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CN-120384817-B - Thrust chamber of liquid rocket engine with wide working condition

CN120384817BCN 120384817 BCN120384817 BCN 120384817BCN-120384817-B

Abstract

The invention relates to the technical field of design of thrust chambers of spacecraft liquid rocket engines, and provides a thrust chamber of a liquid rocket engine under a wide working condition, which comprises a combustion chamber and a direct current mutual impact type injector, wherein the combustion chamber comprises a straight line section, a converging section, a throat and an expanding section which are sequentially connected, the direct current mutual impact type injector is arranged at the end part of the straight line section, the length-diameter ratio of the straight line section is 1.2-1.5, the characteristic length is 290-350 mm, a first central area injection hole, a second central area injection hole and a side area injection hole are arranged on the direct current mutual impact type injector, an oxidant is injected through the first central area injection hole, and fuel is injected through the second central area injection hole and the side area injection hole. The invention greatly widens the adaptable working condition range of the thrust chamber, and solves the problems that the throat temperature of the space engine with high specific impulse performance is too high and is greatly influenced by working condition deviation.

Inventors

  • CHEN RUIDA
  • LIU CHANGGUO
  • YU PENG
  • YE YIXIANG
  • Shi Zhehang
  • XU HUI
  • CHEN HONGYU

Assignees

  • 上海空间推进研究所

Dates

Publication Date
20260512
Application Date
20250410

Claims (10)

  1. 1. The thrust chamber of the liquid rocket engine under the wide working condition is characterized by comprising a combustion chamber (1) and a direct current mutual impact injector (2); The combustion chamber (1) comprises a straight line section (11), a convergence section (12), a throat (13) and an expansion section (14) which are sequentially connected, wherein the direct current mutual impact type injector (2) is configured at the end part of the straight line section (11), the length-diameter ratio of the straight line section (11) is 1.2-1.5, the characteristic length is 290-350 mm, the length-diameter ratio is the ratio of the length L c of the straight line section (11) to the inner diameter D c , and the characteristic length is the ratio of the volume of the combustion chamber (1) to the cross-sectional area of the throat (13); The direct current mutual impact type injector (2) is provided with a first central area injection hole (21), a second central area injection hole (22) and a side area injection hole (23), the oxidant is sprayed out through the first central area injection hole (21), and the fuel is sprayed out through the second central area injection hole (22) and the side area injection hole (23).
  2. 2. The wide-duty liquid rocket engine thrust chamber of claim 1, wherein the direct current mutual-impact injector (2) adopts an autoignition propellant combination, dinitrogen tetroxide is an oxidant, and methyl hydrazine is a fuel.
  3. 3. A wide working range liquid rocket engine thrust chamber according to claim 1, wherein the side zone cooling jet is sprayed out through the side zone spraying holes (23) and then impacts the inner wall surface of the straight section (11) of the combustion chamber for cooling.
  4. 4. A wide-regime liquid rocket engine thrust chamber according to claim 3, wherein the side cooling flow ratio is 20% -30%, the side cooling flow ratio being the ratio of the mass flow of fuel flowing through the side injection holes (23) to the total mass flow of fuel.
  5. 5. A wide-regime liquid rocket engine thrust chamber according to claim 3, wherein the edge zone cooling jet impingement angle θ is 30 ° to 45 °.
  6. 6. The wide working range liquid rocket engine thrust chamber of claim 1, wherein the sotaier average particle Size (SMD) of the spray field formed by the direct current inter-impingement injector (2) is 150-200 μm.
  7. 7. The wide-duty liquid rocket engine thrust chamber of claim 1, wherein the designed vacuum thrust under the rated working conditions of the thrust chamber is 100-300 n, the combustion chamber pressure is 0.8-1.0 mpa, the mixing ratio is 1.65, and the mixing ratio is the ratio of the oxidant to the fuel mass flow.
  8. 8. The wide working condition liquid rocket engine thrust chamber according to claim 1, wherein the first central region injection holes (21), the second central region injection holes (22) and the side region injection holes (23) are uniformly arranged in a single circle, the distribution positions are in radial one-to-one correspondence, and the number N is the same, and N is 6 or 8.
  9. 9. The wide working condition liquid rocket engine thrust chamber according to claim 1, wherein the inner diameter L c of the straight line section (11) is phi 25-35 mm.
  10. 10. The thrust chamber of the wide working condition liquid rocket engine according to claim 1, wherein the base material of the combustion chamber (1) is niobium alloy, the inner surface and the outer surface are provided with high-temperature oxidation-resistant silicide coatings, and the base material of the direct-current mutual impact injector (2) is 7715D high-temperature titanium alloy.

Description

Thrust chamber of liquid rocket engine with wide working condition Technical Field The invention relates to the technical field of design of thrust chambers of spacecraft liquid rocket engines, in particular to a wide-working-condition liquid rocket engine thrust chamber. Background When the traditional space double-component attitude and orbit control liquid rocket engine thrust chamber for the spacecraft is designed, in order to pursue high specific impulse performance, the high-temperature area of the combustion chamber is positioned in the throat area with the maximum heat flux density, the outer wall temperature is even higher than 1400 ℃, and the ablation area at the end of life is also positioned in the throat, but the influence brought by the method is that: (1) The throat bears the highest heat flux density, the highest temperature and the highest gas scouring speed, and ablation failure is easy to occur after long-term work; (2) The side region cooling liquid film covers the front end of the straight line section of the combustion chamber, is far away from the throat, is difficult to effectively extend to the throat region, is greatly influenced by working condition deviation, and particularly has obviously raised throat temperature in the high mixing ratio working condition; (3) The residence time of the fuel gas in the combustion chamber is insufficient, so that the combustion efficiency is obviously fluctuated by the change of working conditions. In the prior art, the problem of overhigh throat temperature is relieved by increasing the side region cooling flow ratio or adopting jet cooling of various angles in the side region, such as a spatial staggered distribution liquid film cooling structure and a flow calculation method thereof disclosed in patent document CN116608058A, but the problem can cause specific impulse performance loss or prolong the iteration period of design parameters, and increase the test run verification cost. Disclosure of Invention Aiming at the defects in the prior art, the invention aims to provide a wide-working-condition liquid rocket engine thrust chamber. The invention provides a wide-working-condition liquid rocket engine thrust chamber, which comprises a combustion chamber and a direct-current mutually-striking injector; The combustion chamber comprises a straight line segment, a converging segment, a throat and an expanding segment which are sequentially connected, the direct current mutual impact type injector is arranged at the end part of the straight line segment, the length-diameter ratio of the straight line segment is 1.2-1.5, preferably 1.3, the characteristic length is 290-350 mm, preferably 320mm, the length-diameter ratio is the ratio of the length L c of the straight line segment to the inner diameter D c, and the characteristic length is the ratio of the volume of the combustion chamber to the cross-sectional area of the throat; The direct current mutual impact type injector is provided with a first central area injection hole, a second central area injection hole and a side area injection hole, wherein the oxidant is sprayed out through the first central area injection hole, and the fuel is sprayed out through the second central area injection hole and the side area injection hole. Preferably, the direct current mutual impact injector adopts an autoignition propellant combination, dinitrogen tetroxide is used as an oxidant, and methyl hydrazine is used as fuel. Preferably, the side zone cooling jet flow is sprayed out through the side zone spraying holes and then impacts the inner wall surface of the straight line section of the combustion chamber for cooling. Preferably, the side cooling flow rate ratio is 20% -30%, and the side cooling flow rate ratio is the ratio of the mass flow rate of fuel flowing through the side injection hole to the total mass flow rate of fuel. Preferably, the edge area cooling jet impingement angle θ is 30 ° to 45 °, for example 30 °, and still for example 40 °. Preferably, the average particle size of the Sotel of the spray field formed by the direct current mutual impact type injector is 150-200 μm, preferably 180 μm. Preferably, the designed vacuum thrust under the rated working condition of the thrust chamber is 100-300N, the pressure of the combustion chamber is 0.8-1.0 MPa, and the mixing ratio is 1.65, and the mixing ratio is the ratio of the mass flow of the oxidant to the mass flow of the fuel. Preferably, the first central region injection holes, the second central region injection holes and the side region injection holes are uniformly arranged in a single circle, the distribution positions are in radial one-to-one correspondence, the number N is the same, and the number N is 6 or 8. Preferably, the inner diameter L c of the straight line segment is phi 25-35 mm, and the inner diameter L c of the straight line segment is 30mm. Preferably, the base material of the combustion chamber is niobium alloy, high-tempe