Wide operating condition liquid rocket engine thrust chamber

Abstract

The present application relates to the technical field of spacecraft liquid rocket engine thrust chamber design, and provides a wide working condition liquid rocket engine thrust chamber, which comprises a combustion chamber and a direct impingement injector, the combustion chamber comprises a straight section, a converging section, a throat and an expansion section connected in sequence, the direct impingement injector is arranged at the end of the straight section, the length-diameter ratio of the straight section is 1.2-1.5, the characteristic length is 290-350 mm, the direct impingement injector is provided with a central zone first injection hole, a central zone second injection hole and a peripheral zone injection hole, the oxidizer is sprayed out through the central zone first injection hole, and the fuel is sprayed out through the central zone second injection hole and the peripheral zone injection hole.The present application greatly widens the working condition range that the thrust chamber can adapt to, and solves the problems that the throat temperature of a high specific impulse performance space engine is too high and is greatly affected by working condition deviation.

Description

Technical Field

[0001] This invention relates to the field of spacecraft liquid rocket engine thrust chamber design technology, specifically, to a wide-condition liquid rocket engine thrust chamber. Background Technology

[0002] In the traditional space dual-element attitude control liquid rocket engine thrust chamber design for spacecraft, in pursuit of high specific impulse performance, the high-temperature zone of the combustion chamber is located in the throat region where the heat flux density is highest, with the outer wall temperature even reaching over 1400℃. The ablation zone at the end of its lifespan is also located in the throat. However, this has the following consequences:

[0003] (1) The throat bears the highest heat flux density, the highest temperature, and the fastest gas scouring speed, and is prone to ablation failure after long-term operation.

[0004] (2) The side area coolant film covers the front end of the straight section of the combustion chamber and is far from the throat. The liquid film coverage area is difficult to effectively extend to the throat area and is greatly affected by the deviation of the operating conditions. Especially under high mixture ratio conditions, the throat temperature rises significantly.

[0005] (3) Insufficient residence time of the gas in the combustion chamber leads to significant fluctuations in combustion efficiency due to changes in operating conditions.

[0006] Existing technologies alleviate the problem of excessive throat temperature by increasing the proportion of edge cooling flow or using edge jet cooling at multiple angles, such as the spatially staggered liquid film cooling structure and its flow calculation method disclosed in patent document CN116608058A. However, this can lead to loss of specific impulse performance or prolong the iteration cycle of design parameters, increasing the cost of trial verification. Summary of the Invention

[0007] In view of the deficiencies in the prior art, the purpose of this invention is to provide a thrust chamber for a liquid rocket engine with a wide operating condition.

[0008] The thrust chamber of a wide-condition liquid rocket engine provided by the present invention includes a combustion chamber and a DC mutual-impact injector;

[0009] The combustion chamber comprises a straight section, a converging section, a throat, and a diverging section connected in sequence. The direct-flow interlocking injector is disposed at the end of the straight section. The aspect ratio of the straight section is 1.2 to 1.5, preferably 1.3, and the characteristic length is 290 to 350 mm, preferably 320 mm. The aspect ratio is L = length of the straight section. c With inner diameter D c The ratio, wherein the characteristic length is the ratio of the volume of the combustion chamber to the cross-sectional area of ​​the throat;

[0010] The DC mutual-impact injector is equipped with a central region first injection hole, a central region second injection hole, and a side region injection hole. Oxidant is injected through the central region first injection hole, and fuel is injected through the central region second injection hole and the side region injection hole.

[0011] Preferably, the DC mutual-impact injector uses a combination of self-igniting propellants, with nitrogen tetroxide as the oxidant and methylhydrazine as the fuel.

[0012] Preferably, the cooling jet from the side zone is ejected through the side zone injection hole and then impacts the inner wall of the straight section of the combustion chamber for cooling.

[0013] Preferably, the edge zone cooling flow rate ratio is 20% to 30%, which is the ratio of the fuel mass flow rate flowing through the edge zone injection hole to the total fuel mass flow rate.

[0014] Preferably, the angle θ of the cooling jet impacting the wall in the edge region is 30° to 45°, for example 30°, or even 40°.

[0015] Preferably, the average Sotel particle size of the spray field formed by the DC mutual impact injector is 150-200 μm, and more preferably 180 μm.

[0016] Preferably, the design vacuum thrust of the thrust chamber under rated operating conditions is 100-300 N, the combustion chamber pressure is 0.8-1.0 MPa, and the mixture ratio is 1.65, wherein the mixture ratio is the ratio of oxidant to fuel mass flow rate.

[0017] Preferably, the first injection hole in the central area, the second injection hole in the central area, and the injection holes in the side area are all uniformly arranged in a single ring, with their distribution positions corresponding one-to-one radially, and the number N is the same, where N is 6 or 8.

[0018] Preferably, the inner diameter L of the straight segment c The inner diameter L of the straight segment is φ25~35mm. c It is 30 mm.

[0019] Preferably, the combustion chamber is made of niobium alloy and has a high-temperature anti-oxidation silicon compound coating on its inner and outer surfaces, and the DC mutual impact injector is made of 7715D high-temperature titanium alloy.

[0020] Compared with the prior art, the present invention has the following beneficial effects:

[0021] 1. This invention employs a smaller combustion chamber inner diameter and a larger length-to-diameter ratio of the straight section of the combustion chamber. Through structural parameter optimization and thermal protection matching design, the high-temperature zone is moved forward to the straight section of the combustion chamber where the heat flux density is lower and the edge coolant film directly covers it. This achieves a dynamic balance between the combustion chamber temperature field and the cooling system, enhances the hydrodynamic effect of the coolant film, reduces the sensitivity of combustion efficiency to changes in operating conditions, significantly widens the operating condition range that the thrust chamber can adapt to, and solves the problem of excessively high throat temperature in high-specific-impulse space engines, which is greatly affected by deviations in operating conditions.

[0022] 2. This invention employs a smaller combustion chamber inner diameter, reducing the radius of curvature of the combustion chamber and enhancing the hydrodynamic effect. Under the same cooling flow rate, it can significantly increase the coolant film velocity, forming a stable liquid film on the wall surface and increasing the local liquid film thickness. The reduced inner diameter of the combustion chamber decreases the area that the coolant film needs to cover, making it easier for the liquid film to cover the entire inner wall surface. It also promotes the merging of adjacent liquid films, forming a thicker cooling layer, while shortening the liquid film coverage distance and reducing evaporation loss.

[0023] 3. In this invention, the coolant film in the side zone can not only fully cool the wall surface, but also participate in the combustion in the central zone, ensuring high combustion efficiency. At the same time, it reduces the sensitivity of combustion efficiency to changes in operating conditions, greatly widening the range of operating conditions that the thrust chamber can adapt to. Especially under low operating conditions, it can still maintain a high liquid film coverage. Attached Figure Description

[0024] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0025] Figure 1 This is a schematic cross-sectional view of the thrust chamber of a wide-condition liquid rocket engine in an embodiment of the present invention. The injection surface shown in the figure is the outlet plane of a DC mutual-impact injector.

[0026] Figure 2 for Figure 1 A magnified view of part A in the middle;

[0027] Figure 3 This is a schematic diagram comparing the axial temperature variation curves of the combustion chamber in the embodiment of the present invention with those in a conventional design.

[0028] The diagram shows:

[0029] Combustion chamber 1;

[0030] Line segment 11;

[0031] Convergence segment 12;

[0032] Throat 13;

[0033] Expansion segment 14;

[0034] DC mutual impact injector 2;

[0035] Central area first injection hole 21;

[0036] Second injection hole 22 in the central area;

[0037] Side area injection hole 23;

[0038] Injection surface 24. Detailed Implementation

[0039] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0040] This invention provides a thrust chamber for a liquid rocket engine with wide operating conditions, such as... Figures 1 to 3 As shown, it includes a combustion chamber 1 and a DC mutual impact injector 2; the combustion chamber 1 includes a straight section 11, a converging section 12, a throat 13 and an expanding section 14, which are connected in sequence. The straight section 11 is a cylindrical structure and the expanding section 14 is a trumpet-shaped structure.

[0041] like Figure 1 As shown, the inner diameter D of straight line segment 11 c It is φ30mm and the length is L. c It has a length of 36mm, an aspect ratio of 1.2, a characteristic length of 320mm, and an aspect ratio equal to the length L of the straight section 11 of the combustion chamber. c With inner diameter D c The ratio of the characteristic length to the volume of the combustion chamber 1 to the cross-sectional area of ​​the throat 13 is given by the characteristic length.

[0042] The DC mutual-impact injector 2 is provided with a central area first injection hole 21, a central area second injection hole 22 and an edge area injection hole 23. The central area first injection hole 21, the central area second injection hole 22 and the edge area injection hole 23 are all uniformly arranged in a single ring, with their distribution positions corresponding one-to-one radially, and the number N is the same, where N is 8.

[0043] The system employs a combination of self-igniting propellants, with nitrogen tetroxide as the oxidizer and methylhydrazine as the fuel. The oxidizer is ejected through the first injection hole 21 in the central region, while the fuel is ejected through the second injection hole 22 in the central region and the injection hole 23 in the side region. The side of the DC mutual-impact injector 2 facing the straight section 11 forms an injection surface 24, and a concave structure is formed on the injection surface 24. The oxidizer ejected from the first injection hole 21 in the central region and the fuel ejected from the second injection hole 22 in the central region are mixed after passing through the concave structure.

[0044] Specifically, the side zone cooling jet is ejected through the side zone injection hole 23 and then impacts the combustion chamber wall for cooling. The side zone cooling flow rate ratio is 30%, and the side zone cooling jet impact angle θ is 30°. The side zone cooling flow rate ratio is the ratio of the fuel mass flow rate through the side zone injection hole 23 to the total fuel mass flow rate.

[0045] The design vacuum thrust of the thrust chamber under rated operating conditions is 160N, the combustion chamber pressure is 0.85MPa, and the mixture ratio is 1.65, which is the ratio of oxidant to fuel mass flow rate; the Sotel average particle size (SMD) of the spray field formed by the DC mutual impact injector 2 is 150μm.

[0046] This invention extends the residence time of the combustion gas in the straight section of the combustion chamber by using a larger length-to-diameter ratio. Combined with the DC mutual-impact injector and combustion chamber parameters, it reduces the temperature at the throat where the heat flux density of the combustion chamber is at its maximum, thus solving the problem of excessively high throat temperature in high specific impulse space engines and the significant impact of operating condition deviations.

[0047] In this embodiment, the substrate of combustion chamber 1 is niobium alloy, and high-temperature anti-oxidation silicide coating is prepared on the inner and outer surfaces. The substrate of DC mutual impact injector 2 is 7715D high-temperature titanium alloy.

[0048] The parameter combination in this embodiment was obtained through high-altitude simulated hot-fire testing of thrust chambers with various design parameters. Under this parameter combination, the thrust chamber can achieve stable and reliable operation under a wide range of working conditions, such as... Figure 3 As shown. The embodiments of the present invention can operate stably and reliably at 50% to 130% of the thrust design value and 65% to 135% of the rated mixture ratio. When the thrust or mixture ratio deviates from the design value by ±10%, the vacuum specific impulse retention rate is ≥97%.

[0049] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0050] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A thrust chamber for a wide-condition liquid rocket engine, characterized in that, It includes a combustion chamber (1) and a direct current mutual impact injector (2); The combustion chamber (1) comprises a straight section (11), a converging section (12), a throat (13), and an expanding section (14) connected in sequence. The DC-DC interlocking injector (2) is disposed at the end of the straight section (11). The length-to-diameter ratio of the straight section (11) is 1.2 to 1.5, and the characteristic length is 290 to 350 mm. The length-to-diameter ratio is the length L of the straight section (11). c With inner diameter D c The ratio of the characteristic length to the volume of the combustion chamber (1) to the cross-sectional area of ​​the throat (13); The DC mutual-impact injector (2) is equipped with a central area first injection hole (21), a central area second injection hole (22), and a side area injection hole (23). The oxidant is injected through the central area first injection hole (21), and the fuel is injected through the central area second injection hole (22) and the side area injection hole (23).

2. The thrust chamber of the wide-condition liquid rocket engine according to claim 1, characterized in that, The DC mutual-impact injector (2) uses a combination of self-igniting propellants, with nitrogen tetroxide as the oxidant and methylhydrazine as the fuel.

3. The thrust chamber of the wide-condition liquid rocket engine according to claim 1, characterized in that, The cooling jet from the side zone is ejected through the side zone injection hole (23) and then impacts the inner wall of the straight section (11) of the combustion chamber for cooling.

4. The thrust chamber of the wide-condition liquid rocket engine according to claim 3, characterized in that, The side zone cooling flow rate ratio is 20% to 30%, which is the ratio of the fuel mass flow rate flowing through the side zone injection hole (23) to the total fuel mass flow rate.

5. The thrust chamber of the wide-condition liquid rocket engine according to claim 3, characterized in that, The angle θ of the cooling jet impacting the wall in the edge region is 30° to 45°.

6. The thrust chamber of a wide-condition liquid rocket engine according to claim 1, characterized in that, The sotel average particle size (SMD) of the spray field formed by the DC mutual impact injector (2) is 150-200 μm.

7. The thrust chamber of a wide-condition liquid rocket engine according to claim 1, characterized in that, The design vacuum thrust of the thrust chamber under rated operating conditions is 100-300 N, the combustion chamber pressure is 0.8-1.0 MPa, and the mixture ratio is 1.65, where the mixture ratio is the ratio of oxidant to fuel mass flow rate.

8. The thrust chamber of the wide-condition liquid rocket engine according to claim 1, characterized in that, The first injection hole (21) in the central area, the second injection hole (22) in the central area, and the injection hole (23) in the side area are all uniformly arranged in a single ring, with their distribution positions corresponding one-to-one radially, and the number N is the same, where N is 6 or 8.

9. The thrust chamber of a wide-condition liquid rocket engine according to claim 1, characterized in that, The inner diameter L of the straight segment (11) c It is φ25~35mm.

10. The thrust chamber of a wide-condition liquid rocket engine according to claim 1, characterized in that, The combustion chamber (1) is made of niobium alloy, and a high-temperature anti-oxidation silicide coating is prepared on its inner and outer surfaces. The DC mutual impact injector (2) is made of 7715D high-temperature titanium alloy.

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