CN112012850B - Method for improving performance of vortex combustion cold wall engine - Google Patents
Method for improving performance of vortex combustion cold wall engine Download PDFInfo
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- CN112012850B CN112012850B CN202010862813.7A CN202010862813A CN112012850B CN 112012850 B CN112012850 B CN 112012850B CN 202010862813 A CN202010862813 A CN 202010862813A CN 112012850 B CN112012850 B CN 112012850B
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- combustion chamber
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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
- F02K9/42—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
- F02K9/64—Combustion or thrust chambers having cooling 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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/52—Injectors
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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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/56—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
- F02K9/42—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
- F02K9/66—Combustion or thrust chambers of the rotary type
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Combustion Methods Of Internal-Combustion Engines (AREA)
Abstract
The invention belongs to the technical field of aerospace, provides a method for improving the performance of a vortex combustion cold wall engine and provides a position and a mode for jetting another component. The method compares the stoichiometric coefficients of reactants in a total reaction formula of combustion of a fuel component and an oxidant component, selects a component with larger volume flow to carry out tangential injection, and nozzles for injecting the other component are uniformly distributed at the head part of a combustion chamber along the circumferential direction, the axis of the nozzle is parallel to the axis of an engine, and the diameter of a distribution circle of the nozzle is 0.72 times of the diameter of the combustion chamber. The invention has the effect and benefit that the temperature of the fuel gas at the inner side of the jet pipe of the vortex combustion cold wall engine can be obviously reduced, and meanwhile, the performance of the engine is not reduced.
Description
Technical Field
The invention belongs to the technical field of aerospace, relates to a liquid rocket engine design technology, and particularly relates to an engine design technology for carrying out heat insulation and cooling on a wall surface by using vortex flow in a combustion chamber.
Background
Liquid rocket engines typically employ a cylindrical combustion chamber with a planar injector at the head. And fuel and oxidant are sprayed into the combustion chamber from the head injector through the nozzle, combustion reaction is carried out in the combustion chamber to generate high-temperature fuel gas, and the high-temperature fuel gas flows through the Laval nozzle to be accelerated and generate thrust. In liquid rocket engine combustors, the combustion temperature is high, for example, the combustion temperature of oxyhydrogen reaches over 3000 ℃, which far exceeds the melting point of many metals, so the combustor must be made of high temperature resistant materials and adopt special cooling technology. The extremely high thermal load not only increases the cost of the engine, but also makes it difficult to reliably operate the engine for long periods of time.
In 2002, american rail technology corporation issued a new rocket engine combustion chamber solution, called a vortex combustion cold wall engine. The wall temperature of the combustion chamber can reach a very low level, the wall temperature of the tested hydrogen-oxygen combustion chamber can reach 300 ℃, and transparent acrylic resin can be used as a combustion chamber material. Such a low wall temperature level not only greatly reduces the manufacturing cost of the thrust chamber, but also easily achieves the purposes of prolonging the service life, being reusable and improving the safety.
The scheme of the existing vortex combustion cold wall engine is that oxidant gas forms a double vortex flow structure in a combustion chamber to protect the side wall surface of the combustion chamber, and the principle is as follows: oxidant gas is tangentially sprayed from the side wall of the combustion chamber at the part of the cylindrical combustion chamber close to the spray pipe, and a vortex which is tightly attached to the side wall and flows to the head is formed at the inner side of the side wall; the oxidant gas turns the axial flow direction after reaching the head of the combustion chamber through vortex flow, and is mixed with the fuel sprayed at the axial flow direction for combustion to start generating fuel gas; the gas forms another vortex with opposite axial movement direction in the outer layer vortex, and continuously carries out combustion reaction, and finally flows to the spray pipe; since the oxidant gas in the outer vortex is not yet burnt and has a low temperature, it separates the wall from the hot gas in the inner vortex, so that the temperature of the side wall of the combustion chamber can be kept low.
Although the scheme can effectively reduce the temperature of the side wall surface of the combustion chamber, the temperature of the fuel gas at the position of the nozzle of the engine is still higher than the melting point of metal, and the engineering application of the engine is limited. In order to reduce the nozzle temperature, some researchers have approached the fuel nozzle at the head toward the axis while increasing the flow of oxidant, but both of these measures reduce engine performance and are undesirable.
Disclosure of Invention
The invention provides a method for selecting a side wall surface vortex propellant component aiming at a vortex combustion cold wall engine, provides a position and a mode for injecting another component, and effectively reduces the temperature of a jet pipe of the vortex combustion cold wall engine on the premise of not reducing the performance of the engine.
The technical scheme of the invention is as follows:
a method for improving the performance of a vortex combustion cold wall engine comprises the following steps:
the combustion chamber 1 of the vortex combustion cold wall engine is of a cylindrical structure, the diameter of the inlet of the spray pipe 2 is about 0.72 times of that of the combustion chamber, and an annular plane 3 is formed at the outlet of the combustion chamber; the tangential straight-flow nozzles 4 are uniformly arranged on the side wall surface of the combustion chamber close to the spray pipe along the circumferential direction; the combustion chamber head part 5 is in a plane or an ellipsoid shape and is smoothly connected with the cylindrical section 1; the nozzles 6 are uniformly arranged on the head 5;
the flow-through nozzle 4 is intended for tangentially spraying gaseous oxidant components or gaseous fuel components on the side wall of the combustion chamber, and the nozzle 6 is intended for spraying propellant components that are distinct from the flow-through nozzle 4.
The selection of the side wall surface vortex propellant component is to select the component sprayed by the direct current nozzle 4, and when the two propellant components are both gases, the selection is carried out according to the following method:
(1) writing a general expression of the general packet reaction of the combustion of propellant component A and propellant component B
aA+bB＝cC+dD
Where C, D denotes the product (the amount of which follows the actual case), a, b, c, d are the corresponding stoichiometric coefficients;
(2) if a is larger than b, the direct current nozzle 4 injects propellant A; if a < B, the nozzle 4 injects propellant B; if a is b, the molecular weight of the jet A, B from the dc nozzle 4 is large.
The position and the mode of injection of the other component are the nozzles 6, and the arrangement mode is as follows:
(1) the number of the nozzles is 6-8, and the nozzles are uniformly distributed at the head of the combustion chamber along the circumferential direction;
(2) the diameter of the distribution circle of the nozzle is 0.72 times of the diameter of the cylindrical section of the combustion chamber;
(3) the nozzle axis is parallel to the combustion chamber axis and extends into the combustion chamber for a length 1/10 the diameter of the combustion chamber.
The theoretical basis of the invention is as follows: in vortex cold-wall combustion chambers, outsideThe stability of the layer vortex plays a leading and controlling role in combustion and cooling, so a propellant component with larger volume flow rate is selected to form the outer layer vortex, and an oxidant is not fixedly used; the theoretical radius of the interface between the inner and outer vortices is 2 of the diameter of the combustion chamber0.5Approximately 0.72 times, injection of another propellant component in the vicinity of this is advantageous to enhance the mixing of the reactants and has less effect on the outer layer vortex and the head flow.
The invention has the advantages of fully playing the advantages of the vortex combustion cold wall engine on the premise of not obviously reducing the performance of the engine, effectively reducing the temperature of the fuel gas at the inner side of the spray pipe, reducing the cooling requirement of the spray pipe and even not needing additional cooling.
Drawings
FIG. 1(a) is a front view of a vortex combustion cold wall engine.
FIG. 1(b) is a side view of a vortex combustion cold wall engine.
Fig. 2 is a temperature distribution diagram in kelvin (K) of the engine outer wall surface in example 1.
In the figure: 1, a combustion chamber cylindrical section; 2, spraying a pipe; 3, a ring plane; 4, a direct current nozzle; 5 a combustion chamber head; 6, a nozzle.
Detailed Description
The following detailed description of the embodiments of the invention is provided in connection with the accompanying drawings.
Example 1:
a vortex combustion cold wall engine is designed, the pressure of a combustion chamber is 1MPa, a propellant is oxygen and hydrogen, and the mass ratio of the oxygen to the hydrogen is 6: 1.
1) The nozzle throat diameter was determined to be 12mm according to engine design theory and the total flow of propellant (oxygen and hydrogen) was 45.4 kg/s.
2) The inner diameter of the combustion chamber is 40mm, and the inner length is 40 mm; the head of the combustion chamber is in an ellipsoid shape, the long axis of the combustion chamber is 40mm, the short axis of the combustion chamber is 20mm, and the ellipsoid is tangent to the cylindrical section of the combustion chamber; the diameter of the inlet of the spray pipe is 28mm, the spray pipe is connected with the cylindrical section of the combustion chamber through an annular plane, the convergence half angle of the spray pipe is 45 degrees, the diameter of the outlet of the spray pipe is 14mm, and the distance between the outlet and the throat part is 2 mm.
3)Write the total envelope reaction equation 8H for oxygen and hydrogen combustion at a mass ratio of 6:12+3O2＝6H2O+2H2Since the stoichiometric coefficient 8 of hydrogen is greater than the stoichiometric coefficient 3 of oxygen, hydrogen is selected as the injection component of the combustion chamber vortex, and the hydrogen is injected tangentially to the side wall of the combustion chamber, while the oxygen is injected from the nozzle at the head.
4) 4 hydrogen nozzles with the diameter of 2mm are arranged on the side wall surface of the combustion chamber, and the axis of each nozzle is tangent to the inner wall surface of the combustion chamber and is 2mm away from the annular plane of the outlet of the combustion chamber.
5) 8 oxygen nozzles are arranged on the circumference of the head part of the combustion chamber with the diameter of 40mm multiplied by 0.72 which is 28mm, the axis is parallel to the axis of the engine, and the length of the oxygen nozzles extending into the combustion chamber is 40mm/10 which is 4 mm.
6) And performing flow field numerical calculation on the designed combustion chamber, and calculating to obtain that the highest gas temperature on the side wall surface of the combustion chamber is 400K, the highest temperature on the inner wall surface of the spray pipe is 1400K, and the highest temperature on the inner wall surface of the head part is 600K, so that the aim of cooling the spray pipe part is fulfilled.
Claims (2)
1. A method for improving the performance of a vortex combustion cold wall engine is characterized in that when two propellant components are both gases, the method is selected as follows:
1) writing a general packet reaction expression of propellant component A and propellant component B combustion
aA+bB＝cC+dD
Where C, D denotes the product, the amounts of which follow the actual situation, a, b, c, d are the corresponding stoichiometric coefficients;
2) if a is larger than b, the direct current nozzle (4) injects propellant A; if a < B, the direct current nozzle (4) injects propellant B; when a is b, the molecular weight of the jet A, B is large in the dc nozzle (4).
2. Method for improving the performance of a vortex combustion cold wall engine according to claim 1, characterized in that the nozzles (6) are arranged in such a way that:
1) the number of the nozzles (6) is 6-8, and the nozzles are uniformly distributed at the head of the combustion chamber along the circumferential direction;
2) the diameter of a distribution circle of the nozzle (6) is 0.72 times of the diameter of the cylindrical section of the combustion chamber;
3) the axis of the nozzle (6) is parallel to the axis of the combustion chamber, and the length of the nozzle extending into the combustion chamber is 1/10 times the diameter of the combustion chamber.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
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| CN202010862813.7A CN112012850B (en) | 2020-08-25 | 2020-08-25 | Method for improving performance of vortex combustion cold wall engine |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202010862813.7A CN112012850B (en) | 2020-08-25 | 2020-08-25 | Method for improving performance of vortex combustion cold wall engine |
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| CN112012850A CN112012850A (en) | 2020-12-01 |
| CN112012850B true CN112012850B (en) | 2021-07-16 |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN85103360A (en) * | 1985-05-13 | 1986-12-10 | 奥斯特博 | Gas turbulator |
| US6298659B1 (en) * | 1999-03-24 | 2001-10-09 | Orbital Technologies Corporation | Vortex flow field and apparatus and method for producing the same |
| US20080264035A1 (en) * | 2007-04-25 | 2008-10-30 | Ricciardo Mark J | Coolant flow swirler for a rocket engine |
| CN201196120Y (en) * | 2008-04-22 | 2009-02-18 | 聂振傲 | Hollow shaft turbojet engine |
| CN102207043B (en) * | 2011-04-27 | 2013-06-19 | 北京航空航天大学 | Gaseous hydrogen/gaseous oxygen eddy current cooling thrust chamber injector |
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