JP4885301B2 - Rocket nozzle and rocket engine combustion gas flow control method - Google Patents

Description

  The present invention relates to a rocket nozzle and a rocket engine combustion gas flow control method.

  FIG. 1 shows a bell type nozzle 100A as a rocket nozzle. Assuming that the channel cross-sectional area at the throat 101A is At and the channel cross-sectional area at the nozzle outlet 102A is A1, the expansion ratio eA of the bell type nozzle 100A is expressed by the ratio of A1 to At (A1 / At). FIG. 2 shows a bell type nozzle 100B as a rocket nozzle. Assuming that the channel cross-sectional area at the throat 101B is At and the channel cross-sectional area at the nozzle outlet 102B is A2, the expansion ratio eB of the bell type nozzle 100B is expressed by a ratio (A2 / At). Here, A2> A1, and therefore eB> eA.

  Specific thrust that represents the performance of a rocket nozzle is known. The specific thrust depends on the expansion ratio and the atmospheric pressure around the rocket nozzle. Therefore, the specific thrust changes during the ascent of the rocket.

  FIG. 3 is a graph showing changes in specific thrust with respect to altitude. The vertical axis of the graph is specific thrust, and the horizontal axis is altitude. The performance curve 121 shows the change with respect to the altitude of the specific thrust of the bell type nozzle 100A. The performance curve 122 shows the change with respect to the altitude of the specific thrust of the bell type nozzle 100B.

  The performance curve 121 and the performance curve 122 intersect at a certain altitude. At an altitude lower than the intersecting altitude, the specific thrust of the bell type nozzle 100A is larger than the specific thrust of the bell type nozzle 100B, and at an altitude higher than the intersecting altitude, the specific thrust of the bell type nozzle 100B is larger than the specific thrust of the bell type nozzle 100A.

  If the expansion ratio of the rocket nozzle can be changed while the rocket is rising, the specific thrust of the rocket nozzle can be kept high over a wide altitude range.

  FIG. 4 shows a dual bell nozzle 110 as a rocket nozzle. The dual bell nozzle 110 includes a first stage nozzle 115 as a part from the throat 111 to an inflection point (inflection point) 112 and a second stage nozzle 116 as a part from the inflection point 112 to the nozzle outlet 113. . Each of the first stage nozzle 115 and the second stage nozzle 116 has a bell shape. Here, the channel cross-sectional area at the throat 111 is represented by At, the channel cross-sectional area at the inflection point 112 is represented by A1, and the channel cross-sectional area at the nozzle outlet 113 is represented by A2.

  As shown in FIG. When the atmospheric pressure around the dual bell nozzle 110 is high, the flow formed by the combustion gas separates from the inner wall surface of the dual bell nozzle 110 at the inflection point 112. Hereinafter, such a flow is referred to as a low expansion flow. The expansion ratio of the dual bell nozzle 110 in the low expansion flow state substantially matches the expansion ratio of the bell type nozzle 100A.

  As shown in FIG. When the atmospheric pressure around the dual bell nozzle 110 is low, the flow formed by the combustion gas separates from the inner wall surface of the dual bell nozzle 110 at the nozzle outlet 113. Hereinafter, such a flow is referred to as a high expansion flow. The expansion ratio of the dual bell nozzle 110 in the state of high expansion flow substantially matches the expansion ratio of the bell type nozzle 100B.

  Referring to FIG. 3, it is ideal that the flow of the combustion gas in the dual bell nozzle 110 transitions from the low expansion flow state to the high expansion flow state at the altitude at the intersection of the performance curve 121 and the performance curve 122. . The intersection of the performance curve 121 and the performance curve 122 is referred to as an ideal transition point 120. However, as described in Non-Patent Document 1, the flow of the combustion gas in the dual bell nozzle 110 at an altitude much lower than the altitude of the ideal transition point 120 changes from a low expansion flow state to a high expansion flow state. It is known to transition. Therefore, the change of the specific thrust of the dual bell nozzle 110 with respect to the altitude is as shown by a performance curve 123.

  Since the state of the flow of combustion gas in the dual bell nozzle 110 transitions at an altitude much lower than the ideal transition point 120, the specific thrust of the dual bell nozzle 110 decreases during the transition. The decrease in the specific thrust is greater as the altitude at which the transition occurs is lower. In addition, when the altitude immediately after the transition is low, the flow of combustion gas tends to separate from the inner wall surface of the second-stage nozzle 116 before the nozzle outlet 113, and thus vibration increases.

G. Hagemann, M.M. Frey, D.D. Manski, “A Critical Assessment of Dual-Bell Nozzles”, 1997

  An object of the present invention is to provide a control method for a rocket nozzle and a rocket engine combustion gas flow in which the transition of the state of the combustion gas flow is appropriately performed.

  The means for solving the problem will be described below using the numbers used in the (DETAILED DESCRIPTION). These numbers are added to clarify the correspondence between the description of (Claims) and (Mode for Carrying Out the Invention). However, these numbers should not be used to interpret the technical scope of the invention described in (Claims).

  The rocket nozzle according to the present invention includes a dual bell nozzle (10) and gas introduction portions (40A to 40D) for introducing gas into a space (80) surrounded by the dual bell nozzle. Combustion gas flows through the space. The dual bell nozzle includes a bell-shaped first-stage nozzle (20) surrounding the upstream portion (81) of the space and a bell-shaped second-stage nozzle (30) surrounding the downstream portion (82) of the space. Is provided. The dual bell nozzle has an inflection point (12) between the first stage nozzle and the second stage nozzle. The gas introduction unit includes first gas introduction ports (45, 63) provided on the first inner wall surface of the first stage nozzle. The gas is introduced into the space from the first gas inlet.

  Turbine exhaust gas after driving the turbo pump (4, 5) is preferably used as the gas.

  It is preferable that the first gas introduction port ejects the gas in a film shape along the first inner wall surface.

  The gas introduction part includes a first manifold (42) provided around the first stage nozzle, and a first passage (43) connecting the first manifold and the first introduction port. preferable. The first passage includes a portion (44) in which a cross-sectional area of the flow path decreases as it approaches the introduction port. The turbine exhaust gas flows from the turbo pump into the manifold and flows to the first inlet through the first passage.

  The gas introduction unit includes a pipe (41) connecting the turbo pump and the first manifold, an exhaust pipe (46) branched from the pipe, and a valve (47) provided in the exhaust pipe. It is preferable.

  The valve preferably operates based on the altitude of the rocket equipped with the rocket nozzle.

  The gas introduction part preferably includes a second gas introduction port (55) provided in the second inner wall surface of the second stage nozzle and a three-way valve (51). Each of the first gas inlet, the second gas inlet, and the turbo pump is connected to the three-way valve. The second gas introduction port ejects the turbine exhaust gas in a film shape along the second inner wall surface.

  The three-way valve is preferably connected to the turbo pump by switching between the first gas inlet and the second gas inlet based on the altitude of the rocket equipped with the rocket nozzle.

  Air is preferably used as the gas.

  The gas inlet includes an air inlet (62) provided on an outer wall surface of the first stage nozzle, an air passage (61) connecting the air inlet and the first gas inlet, and the air It is preferable to provide a slide body (64) for opening and closing the intake port.

  The rocket engine combustion gas flow control method according to the present invention includes the step of introducing gas into a space (80) surrounded by a dual bell nozzle (10). Combustion gas flows through the space. The dual bell nozzle includes a bell-shaped first-stage nozzle (20) surrounding the upstream portion (81) of the space and a bell-shaped second-stage nozzle (30) surrounding the downstream portion (82) of the space. Is provided. The dual bell nozzle has an inflection point (12) between the first stage nozzle and the second stage nozzle. In the step of introducing the gas, the gas is introduced into the space from a first gas inlet (45, 63) provided on a first inner wall surface of the first stage nozzle.

  In the step of introducing the gas, it is preferable to control the introduction amount of the gas based on the altitude of the rocket equipped with the dual bell nozzle.

  Turbine exhaust gas after driving the turbo pump (4, 5) is preferably used as the gas. An exhaust pipe (46) branches from a pipe (41) that guides the turbine exhaust gas from the turbo pump to the first gas inlet. A valve (47) is provided in the exhaust pipe. In the step of introducing the gas, when it is detected that the altitude of the rocket has exceeded the first altitude, it is preferable to open the valve that has been closed.

  Preferably, the rocket is lifted off with the valve open. In the step of introducing the gas, when it is detected that the altitude of the rocket has exceeded a second altitude lower than the first altitude, it is preferable to close the valve that has been opened.

  The turbine exhaust gas after driving the turbo pump is preferably used as the gas. A second gas inlet (55) is provided on the second inner wall surface of the second stage nozzle. The second gas introduction port ejects the turbine exhaust gas in a film shape along the second inner wall surface. Each of the first gas inlet and the second gas inlet is connected to the turbo pump via a three-way valve (51). In the step of introducing the gas, when it is detected that the altitude of the rocket has exceeded the first altitude, the three-way valve causes the turbine exhaust gas to flow from a first state in which the turbine exhaust gas flows to the first gas inlet. It is preferable to switch to the second state for flowing to the second gas inlet.

The rocket is preferably lifted off when the three-way valve is in the second state.
In the step of introducing the gas, when it is detected that the altitude of the rocket has exceeded a second altitude lower than the first altitude, the three-way valve is preferably switched from the second state to the first state. .

  Air is preferably used as the gas. The first gas inlet is connected by an air passage (61) to an air intake (62) provided on the outer wall surface of the first stage nozzle. In the step of introducing the gas, when it is detected that the altitude of the rocket exceeds the first altitude, the slide body (64) moves from the first position where the air intake is not closed to the second position where the air intake is closed. ) Is preferably slid.

  It is preferable that the rocket lifts off in a state where the slide body is in the second position. In the step of introducing the gas, when it is detected that the altitude of the rocket exceeds the second altitude which is lower than the first altitude, the slide body is slid from the second position to the first position. preferable.

  ADVANTAGE OF THE INVENTION According to this invention, the control method of the rocket nozzle and the rocket engine combustion gas flow by which the transition of the state of the flow of combustion gas is performed appropriately is provided.

FIG. 1 shows a bell type nozzle with a low expansion ratio. FIG. 2 shows a bell type nozzle with a high expansion ratio. FIG. 3 is a graph comparing the relationship between specific thrust and altitude between nozzles. FIG. 4 shows a dual bell nozzle. FIG. 5 shows a state where the combustion gas flow is a low expansion flow in the dual bell nozzle. FIG. 6 shows a state in which the combustion gas flow is a highly expanded flow in the dual bell nozzle. FIG. 7 is a schematic view of a rocket engine equipped with a dual bell nozzle according to the first embodiment of the present invention. FIG. 8 is an enlarged view near the inflection point of the dual bell nozzle according to the first embodiment. FIG. 9 is a graph for explaining the effect of the dual bell nozzle according to the first embodiment. FIG. 10 is a schematic view of a dual bell nozzle according to the second embodiment of the present invention. FIG. 11 is a schematic view of a dual bell nozzle according to the third embodiment of the present invention. FIG. 12 is an enlarged view near the inflection point of the dual bell nozzle according to the third embodiment. FIG. 13 is a schematic view of a dual bell nozzle according to the fourth embodiment of the present invention.

  With reference to the accompanying drawings, a mode for carrying out a rocket nozzle and rocket engine combustion gas flow control method according to the present invention will be described below.

(First embodiment)
FIG. 7 shows a rocket engine 1 according to the first embodiment of the present invention. The rocket engine 1 includes a fuel tank 2, an oxidant tank 3, a fuel turbo pump 4, an oxidant turbo pump 5, and a rocket nozzle 8. The fuel turbo pump 4 includes a pump 4a and a turbine 4b that rotates in the same body as the pump 4a. The oxidant turbo pump 5 includes a pump 5a and a turbine 5b that rotates in the same body as the pump 5a. The rocket nozzle 8 includes an injector 6, a combustion chamber 7, a dual bell nozzle 10, and a gas introduction part 40A. The combustion chamber 7 is provided with a cooling passage 7 a for cooling the wall surface of the combustion chamber 7. The dual bell nozzle 10 includes a first stage nozzle 20 as a part from the throat 11 to an inflection point (inflection point) 12 and a second stage nozzle 30 as a part from the inflection point 12 to the nozzle outlet 13. . The inflection point 12 is disposed between the first stage nozzle 20 and the second stage nozzle 30. Each of the first stage nozzle 20 and the second stage nozzle 30 has a bell shape. The dual bell nozzle 10 surrounds the space 80. The first stage nozzle 20 surrounds the upstream portion 81 of the space 80. The second stage nozzle 30 surrounds the downstream portion 82 of the space 80. The first stage nozzle 20 may be called a base nozzle. The second stage nozzle 30 may be called an expansion nozzle.

  Manifolds 21 to 23 are provided along the periphery of the first stage nozzle 20. Each of the manifolds 21 to 23 has a donut shape. A cooling passage 20 a for cooling the wall surface of the first stage nozzle 20 is provided in a portion between the manifold 21 and the throat 11 of the first stage nozzle 20. The manifold 22 is arranged on the inflection point 12 side of the manifold 21 so as to be adjacent to the manifold 21. The manifold 23 is disposed between the manifold 22 and the inflection point 12. A cooling passage 20 b is provided in the portion between the manifold 22 and the manifold 23 of the first stage nozzle 20 to cool the wall surface of the first stage nozzle 20.

  The gas introduction part 40 </ b> A includes a pipe 41 and a manifold 42 provided along the periphery of the first stage nozzle 20. The manifold 42 has a donut shape. The manifold 42 is disposed on the inflection point 12 side of the manifold 23 so as to be adjacent to the manifold 23. The pipe 41 connects the oxidant turbo pump 5 and the manifold 42.

  The fuel tank 2 stores fuel such as liquid hydrogen. The oxidant tank 3 stores an oxidant such as liquid oxygen. The fuel turbo pump 4 supplies fuel from the fuel tank 2 to the injector 6. The oxidant turbo pump 5 supplies the oxidant from the oxidant tank 3 to the injector 6. The injector 6 mixes fuel and oxidant and injects them into the combustion chamber 7. The injected fuel and oxidant are ignited to produce combustion gases. Combustion gas flows from the combustion chamber 7 through the throat 11 into the space 80. The combustion gas flows through the upstream portion 81 and the downstream portion 82 in order, and flows out from the nozzle outlet 13.

  The fuel turbo pump 4 also supplies fuel from the fuel tank 2 to the manifold 21. The fuel passes from the manifold 21 through the cooling passage 20a and the cooling passage 7a in order, and then enters the manifold 22. The fuel enters the manifold 23 from the manifold 22 through the cooling passage 20 b and is sent from the manifold 23 to the fuel turbo pump 4. The fuel drives the fuel turbo pump 4 and the oxidant turbo pump 5. Turbine exhaust gas as fuel after driving the fuel turbo pump 4 and the oxidant turbo pump 5 flows into the manifold 42 through the pipe 41. The gas introduction unit 40A introduces turbine exhaust gas into the space 80 in order to control the combustion gas flow.

  Referring to FIG. 8, the gas introduction unit 40 </ b> A includes an introduction port 45 provided on the inner wall surface of the first stage nozzle 20, and an introduction passage 43 that connects the manifold 42 and the introduction port 45. The introduction port 45 is disposed upstream of the inflection point 12. The introduction passage 43 includes a tapered portion 44 in which the flow path cross-sectional area decreases as the introduction passage 45 is approached. Turbine exhaust gas flowing into the manifold 42 is ejected from the introduction port 45 into the space 80 through the introduction passage 43. Turbine exhaust is accelerated as it passes through the taper 44. The introduction port 45 ejects the turbine exhaust gas in a film shape in the direction of the inflection point 12 along the inner wall surface of the first stage nozzle 20.

  The turbine exhaust gas ejected in a film form from the inlet 45 prevents the combustion gas flowing in the space 80 from being in a state of high expansion flow at a timing (low altitude) earlier than the optimum point.

  An effect of cooling the wall surface of the second stage nozzle 30 by the turbine exhaust gas introduced into the space 80 is also expected.

  The effect by this embodiment is demonstrated with reference to FIG. FIG. 9 is a graph showing a change with time in performance (specific thrust) after lift-off. The vertical axis of the graph is performance (specific thrust), and the horizontal axis is the time from lift-off. A performance curve 131 shows a change with time of the performance (specific thrust) of the rocket nozzle 8 when the combustion gas flow is in a low expansion flow state where the combustion gas flow is separated at the inflection point 12. The performance curve 132 shows the change with time of the performance (specific thrust) of the rocket nozzle 8 when the combustion gas flow is in a high expansion flow state where the combustion gas flow is separated at the nozzle outlet 13. It is ideal that the flow of the combustion gas in the rocket nozzle 8 transitions from the low expansion flow state to the high expansion flow state at the timing of the intersection of the performance curve 131 and the performance curve 132. The intersection of the performance curve 131 and the performance curve 132 is referred to as an ideal transition point 130. The performance curve 133 shows the simulation result of the performance (specific thrust) of the rocket nozzle 8 when the turbine exhaust gas is not introduced into the space 80 from the introduction port 45. The performance curve 134 shows the simulation result of the performance (specific thrust) of the rocket nozzle 8 when the turbine exhaust gas is introduced into the space 80 from the introduction port 45.

  By introducing the turbine exhaust gas into the space 80 from the introduction port 45, the transition timing from the low expansion flow state to the high expansion flow state can be delayed, and the transition timing is controlled near the ideal transition point 130. It was done. When the timing of the transition is controlled near the ideal transition point 130, the amount of decrease in the specific thrust during the transition is reduced, and the vibration immediately after the transition is suppressed.

(Second Embodiment)
The rocket engine 1 according to the second embodiment of the present invention is configured in the same manner as the rocket engine 1 according to the first embodiment, except that the gas introduction part 40A is replaced with a gas introduction part 40B. Yes. Referring to FIG. 10, in addition to the pipe 41, the manifold 42, the introduction passage 43, and the introduction port 45, the gas introduction unit 40B includes an exhaust pipe 46 branched from the pipe 41 and a valve 47 provided in the exhaust pipe 46. Is provided. The valve 47 opens and closes based on the control signal 71. When the valve 47 is open, a part of the turbine exhaust gas from the oxidant turbo pump 5 is introduced into the space 80 from the introduction port 45, and the rest is exhausted to the atmosphere through the exhaust pipe 46. When the valve 47 is closed, the entire amount of turbine exhaust gas from the oxidant turbo pump 5 is introduced into the space 80 from the inlet 45.

  The altitude H of the rocket equipped with the rocket engine 1 is monitored. The rocket lifts off with the valve 47 open. Thereafter, the valve 47 is kept open until it is detected that the altitude H exceeds the predetermined altitude H1. The altitude H1 is determined based on the altitude at which the performance curve 133 deviates from the performance curve 131, for example. At the time of lift-off and immediately after the lift-off, since the atmospheric pressure around the rocket is sufficiently high, the flow rate of the turbine exhaust gas introduced into the space 80 from the inlet 45 is at least, and the flow of the combustion gas is kept in a low expansion flow state. It is.

  When it is detected that the altitude H exceeds the altitude H1, the valve 47 is closed. Thereafter, the valve 47 is kept closed until it is detected that the altitude H exceeds the predetermined altitude H2. The altitude H2 is determined based on the ideal transition point 130 described above, for example. The altitude H2 is higher than the altitude H1. While the rocket rises from altitude H1 to altitude H2, the flow of the turbine exhaust gas introduced into the space 80 from the inlet 45 is controlled to a large flow rate. Therefore, the flow of the combustion gas is kept in a low expansion flow state.

  When it is detected that the altitude H exceeds the altitude H2, the valve 47 is opened. Thereafter, the valve 47 is kept open. By opening the valve 47 and reducing the flow rate of the turbine exhaust gas introduced into the space 80 from the introduction port 45, the transition of the combustion gas flow from the low expansion flow state to the high expansion flow state is promoted.

  Although it is possible to keep the valve 47 closed from the lift-off until it is detected that the altitude H exceeds the altitude H1, the turbine exhaust gas introduced into the space 80 from the inlet 45 by opening the valve 47 It is expected that the flow of combustion gas immediately after ignition is stabilized by reducing the flow rate of.

(Third embodiment)
The rocket engine 1 according to the third embodiment of the present invention is configured in the same manner as the rocket engine 1 according to the first embodiment except that the gas introduction part 40A is replaced with a gas introduction part 40C. Yes. Referring to FIG. 11, in addition to the pipe 41, the manifold 42, the introduction passage 43, and the introduction port 45 described above, the gas introduction unit 40C includes a three-way valve 51 provided in the pipe 41 and the inflection point 12 side of the manifold 42. Includes a manifold 52 disposed adjacent to the manifold 42. The three-way valve 51 is switched between the manifold 42 and the manifold 52 based on the control signal 72 and connected to the oxidant turbo pump 5. Turbine exhaust gas from the oxidant turbo pump 5 flows into the manifold 42 or 52 according to the switching state of the three-way valve 51.

  Referring to FIG. 12, the manifold 52 is provided along the periphery (or the inflection point 12) of the first stage nozzle 20. The manifold 52 has a donut shape. The gas introduction part 40 </ b> C includes an introduction port 55 provided on the inner wall surface of the second stage nozzle 30, and an introduction passage 53 that connects the manifold 52 and the introduction port 55. The introduction passage 53 includes a tapered portion 54 in which the flow path cross-sectional area decreases as the introduction passage 55 is approached. The turbine exhaust gas that has flowed into the manifold 52 is ejected from the introduction port 55 into the space 80 through the introduction passage 53. The turbine exhaust gas is accelerated as it passes through the taper 54. The introduction port 55 ejects the turbine exhaust gas in a film shape in the direction of the nozzle outlet 13 along the inner wall surface of the second stage nozzle 30.

  The rocket's altitude H is monitored. The rocket lifts off with the three-way valve 51 connecting the manifold 52 to the oxidant turbo pump 5. When the manifold 52 is connected to the oxidant turbo pump 5, the introduction port 55 is connected to the oxidant turbo pump 5. The three-way valve 51 continues to connect the manifold 52 to the oxidant turbo pump 5 until it is detected that the altitude H has exceeded the altitude H1 described above. At the time of lift-off and immediately after the lift-off, the atmospheric pressure around the rocket is sufficiently high, so even if the turbine exhaust gas is not introduced into the space 80 from the inlet 45 on the upstream side of the inflection point 12, the flow of the combustion gas is A low expansion flow state is maintained.

  When it is detected that the altitude H exceeds the altitude H1, the three-way valve 51 switches from a state in which the manifold 52 is connected to the oxidant turbo pump 5 to a state in which the manifold 42 is connected to the oxidant turbo pump 5. Thereafter, until it is detected that the altitude H exceeds the above-described altitude H2, the three-way valve 51 keeps the state in which 42 is connected to the oxidant turbo pump 5. When the manifold 42 is connected to the oxidant turbo pump 5, the introduction port 45 is connected to the oxidant turbo pump 5. While the rocket rises from altitude H 1 to altitude H 2, the turbine exhaust gas flows from the oxidant turbo pump 5 through the manifold 42 to the inlet 45 and is introduced into the space 80. Therefore, the flow of the combustion gas is kept in a low expansion flow state.

  When it is detected that the altitude H exceeds the altitude H2, the three-way valve 51 switches from a state in which the manifold 42 is connected to the oxidant turbo pump 5 to a state in which the manifold 52 is connected to the oxidant turbo pump 5. Thereafter, the three-way valve 51 maintains a state in which the manifold 52 is connected to the oxidant turbo pump 5. By stopping the introduction of the turbine exhaust gas from the inlet 45 to the space 80 by the three-way valve 51, the transition of the combustion gas flow from the low expansion flow state to the high expansion flow state is promoted. While the rocket rises above the altitude H 2, the turbine exhaust gas flows from the oxidant turbo pump 5 through the manifold 52 to the inlet 55 and is introduced into the space 80. Therefore, the inner wall surface of the second stage nozzle 30 is film-cooled.

  It is possible to keep the three-way valve 51 connected to the manifold 42 and the oxidant turbo pump 5 from the lift-off until it is detected that the altitude H exceeds the altitude H1. By keeping the agent turbo pump 5 connected, it is expected that the combustion gas flow immediately after ignition is stabilized.

(Fourth embodiment)
In the rocket engine 1 according to the fourth embodiment of the present invention, the gas introduction unit 40A is replaced with the gas introduction unit 40D and the fuel before driving the fuel turbo pump 4 and the oxidant turbo pump 5 is used. Except that the wall surface of the two-stage nozzle 30 is cooled, the configuration is the same as that of the rocket engine 1 according to the first embodiment.

  Referring to FIG. 13, in the rocket engine 1 according to the present embodiment, a manifold 24 is provided along the periphery of the first stage nozzle 20. The manifold 24 is disposed between the manifold 23 and the inflection point 12. The manifold 23 and the manifold 24 are connected by a pipe 48 provided outside the dual bell nozzle 10. A manifold 31 is provided along the nozzle outlet 13. Each of the manifolds 24 and 31 has a donut shape. The second stage nozzle 30 is provided with a cooling passage 30 a for cooling the wall surface of the second stage nozzle 30. The fuel that has entered the manifold 23 through the cooling passage 20 b enters the manifold 24 through the pipe 48. Then, it enters the manifold 31 through the cooling passage 30 a and is sent from the manifold 31 to the turbine 4 b of the fuel turbo pump 4. The fuel is exhausted after driving the fuel turbo pump 4 and the oxidant turbo pump 5.

  The gas introduction unit 40D includes an air intake 62 provided on the outer wall surface of the first stage nozzle 20, an introduction port 63 provided on the inner wall surface of the first stage nozzle 20, an air intake 62, and the introduction port 63. Are provided with an air passage 61, a slide body 64, and a drive mechanism 65. The slide body 64 is driven by the drive mechanism 65 and slides so as to open and close the air intake port 62. The introduction port 63 is disposed upstream of the inflection point 12.

  The rocket's altitude H is monitored. The rocket lifts off in a state where the slide body 64 is disposed at a position where the air intake 62 is closed. Thereafter, the slide body 64 is held at a position where the air intake port 62 is closed until it is detected that the altitude H exceeds the altitude H1. Since the atmospheric pressure around the rocket is sufficiently high at the time of lift-off and immediately after the lift-off, the flow of the combustion gas is maintained in a low expansion flow state even if air is not introduced into the space 80 from the introduction port 63.

  When it is detected that the altitude H exceeds the altitude H1, the drive mechanism 65 slides the slide body 64 to a position where the air intake 62 is not closed. Thereafter, the slide body 64 is held at a position where the air intake port 62 is not closed until it is detected that the altitude H exceeds the altitude H2. While the rocket rises from the altitude H1 to the altitude H2, air is introduced into the space 80 from the inlet 63. Therefore, the flow of the combustion gas is kept in a low expansion flow state.

  When it is detected that the altitude H exceeds the altitude H2, the drive mechanism 65 slides the slide body 64 to a position where the air intake port 62 is closed. By stopping the introduction of air from the inlet 63 into the space 80, the transition of the combustion gas flow from the low expansion flow state to the high expansion flow state is promoted.

  Thereafter, the slide body 64 is held at a position where the air intake 62 is closed. The combustion gas is prevented from leaking outside through the air passage 61.

  It is possible to hold the slide body 64 in a position where the air intake 62 is not closed during the period from the lift-off until it is detected that the altitude H exceeds the altitude H1, but by closing the air intake 62, The combustion gas flow immediately after ignition is expected to be stable.

  The gas introduction part 40D can also be applied to a fixed fuel rocket. In the fourth embodiment, the wall surface of the second stage nozzle 30 may be film cooled.

  If the gas introduction parts 40A to 40D are applied to a rocket equipped with a cluster engine, it becomes easy to align the transition timings in a plurality of engines. As a result, the attitude control of the rocket equipped with the cluster engine is stabilized.

  In addition, the gas introduction parts 40A to 40D are also effective when the rocket returns.

  In the second to fourth embodiments, by controlling the flow rate of the gas introduced into the space 80 from the inlets 45 and 63 provided on the inner wall surface of the first stage nozzle 20 based on the altitude H of the rocket, It is possible to control the timing of transition more precisely.

DESCRIPTION OF SYMBOLS 1 ... Rocket engine 2 ... Fuel tank 3 ... Oxidant tank 4 ... Fuel turbo pump 4a ... Pump 4b ... Turbine 5 ... Oxidant turbo pump 5a ... Pump 5b ... Turbine 6 ... Injector 7 ... Combustion chamber 7a ... Cooling passage 8 ... rocket nozzle 10 ... dual bell nozzle 11 ... throat 12 ... inflection point 13 ... nozzle outlet 20 ... first stage nozzles 20a, 20b ... cooling passages 21-24 ... manifold 30 ... second stage nozzle 30a ... cooling passage 31 ... manifolds 40A to 40D ... gas introduction parts 41, 48 ... piping 42 ... manifold 43 ... introduction passage 44 ... tip 45 ... introduction port 46 ... exhaust pipe 47 ... valve 51 ... three-way valve 52 ... manifold 53 ... introduction passage 54 ... tip Details 55 ... Inlet 61 ... Air passage 62 ... Air intake 63 ... Inlet 64 ... Sliding body 65 ... Drive mechanisms 71, 72 ... Control signal 80 ... Space 81 ... Upstream portion 82 ... Downstream portion 100A, 100B ... Bell type nozzles 101A, 101B ... Throat 102A, 102B ... Nozzle outlet 110 ... Dual bell nozzle 111 ... Throat 112 ... Inflection point 113 ... Nozzle outlet 115 ... 1st stage nozzle 116 ... 2nd stage nozzle 120, 130 ... Ideal transition points 121-123, 131-134 ... Performance curve

Claims (18)

Dual bell nozzle,
A gas introduction part for introducing gas into a space surrounded by the dual bell nozzle,
Combustion gas flows through the space,
The dual bell nozzle is
A bell-shaped first stage nozzle surrounding the upstream portion of the space;
A bell-shaped second stage nozzle surrounding the downstream portion of the space;
With
The dual bell nozzle has an inflection point between the first stage nozzle and the second stage nozzle,
The gas introduction part includes a first gas introduction port provided on a first inner wall surface of the first stage nozzle,
The gas is introduced into the space from the first gas inlet,
The gas introduction part introduces the gas into the space to keep the flow of the combustion gas in a low expansion flow state,
In the state of the low expansion flow, the flow of the combustion gas is separated from the inner wall of the dual bell nozzle at the inflection point and flows not along the second stage nozzle. The rocket nozzle according to claim 1, wherein a turbine exhaust gas after driving a turbo pump is used as the gas. The rocket nozzle according to claim 2, wherein the first gas introduction port ejects the gas in a film shape along the first inner wall surface. The gas introduction part is
A first manifold provided around the first stage nozzle;
A first passage connecting the first manifold and the first inlet,
The first passage includes a portion in which a cross-sectional area of the flow path decreases as it approaches the introduction port,
The rocket nozzle according to claim 3, wherein the turbine exhaust gas flows into the manifold from the turbo pump and flows to the first inlet through the first passage. The gas introduction part is
A pipe connecting the turbo pump and the first manifold;
An exhaust pipe branched from the pipe;
The rocket nozzle according to claim 4, further comprising a valve provided in the exhaust pipe. The rocket nozzle according to claim 5, wherein the valve operates based on an altitude of a rocket including the rocket nozzle. The gas introduction part is
A second gas inlet provided in a second inner wall surface of the second stage nozzle;
With a three-way valve,
Each of the first gas inlet, the second gas inlet, and the turbo pump is connected to the three-way valve,
The rocket nozzle according to claim 3 or 4, wherein the second gas introduction port ejects the turbine exhaust gas in a film shape along the second inner wall surface. The rocket nozzle according to claim 7, wherein the three-way valve is connected to the turbo pump by switching between the first gas inlet and the second gas inlet based on an altitude of a rocket including the rocket nozzle. The rocket nozzle according to claim 1, wherein air is used as the gas. The gas introduction part is
An air intake provided in an outer wall surface of the first stage nozzle;
An air passage connecting the air inlet and the first gas inlet;
The rocket nozzle of Claim 9 provided with the slide body which opens and closes the said air intake. Introducing a gas into the space surrounded by the dual bell nozzle;
Combustion gas flows through the space,
The dual bell nozzle is
A bell-shaped first stage nozzle surrounding the upstream portion of the space;
A bell-shaped second stage nozzle surrounding the downstream portion of the space;
The dual bell nozzle has an inflection point between the first stage nozzle and the second stage nozzle,
In introducing the gas,
Introducing the gas into the space from a first gas inlet provided on the first inner wall surface of the first stage nozzle to keep the flow of the combustion gas in a low expansion flow state;
A control method of a rocket engine combustion gas flow in which the flow of the combustion gas is separated from the inner wall of the dual bell nozzle at the inflection point and flows along the second stage nozzle in the low expansion flow state. In introducing the gas,
The rocket engine combustion gas flow control method according to claim 11, wherein the gas introduction amount is controlled based on an altitude of a rocket having the dual bell nozzle. The turbine exhaust gas after driving the turbo pump is used as the gas,
An exhaust pipe branches from a pipe that guides the turbine exhaust gas from the turbo pump to the first gas inlet,
A valve is provided in the exhaust pipe;
In introducing the gas,
The rocket engine combustion gas flow control method according to claim 12, wherein when the altitude of the rocket exceeds the first altitude, the closed valve is opened. The rocket lifts off with the valve open,
In introducing the gas,
The rocket engine combustion gas flow control method according to claim 13, wherein when the rocket altitude is detected to exceed a second altitude lower than the first altitude, the valve that has been opened is closed. Turbine exhaust gas after driving the turbo pump is used as the gas,
A second gas inlet is provided in the second inner wall surface of the second stage nozzle;
The second gas introduction port blows out the turbine exhaust gas in a film shape along the second inner wall surface,
Each of the first gas inlet and the second gas inlet is connected to the turbo pump via a three-way valve,
In introducing the gas,
When it is detected that the altitude of the rocket exceeds the first altitude, the three-way valve causes the turbine exhaust gas to flow to the second gas inlet from the first state in which the turbine exhaust gas flows to the first gas inlet. The rocket engine combustion gas flow control method according to claim 12, wherein the rocket engine combustion gas flow is switched to a second state. The rocket lifts off when the three-way valve is in the second state,
In introducing the gas,
When detecting that the altitude of the rocket exceeds a second altitude lower than the first altitude,
The rocket engine combustion gas flow control method according to claim 15, wherein the three-way valve switches from the second state to the first state. Air is used as the gas,
The first gas introduction port is connected to an air intake port provided on an outer wall surface of the first stage nozzle by an air passage;
In introducing the gas,
The rocket engine combustion according to claim 12, wherein when detecting that the altitude of the rocket exceeds the first altitude, the slide body is slid from a first position where the air intake is not closed to a second position where the air intake is closed. Gas flow control method. The rocket lifts off with the slide body in the second position,
In introducing the gas,
The rocket engine combustion gas flow according to claim 17, wherein when detecting that the altitude of the rocket exceeds the second altitude which is lower than the first altitude, the slide body is slid from the second position to the first position. Control method.

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