CN109882886B - RBCC engine combustion chamber with slope rocket layout mode and design method thereof - Google Patents
RBCC engine combustion chamber with slope rocket layout mode and design method thereof Download PDFInfo
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Abstract
A ramp type rocket layout mode RBCC combustion chamber and its design method, the combustion chamber is from the air inlet to the gas outlet that the burning produces is the isolation section, rocket slope installation section, expansion section sequentially; the gas outlet from the air inlet to the combustion is sequentially provided with an isolation section, a rocket slope installation section and an expansion section; the rocket thrust chamber is arranged at the rocket slope installation section, so that rocket jet flow is in contact with three sides of the main ram flow and is positioned in front of the flame front of the expansion section, ignition and flame stabilization of the combustion chamber are realized by using the rocket jet flow, and the problem that the arrangement mode of a step rocket cannot effectively utilize high-temperature high-enthalpy active jet flow ignition and flame stabilization of the rocket is solved.
Description
Technical Field
The invention relates to a combustion device for a Combined power system of a hypersonic aircraft, in particular to a combustion device for realizing high-efficiency conversion of chemical energy of fuel into internal energy of airflow by a Rocket Based Combined Cycle (RBCC) engine.
Background
The rocket layout mode has important influence on the structure, the working range and the performance of the combustion chamber of the RBCC engine. At present, the rocket layout modes of the combustion chamber of the RBCC engine adopted at home and abroad are divided into two types: 1) step rocket overall arrangement, rocket arrange in combustion chamber one side, rocket efflux and punching press mainstream unilateral contact, two air currents are relative decoupling zero in the pneumatics for combustion chamber design and intake duct design are relatively independent, have reduced the combustion chamber design degree of difficulty. But the step rocket layout mode does not effectively utilize the high-temperature and high-enthalpy active jet ignition and flame stabilization capabilities of the rocket, and the structural utilization rate of the combustion chamber is low. 2) The support plate rocket layout is adopted, the rocket is arranged in the center of the combustion chamber, the rocket jet flow is in bilateral contact with the main ram flow, the two air flows are in close pneumatic coupling, the ignition and flame stabilization of the combustion chamber can be realized by utilizing the high-temperature and high-enthalpy active jet flow of the rocket, the structural utilization rate of the combustion chamber is high by adopting the support plate rocket layout mode, and the length of the combustion chamber is generally small. However, the combustion chamber adopts a support plate rocket layout, and the design coupling of the combustion chamber and an air inlet channel is high, so that the design of the engine lacks flexibility.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method overcomes the contradiction that the existing RBCC engine combustion chamber rocket layout mode cannot give consideration to the realization of combustion chamber ignition and flame stabilization by using high-temperature high-enthalpy active jet of the rocket and no constraint on the type of an air inlet, provides a slope rocket layout mode for the RBCC engine combustion chamber, can realize a scheme of combustion chamber ignition, rocket jet and rocket step double-flame stabilization by using the high-temperature high-enthalpy active jet of the rocket, and is suitable for all types of air inlets.
The technical solution of the invention is as follows: a RBCC combustion chamber with a slope type rocket layout mode is characterized in that a gas outlet generated by combustion from an air inlet is sequentially provided with an isolation section, a rocket slope installation section and an expansion section;
the rocket thrust chamber is arranged at the rocket slope installation section, so that rocket jet flow is in contact with three sides of the main ram flow and is positioned in front of the flame front of the expansion section, and ignition and flame stabilization of the combustion chamber are realized by the rocket jet flow.
Preferably, the rocket mounting slope section is located between the isolation section and the expansion section, the main body is the expansion section, a rocket mounting slope is arranged in the center of the flow channel of the expansion section of the main body from the exit position of the isolation section along the airflow direction, the rocket mounting slope is wedge-shaped, and the rocket thrust chamber is mounted on the slope.
Preferably, the expansion angle alpha 2 of the main body expansion section of the rocket mounting slope section ranges from 5 degrees to 10 degrees.
Preferably, the divergence angle 2 of the slope ranges from 5 to 10 °.
Preferably, the compression angle 1 of the rocket installation slope is in the range of 10-15 degrees.
Preferably, the length L1 of the isolation section is 6-10H, and the expansion angle α 1 of the isolation section is 0-2 °; and H is the height of the inlet of the isolating section.
Preferably, the number of stages of the expansion section is determined according to the range of the flying mach number of the combustor.
Preferably, for an RBCC combustion chamber with a flight Mach number of Ma0-7, two-stage expansion sections are adopted, wherein the length L3 of the one-stage expansion section is 6-10H, and the expansion angle alpha 3 is 2-4 degrees; the length L4 of the secondary expansion section is 6-10H, and alpha 4 is 3-5 degrees; and H is the height of the inlet of the isolating section.
Preferably, the rocket is a variable working condition rocket, the rocket works under the maximum working condition of the flying Mach number Ma being 0-4, and thermodynamic choked tissue combustion is realized in the secondary expansion section; the flying Mach number Ma is 3.5-5.5, the rocket jet flow is closed, and thermodynamic choked tissue combustion is realized in the secondary expansion section; the flying Mach number Ma is 5-7, the rocket works under the minimum working condition, the tissue center burns in the first-stage expansion section, and thermal congestion is realized.
A design method of a combustion chamber of a slope rocket layout type RBCC engine is realized by the following steps:
(1) determining the height H and the width W of an inlet of an isolation section according to a design result of an air inlet channel;
(2) determining the length L1 of the isolation section according to the height H of the inlet of the isolation section; selecting an expansion angle alpha 1 of the isolation section according to the correction result of the boundary layer of the isolation section;
(3) determining rocket installation slope structure parameters according to the installation size of the thrust chamber: length L2, compression angle 1, expansion angle 2, such that the rocket jet makes three-sided contact with the ram main flow;
(4) determining an expansion angle alpha 2 of the rocket ramp installation section according to the rocket installation ramp structure parameters in the step (3), wherein the expansion angle alpha 2 at least meets the condition that the flow area of each section of the rocket ramp installation section along the airflow flowing direction is consistent with the area of an outlet of the isolation section;
(5) determining the expansion stage number of the combustion chamber according to the combustion organization idea of graded thermal choke and the combination of the range of the flying Mach number of the combustion chamber from 0 to 7;
(6) determining the length L3 and the expansion angle alpha 3 of the primary expansion section according to the upper boundary of the combustion chamber flight Mach number Ma; and determining the length L4 of the secondary expansion section and the expansion angle alpha 4 according to the lower boundary of the flight Mach number Ma of the combustion chamber.
Compared with the prior art, the invention has the beneficial effects that:
the rocket thrust chamber is arranged on a wall surface slope of a combustion chamber, rocket jet flow is in contact with three sides of a main ram flow, two air flows are pneumatically coupled, ignition of the combustion chamber is realized by using high-temperature high-enthalpy active jet flow of the rocket, and a double-flame stabilization scheme of the rocket jet flow and the rocket slope is adopted; the combustion chamber has smooth channels, avoids the problem of aerodynamic flow in a multi-channel supersonic velocity caused by a support plate rocket layout mode, and can be suitable for any type of air inlet channel.
The rocket thrust chamber is arranged at the rocket slope installation section, so that rocket jet flow is in contact with three sides of the main ram flow, ignition and flame stabilization of the combustion chamber are realized by using the rocket jet flow, and the problem that the high-temperature and high-enthalpy active jet flow ignition and flame stabilization of the rocket cannot be effectively utilized in a step rocket layout mode is solved. The rocket thrust chamber is arranged on the wall surface slope of the combustion chamber, the design of the combustion chamber is relatively independent from that of the air inlet channel, the channel of the combustion chamber is smooth, and the problem of multi-channel supersonic velocity internal flow starting caused by the layout mode of the central support plate rocket is solved, so that the rocket thrust chamber can adapt to any type of air inlet channel. The invention adopts a rocket high-enthalpy active jet ignition, rocket jet and rocket slope starting/mechanical double-flame stabilization scheme and a multistage thermal choked combustion organization, and solves the problems of wide-range efficient and stable combustion of the RBCC combustion chamber and reliable working under extreme and severe working conditions.
Drawings
Fig. 1-3 are schematic structural diagrams of a combustion chamber of an RBCC engine in a ramp rocket layout mode.
Detailed Description
The invention is described in detail below with reference to the figures and examples.
As shown in fig. 1-3, a ramp type rocket layout type RBCC combustion chamber comprises an isolation section 1, a rocket ramp installation section 2, a primary expansion section 3, a secondary expansion section 4 and a rocket installation ramp 5; the combustion chamber sequentially comprises from an air inlet to a gas outlet generated by combustion: the rocket comprises an isolation section, a rocket slope installation section, a primary expansion section and a secondary expansion section; the rocket mounting slope is positioned at the rocket slope mounting section. The invention adopts the combustion organization idea of graded thermal choke to realize the wide-range reliable ignition and flame stabilization of the combustion chamber. The number of stages of the expansion section is determined according to the range of the flying Mach number of the combustion chamber, and the expansion section is generally in a two-stage structure, and the two-stage expansion section structure is provided in the example.
The rocket mounting slope section is located between the isolation section and the expansion section, the main body is an expansion section, a rocket mounting slope is arranged in the center of a flow channel of the expansion section of the main body from the junction position of the main body and an outlet of the isolation section along the airflow direction, the rocket mounting slope is wedge-shaped, and the rocket thrust chamber is mounted on the slope. Through the arrangement mode of the slope rocket, rocket jet flow is in contact with three sides of a main ram flow, two air flows are pneumatically coupled, ignition and flame stabilization of an RBCC combustion chamber are realized by utilizing high-temperature high-enthalpy active gas jet flow of a rocket thrust chamber, and the device is suitable for all types of air inlets.
The combustion chamber adopts a rocket jet ignition, rocket slope and rocket jet double-flame stabilization scheme, and two-stage thermodynamic choked tissue combustion. The working range of the combustion chamber is the flight Mach number Ma which is 0-7. Working under the working condition that the flight Mach number Ma is 0-4 and the maximum working condition of the rocket works, adopting a rocket jet ignition, rocket slope and rocket jet double-flame stabilization scheme, and realizing thermodynamic choked tissue combustion in a secondary expansion section; the flying Mach number Ma is 3.5-5.5, the rocket jet flow is closed, the scheme of rocket jet flow ignition and rocket slope single stable flame is adopted, and thermodynamic choked tissue combustion is realized in a secondary expansion section; the flying Mach number Ma is 5-7, the rocket jet works under a small working condition, a rocket jet ignition scheme, a rocket slope scheme and a rocket jet double-flame stabilization scheme are adopted, the central combustion is organized in a first-stage expansion section, and thermal congestion is realized.
The combustion chamber with the structure determines each structural parameter through the following steps:
(1) determining the height H and the width W of an inlet of an isolation section according to a design result of an air inlet channel;
(2) determining the length L1 of the isolation section according to the height H of the inlet of the isolation section; selecting an expansion angle alpha 1 of the isolation section according to the correction result of the boundary layer of the isolation section;
(3) determining rocket installation slope structure parameters according to the installation size of the thrust chamber: length L2, compression angle 1, expansion angle 2, such that the rocket jet makes three-sided contact with the ram main flow;
(4) and (4) determining an expansion angle alpha 2 of the rocket ramp installation section according to the rocket installation ramp structure parameters in the step (3), wherein the expansion angle alpha 2 at least meets the condition that the flow area of each section of the rocket ramp installation section along the airflow flowing direction is consistent with the area of an outlet of the isolation section.
(5) And determining the expansion stage number of the combustor according to the flight Mach number range of the combustor. For the RBCC combustion chamber with the range of Ma0-7, the two-stage expansion combustion chamber can realize wide-range reliable ignition and flame stabilization. A wider range of combustion chambers, depending on the particular operating range of the combustion chamber.
(6) For an RBCC combustor in the range of Ma0-7, determining the length L3 of a first-stage expansion section and an expansion angle alpha 3 according to the upper boundary of the flight Mach number Ma of the combustor;
(7) for RBCC combustors in the range of Ma0-7, the length L4 and the divergence angle alpha 4 of the secondary expansion section are determined according to the lower boundary of the flight Mach number Ma of the combustors.
Example 1
For an RBCC combustor in the Ma0-7 range:
(1) according to the design result of the air inlet channel, determining that the height H of the inlet of the combustion chamber is 50mm, and the width W of the inlet of the combustion chamber is 200 mm;
(2) selecting the length L1 of the isolation section to be 500mm according to the height H of the inlet of the combustion chamber; and selecting the expansion angle alpha 1 of the isolation section to be 1.0 degrees according to the correction calculation result of the boundary layer of the isolation section.
(3) According to the size of a rocket thrust chamber provided by the engine, the length L2 of a rocket installation slope is determined to be 200mm, the compression angle 1 is 11.4 degrees, and the expansion angle 2 is determined to be 11.4 degrees;
(4) determining that the rocket slope installation expansion angle alpha 2 is 5.7 degrees according to the rocket slope length L2, the compression angle 1 and the expansion angle 2;
(5) determining the length L3 of the primary expansion segment to be 500mm and the expansion angle alpha 3 to be 2.3 according to the upper boundary (Ma to be 6) of the number of the flying Ma of the combustion chamber;
(6) and determining the length L4 of the secondary expansion section to 465mm and the expansion angle alpha 4 to 3.7 degrees according to the lower boundary (Ma to 4) of the number of the flying Ma of the combustion chamber.
(7) The combustion chamber designed according to the design parameters can realize reliable ignition and stable flame in a wide range of the flight Mach number Ma0-7, and the combustion efficiency is not less than 0.85.
The invention has not been described in detail in part of the common general knowledge of those skilled in the art.
Claims (8)
1. A ramp type rocket layout mode's RBCC combustion chamber which characterized in that: the gas outlet from the air inlet to the combustion is sequentially provided with an isolation section, a rocket slope installation section and an expansion section;
the rocket thrust chamber is arranged at the rocket slope installation section, so that rocket jet flow is in contact with three sides of the main ram flow and is positioned in front of the flame front of the expansion section, and ignition and flame stabilization of the combustion chamber are realized by the rocket jet flow; the main body of the rocket slope installation section is an expansion section, a rocket installation slope is arranged in the center of a flow channel of the main body expansion section from an outlet of the isolation section along the airflow direction, the rocket installation slope is wedge-shaped, and a rocket thrust chamber is installed on the slope.
2. A ramped rocket layout RBCC combustor as claimed in claim 1 wherein: the expansion angle alpha 2 of the main body expansion section of the rocket mounting slope section ranges from 5 degrees to 10 degrees.
3. A ramped rocket layout RBCC combustor as claimed in claim 1 wherein: the divergence angle 2 of the slope ranges from 5 to 10 degrees.
4. A ramped rocket layout RBCC combustor as claimed in claim 1 wherein: the compression angle 1 of the rocket installation slope ranges from 10 degrees to 15 degrees.
5. A ramped rocket layout RBCC combustor as claimed in claim 1 wherein: the length L1 of the isolation section is 6-10H, and the expansion angle alpha 1 of the isolation section is 0-2 degrees; and H is the height of the inlet of the isolating section.
6. A ramped rocket layout RBCC combustor as claimed in claim 1 wherein: the stage number of the expansion section is determined according to the range of the flying Mach number of the combustion chamber; for an RBCC combustion chamber with a flight Mach number of Ma0-7, two stages of expansion sections are adopted, wherein the length L3 of the first stage expansion section is 6-10H, and the expansion angle alpha 3 is 2-4 degrees; the length L4 of the secondary expansion section is 6-10H, and alpha 4 is 3-5 degrees; and H is the height of the inlet of the isolating section.
7. A ramped rocket layout RBCC combustor as claimed in claim 6, wherein: the rocket is a variable working condition rocket, the rocket works under the maximum working condition of the flying Mach number Ma being 0-4, and thermodynamic choked tissue combustion is realized in the secondary expansion section; the flying Mach number Ma is 3.5-5.5, the rocket jet flow is closed, and thermodynamic choked tissue combustion is realized in the secondary expansion section; the flying Mach number Ma is 5-7, the rocket works under the minimum working condition, the tissue center burns in the first-stage expansion section, and thermal congestion is realized.
8. A design method of a combustion chamber of a slope rocket layout type RBCC engine is characterized by being realized by the following steps:
(1) determining the height H and the width W of an inlet of an isolation section according to a design result of an air inlet channel;
(2) determining the length L1 of the isolation section according to the height H of the inlet of the isolation section; selecting an expansion angle alpha 1 of the isolation section according to the correction result of the boundary layer of the isolation section;
(3) determining rocket installation slope structure parameters according to the installation size of the thrust chamber: length L2, compression angle 1, expansion angle 2, such that the rocket jet makes three-sided contact with the ram main flow;
(4) determining an expansion angle alpha 2 of the rocket ramp installation section according to the rocket installation ramp structure parameters in the step (3), wherein the expansion angle alpha 2 at least meets the condition that the flow area of each section of the rocket ramp installation section along the airflow flowing direction is consistent with the area of an outlet of the isolation section;
(5) determining the expansion stage number of the combustion chamber according to the combustion organization idea of graded thermal choke and the combination of the range of the flying Mach number of the combustion chamber from 0 to 7;
(6) determining the length L3 and the expansion angle alpha 3 of the primary expansion section according to the upper boundary of the combustion chamber flight Mach number Ma; and determining the length L4 of the secondary expansion section and the expansion angle alpha 4 according to the lower boundary of the flight Mach number Ma of the combustion chamber.
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| CN110307987B (en) * | 2019-06-19 | 2020-07-03 | 西北工业大学 | Variable structure rocket base combined power cycle combustion chamber experimental device |
| CN112747335B (en) * | 2021-01-06 | 2022-05-13 | 中国人民解放军国防科技大学 | Lobe type backward step structure, lobe type concave cavity device and afterburning chamber |
| CN114484503B (en) * | 2022-01-05 | 2023-03-21 | 中国科学院力学研究所 | Self-adaptive geometric throat combustion chamber of wide-range ramjet engine |
| CN115434823A (en) * | 2022-08-31 | 2022-12-06 | 西安航天动力研究所 | Rocket stamping combined engine with parallel compressor runners |
| CN116335852B (en) * | 2023-02-07 | 2023-09-01 | 中国空气动力研究与发展中心空天技术研究所 | Stamping engine tail nozzle of integrated enhanced rocket and design and working methods thereof |
| CN116379477B (en) * | 2023-04-20 | 2024-08-27 | 中国人民解放军国防科技大学 | Slope concave cavity combustion chamber and scramjet engine |
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| US6584774B1 (en) * | 2001-10-05 | 2003-07-01 | The United States Of America As Represented By The Secretary Of The Air Force | High frequency pulsed fuel injector |
| CN103711610B (en) * | 2013-12-18 | 2016-01-20 | 中国航天科技集团公司第六研究院第十一研究所 | A kind of integration of the RBCC gas generator based on liquid oxygen supply and regulating system |
| CN103726954B (en) * | 2013-12-23 | 2015-11-18 | 西北工业大学 | The T-shaped layout of a kind of Rocket based combined cycle motor Rocket ejector |
| JP6736998B2 (en) * | 2016-06-16 | 2020-08-05 | 株式会社Ihi | Combustor liner |
| CN106765311B (en) * | 2016-12-13 | 2019-04-09 | 北京航空航天大学 | A kind of ultra-combustion ramjet combustion chamber supporting plate with right angle trigonometry connected in star |
| CN106907272B (en) * | 2017-03-23 | 2019-01-01 | 西北工业大学 | Structure changes Rocket based combined cycle engine |
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