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the number of the high-pressure helium storage tanks (5) is 4, and the high-pressure helium storage tanks are uniformly distributed on the periphery of the thrust chamber shell (6) along the circumferential direction and used for pressurizing the oxidant storage tanks.
7. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the secondary flow pipeline (32) is an L-shaped pipe, the long arm end of the secondary flow pipeline is arranged on the outer wall surface of the spray pipe along the axial direction of the expansion section of the spray pipe, and the short arm end of the secondary flow pipeline is inserted into a nozzle of the spray pipe; the upper ends of the long arms are communicated with the thrust vector control pipeline (30) after being communicated through the annular pipeline.
8. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the number of the secondary flow pipelines (32) is 4, and the secondary flow pipelines are uniformly distributed along the circumferential direction.
9. The Mars aircraft solid-liquid hybrid engine of claim 1, characterized in that: the thrust vector control system further comprises secondary flow pipeline control valves (21), and on-off of each secondary flow pipeline (32) is informed through the secondary flow pipeline control valves (21).
10. A method for supplying an oxidant to a solid-liquid hybrid engine for a mars-ascent aircraft according to any one of claims 1 to 9, comprising the steps of:
the method comprises the following steps: after an engine standby instruction is given, opening a pressure increasing control valve (19), increasing pressure in the oxidant storage tank (1), and enabling the system to enter a standby state;
step two: after a formal working instruction is issued, the boron/potassium nitrate igniter (3) is started to generate high-temperature fuel gas, the extrusion system control valve (23) is opened after 0.5s, the paraffin fuel (11) is ignited under the action of the high-temperature fuel gas and nitrous oxide, the ignition and the starting of the engine are finished, and at the moment, the engine adopts an extrusion type oxidant supply mode;
step three: after ignition is started, a cooling circulation pipeline control valve (20) and a circulation loop pipeline control valve (22) are opened, nitrous oxide flows through an interlayer in a regeneration cooling spray pipe (7) to absorb heat and then enters a turbine, a turbine pump (2) is started, a pump pressure system control valve (24) is opened, and the turbine pump (2) absorbs nitrous oxide from an oxidant storage tank (1) and supplies the nitrous oxide to a thrust chamber; the turbine pump works stably, after the pressure in the combustion chamber is monitored to be stable, the control valve (23) of the extrusion system is closed, and at the moment, the engine adopts an expansion circulation type oxidant supply mode;
step four: if the attitude of the riser needs to be adjusted in a certain direction in the flying process, one or two secondary flow pipeline control valves (21) on the four groups of secondary flow pipelines (32) are opened, the flow of the oxidant entering the nozzle expansion section is controlled by the vector control flow regulator (27), nitrous oxide enters the nozzle expansion section, is rapidly atomized and evaporated and is mixed with fuel gas to interfere the supersonic gas flow to generate oblique shock waves, so that the supersonic gas flow is deflected in the direction to generate lateral force, and the attitude adjustment of the riser is realized.
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| CN114030656B (en) * | 2021-11-09 | 2022-08-05 | 西安交通大学 | Novel variable-thrust nuclear heat propulsion system |
| CN114084378B (en) * | 2021-11-16 | 2023-09-26 | 中国人民解放军国防科技大学 | Microwave heating water propulsion system and propulsion control method |
| CN114278464B (en) * | 2021-12-03 | 2023-04-14 | 西北工业大学太仓长三角研究院 | Self-heat-dissipation microminiature rocket propulsion device based on liquid fuel |
| CN114263548B (en) * | 2021-12-22 | 2022-07-12 | 宁波天擎航天科技有限公司 | Solid-liquid mixed engine and aircraft |
| CN118375532B (en) * | 2024-06-24 | 2024-08-23 | 北京航空航天大学 | Turbine pump pressure type solid-liquid hybrid rocket engine system |
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