Combined engine system, aerospace vehicle, and flight control method and apparatus

Abstract

The application discloses a combined engine system, an aerospace vehicle and a flight control method and device, and relates to the field of aerospace vehicles.The combined engine system comprises a detonation ramjet engine device, a detonation rocket engine device and a turbojet engine device, wherein the detonation rocket engine device is arranged inside the detonation ramjet engine device, and the turbojet engine device is arranged outside the detonation ramjet engine device; the detonation rocket engine device and the turbojet engine device are arranged to be capable of working independently or working in combination to form a combined mode; and the working modes of the combined engine system include a ramjet mode, a rocket mode, a turbojet mode and a combined mode.The combined engine system is suitable for more application scenarios and can reduce the design of an oxidant cabin.

Description

Technical Field

[0001] The present invention relates to the field of aerospace engines, and more particularly, to a combined detonation engine system and a flight control method. Background Art

[0002] There are two types of combustion in nature: slow combustion and detonation combustion. The flame propagation rate of slow combustion is relatively low. The combustion mode in power units such as internal combustion engines, aircraft engines and gas turbines is slow combustion. The characteristic of detonation combustion is that the upstream of the combustion zone is a shock wave structure, and the shock wave and the combustion zone are coupled together for propagation. The flame propagation speed of detonation combustion is much higher than that of slow combustion, usually reaching several thousand meters per second.

[0003] Competition in the aerospace sector is becoming increasingly fierce, and research into key innovative technologies in this area is drawing increasing attention from various countries. In recent years, with the continued advancement of research into hypersonic vehicles and single-stage orbital propulsion systems, novel continuous rotating detonation engine technology has seen rapid development. Research has demonstrated that propulsion technology based on detonation combustion can significantly reduce fuel consumption and significantly improve the specific impulse characteristics of power plants. This holds significant significance for expanding the operating envelope of air-breathing vehicles and enhancing the economic efficiency and combat performance of existing weaponry.

[0004] The continuous rotating detonation engine is a power technology that utilizes detonation combustion. Its characteristics and advantages are: (1) Only one successful detonation is required for the detonation wave to propagate continuously along the circumference of the combustion chamber; (2) the combustion rate is fast, the heat release intensity is large, and the combustion chamber structure is compact, which can shorten the engine length; (3) it has a supercharging characteristic, which can reduce the number of compressor stages of a turbine engine or reduce the total pressure loss in the inlet of a ramjet engine, which is conducive to simplifying the propulsion system design and improving the engine thrust-to-weight ratio; (4) it can operate in an air-breathing mode or a rocket mode, and its operating range can vary from subsonic to supersonic speeds with high Mach numbers.

[0005] At present, although the research on continuous rotating detonation engines has achieved many results and accumulated relatively much experience, the shortcomings of single-mode detonation engines are also obvious. In order to study the feasibility under different application scenarios, research on detonation combination engines and detonation engines and traditional engine combined power has been carried out.

[0006] Patent CN114439646A discloses an air-turbine rocket-ramjet combined propulsion system; Patent CN114810417A discloses a full-rotating detonation mode rocket-ramjet combined engine and operating method; Patent CN113513430A discloses a dual- or tri-component continuous rotating detonation engine; Patent CN111140399A discloses an additively manufactured continuous rotating detonation rocket engine and its additive manufacturing method. The combined continuous rotating detonation engine or single-mode engine of the existing technology has the following disadvantages: (1) The current detonation ramjet engine has a relatively narrow flight speed range. Only when it has a certain speed (flight speed reaches 2.5Ma) can it achieve self-supercharging and achieve better engine starting and working conditions. There is no adjustment device for the air intake ducts at different altitudes to adapt to the air intake requirements at different altitudes; (2) The rocket mode detonation engine can start at zero speed, but it needs to carry sufficient oxidizer and fuel throughout the entire flight range. The required oxidizer and fuel are heavy and occupy a large space; (3) The combined detonation ramjet and detonation rocket engine has poor cruise economy when flying below 2.5Ma (the economic speed range of traditional engines), and the oxidizer and fuel consumption is large. The oxidizer required by the detonation rocket engine cannot be replenished in flight mode; (4) There is no adjustable tail nozzle for adaptation, so the aircraft cannot be adjusted along the flight trajectory. Summary of the Invention

[0007] In view of the technical problems of limitations in the combined rotating detonation engine system and single-mode detonation engine of the prior art, the embodiments of the present application provide a combined engine system, an aerospace vehicle, and a flight control method and device. The embodiments of the present application aim to combine a new detonation engine with a traditional turbojet engine. The four engine modes of rocket mode, ramjet mode, turbojet mode, and combined mode are combined to work together, can be switched between and operate independently, and combine traditional power with new power technology. The combined engine system can adapt to more application scenarios. The turbojet mode technology is mature and the turbojet mode can supplement oxidizer for the rocket mode. The unique oxidizer supplementation design greatly reduces the design of the engine oxidation chamber, allowing the aircraft engine to adopt a smaller size while achieving the same flight range. In addition, the turbojet mode provides supplementary oxidizer to the rocket mode, which can cool the combustion chamber air film and improve the life of the electromechanical components of the detonation ramjet engine device and the detonation rocket engine device.

[0008] The technical solution of this application is as follows:

[0009] A combined engine system comprising: a detonation ramjet engine device, a detonation rocket engine device and a turbojet engine device;

[0010] The detonation rocket engine device is arranged inside the detonation ramjet engine device, and the turbojet engine device is arranged outside the detonation ramjet engine device;

[0011] The detonation rocket engine device and the turbojet engine device are configured to be able to work independently or in combination to form a combined mode. The detonation ramjet engine device works independently, and the working modes of the combined engine system include ramjet mode, rocket mode, turbojet mode and combined mode.

[0012] In some exemplary embodiments, the detonation ramjet engine device includes an intake module, a combustion chamber module, and a tail jet module connected in sequence; the intake module is provided with an intake flow passage communicating with the outside; the combustion chamber module is provided with an annular detonation ramjet combustion chamber and a mounting cavity located inside the detonation ramjet combustion chamber; the tail jet module is provided with a tail jet flow passage communicating with the outside, the intake flow passage, the detonation ramjet combustion chamber, and the tail jet flow passage are connected in sequence, and the tail jet flow passage is also connected with the mounting cavity;

[0013] The detonation rocket engine device is arranged in the installation cavity, and the turbojet engine device is arranged outside the tail nozzle module.

[0014] In some exemplary embodiments, the air intake module includes an air intake housing and an air intake guide, the air intake housing is sleeved on the outside of the air intake guide, and the air intake flow channel is located between the air intake housing and the air intake guide;

[0015] The combustion chamber module includes a combustion chamber shell, the combustion chamber shell is connected to the air intake shell, the installation cavity is connected to the air intake guide member, and the detonation ramjet combustion chamber is located between the combustion chamber shell and the installation cavity;

[0016] The tail nozzle module includes a tail nozzle shell and a tail nozzle pipe assembly. The tail nozzle shell is connected to the combustion chamber shell. The tail nozzle pipe assembly is located inside the tail nozzle shell. The internal space of the tail nozzle pipe assembly forms the tail nozzle flow channel.

[0017] In some exemplary embodiments, the air intake module further includes: an air intake regulating assembly, which is arranged between the air intake housing and the air intake guide, and the air intake housing, the air intake regulating assembly, and the air intake guide together enclose the air intake flow channel, and the air intake regulating assembly is configured to adjust the flow area of ​​the air intake flow channel.

[0018] In some exemplary embodiments, the air intake adjustment assembly includes:

[0019] an adjusting member, slidably connected to the air intake housing;

[0020] an adjusting guide rail, cooperating with the adjusting member and extending along the axial direction of the detonation ramjet engine device; and

[0021] The first driving member is connected to the adjusting member and is configured to drive the adjusting member to move along the adjusting guide rail to change the width of the gap between the adjusting member and the air intake guide member to adjust the flow area of ​​the air intake channel.

[0022] In some exemplary embodiments, the adjustment member includes a flow channel adjustment portion for enclosing the intake flow channel, and a support connection portion connected to the adjustment guide rail and the first driving member;

[0023] The first driving member and the adjusting guide rail are located between the outer side wall of the flow channel adjusting portion and the inner side wall of the air intake housing.

[0024] In some exemplary embodiments, the air intake guide is configured as a hollow structure, and a fuel storage cavity is provided in the air intake guide;

[0025] The detonation ramjet engine device also includes a fuel injection module, which is provided with a feed port connected to the fuel storage chamber and an injection port connected to the detonation ramjet combustion chamber. The fuel injection module is configured to inject fuel in the fuel storage chamber into the detonation ramjet combustion chamber.

[0026] In some exemplary embodiments, the tail jet module further includes a tail jet adjustment assembly connected to the tail jet pipe assembly and configured to adjust a flow area of ​​the tail jet flow channel.

[0027] In some exemplary embodiments, the tail nozzle assembly includes a plurality of guide vanes, which are arranged along the circumference of the tail jet flow channel and enclose the tail jet flow channel; along the exhaust direction, the tail jet flow channel includes a contraction section and an expansion section that are connected in sequence, and the connection between the contraction section and the expansion section forms a throat portion;

[0028] The tail jet adjustment assembly includes a plurality of second driving members provided on the outside of the tail jet pipe assembly, the plurality of second driving members being connected to the plurality of guide vanes in a one-to-one correspondence to drive the plurality of guide vanes located at the throat portion to contract inwardly or expand outwardly along the radial direction of the tail jet flow channel, so that the flow area of ​​the throat portion is adjusted;

[0029] The tail jet adjustment assembly further includes a plurality of fixing members, and the plurality of fixing members are connected to the plurality of second driving members in a one-to-one correspondence.

[0030] In some example embodiments, the combined engine system further comprises an oxidizer supplement channel connecting the turbojet engine device and the detonation rocket engine device, the oxidizer supplement channel being configured to supplement oxidizer from the turbojet engine device to the detonation rocket engine device.

[0031] In some example embodiments, the turbojet engine device comprises a compressor and a combustion chamber; the detonation rocket engine device comprises a supply module; an outlet of the compressor is in communication with the combustion chamber and an input of the oxidizer supplement channel; an output of the oxidizer supplement channel is connected to the supply module.

[0032] In some example embodiments, the oxidizer supplement channel comprises a first channel, a second channel and a third channel in sequence;

[0033] the first channel is at least partially embedded in the combustion chamber shell, the second channel extends inwardly from the combustion chamber shell and into the mounting cavity, and the third channel is located in the mounting cavity.

[0034] In some example embodiments, the turbojet engine device comprises a plurality of turbojet engines, the plurality of turbojet engines being arranged in a circumferential direction of the tail nozzle shell;

[0035] The number of the oxidizer supplement channels is a plurality, and the plurality of oxidizer supplement channels are connected to the plurality of turbojet engines one by one.

[0036] In some example embodiments, the detonation rocket engine device comprises a plurality of detonation rocket engines, the plurality of detonation rocket engines being arranged in a circumferential direction of the mounting cavity;

[0037] In the exhaust direction, the detonation rocket engine is arranged outwardly inclined relative to the axial direction of the detonation ramjet engine device.

[0038] An aerospace vehicle comprising the combined engine system according to any one of the example embodiments described above.

[0039] A flight control method for performing a flight task by using the aerospace vehicle according to the example embodiments described above, comprising:

[0040] Based on satisfying a first flight condition, controlling the combined engine system to work in a turbojet mode, a rocket mode or a combined mode; wherein the detonation rocket engine device and the turbojet engine device of the combined mode both provide thrust.

[0041] Based on the second flight condition being satisfied, controlling the combined engine system to operate in a combined mode, wherein the detonation rocket engine assembly provides thrust and the turbojet engine assembly supplies oxidizer to the detonation rocket engine assembly;

[0042] Based on the satisfaction of the third flight condition, controlling the combined engine system to operate in a ramjet mode;

[0043] Based on satisfying the fourth flight condition, the combined engine system is controlled to operate in rocket mode.

[0044] In some exemplary embodiments, the first flight condition includes: an aviation altitude less than or equal to 25 kilometers, or a flight speed between 0 Ma and 1.6 Ma;

[0045] The second flight condition includes: a flight speed between 1.6 Ma and 2.5 Ma;

[0046] The third flight condition includes: the aviation altitude is less than 100 kilometers and the flight speed is greater than or equal to 2.5 Ma;

[0047] The fourth flight condition includes: the aviation altitude is greater than 100 kilometers.

[0048] In some exemplary embodiments, the flight control method further includes: starting the turbojet engine device and / or the detonation rocket engine device in response to a zero-speed start instruction.

[0049] A control device includes a processor and a memory storing a computer program, wherein when the processor executes the computer program, the steps of the flight control method as described in any one of the above exemplary embodiments are implemented.

[0050] The combined engine system provided by the embodiment of the present application has the following beneficial effects:

[0051] In the combined engine system provided in the embodiment of the present application, the detonation rocket engine device and the turbojet engine device are configured to be able to work independently or in combination to form a combined mode. The detonation ramjet engine device works independently, and the working modes of the combined engine system include ramjet mode, rocket mode, turbojet mode and combined mode. The combined engine system provided in the embodiment of the present application adopts a combined mode that combines the rocket mode and the turbojet mode. The turbojet mode technology is mature, and the turbojet mode can supplement the oxidizer for the rocket mode. The unique oxidizer supplementation design greatly reduces the design of the engine oxidation compartment, and the aircraft engine adopts a smaller size to achieve the same flight mileage. In addition, the turbojet mode provides supplementary oxidizer for the rocket mode, which can cool the combustion chamber air film and increase the life of the electromechanical components of the detonation ramjet engine device and the detonation rocket engine device.

[0052] Moreover, the four engine modes work together, can be switched between each other and operate independently, and combine traditional power with new power technology. The combined engine system can adapt to more application scenarios. For example, after the aircraft enters orbit, if the power unit is recovered, the turbojet mode is adopted. By changing the rotation speed, the thrust can be accurately adjusted to allow the power unit to land safely, and power can be provided for the recovery of the power unit. However, it is difficult to control the combined thrust using the detonation mode alone, and it cannot gradually change to zero thrust. The combined power mode enables the combined engine system of the present application to start at zero speed (select turbojet mode, detonation rocket mode or a combination of the two modes as needed), and can fly in the economic speed range of the turbojet engine and the economic speed range of the detonation ramjet engine, thereby improving fuel economy and reducing fuel emissions. When the turbojet module is working, it can provide part of the oxidizer for the detonation rocket engine, so the oxidizer required by the detonation rocket during the entire flight is greatly reduced.

[0053] Furthermore, the structural arrangement in which the detonation rocket engine device is arranged inside the detonation ramjet engine device enables the combined engine to broaden the speed range, have a smaller design size, and improve the thrust-to-weight ratio of the aircraft.

[0054] Other features and advantages of the present invention will be set forth in the description that follows. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] The accompanying drawings are used to provide a further understanding of the technical solution of the present invention and constitute a part of the specification. Together with the embodiments of the present application, they are used to explain the technical solution of the present invention and do not constitute a limitation on the technical solution of the present invention.

[0056] Figure 1 A schematic cross-sectional view of the combined engine system according to an embodiment of the present application;

[0057] Figure 2 for Figure 1 A schematic diagram of the three-dimensional structure of the tail nozzle adjustment assembly and the tail nozzle assembly;

[0058] Figure 3 for Figure 1 A three-dimensional structural example diagram of a combined engine system;

[0059] Figure 4 for Figure 1 A right-side structural diagram of a combined engine system;

[0060] Reference numerals:

[0061] 1- detonation ramjet engine device, 11- air intake module, 111- inlet flow channel, 112- air intake housing, 113- air intake guide, 114- air intake adjustment assembly, 1141- adjustment member, 1141-1- flow channel adjustment part, 1141-2- support connection part, 1142- adjustment guide rail, 1143- first drive member, 12- combustion chamber module, 121- detonation ramjet combustion chamber, 122- installation cavity, 123- combustion chamber housing, 13- tail nozzle module, 131- tail nozzle flow channel, 1311- contraction section, 1312- expansion section, 1313- throat part, 132-Tail nozzle shell, 133-Tail nozzle assembly, 1331-Guide vane, 1332-Guide outer vane, 1333-Guide inner vane, 134-Tail nozzle adjustment assembly, 1341-Second drive member, 1342-Fixer, 14-Fuel storage chamber, 15-Fuel injection module, 2-Detonation rocket engine device, 21-Supply module, 22-Detonation rocket engine, 3-Turbojet engine device, 31-Turbojet engine, 4-Oxidant supplement channel, 41-Input end, 42-Output end, 43-First channel, 44-Second channel, 45-Third channel DETAILED DESCRIPTION

[0062] To make the objectives, technical solutions, and advantages of the present invention more clearly understood, embodiments of the present invention will be described in detail below with reference to the accompanying drawings. It is apparent that the described embodiments are only a portion of the embodiments of the present invention, not all of them. All other embodiments derived by persons of ordinary skill in the art based on the embodiments of the present invention without inventive effort are intended to fall within the scope of protection of the present invention.

[0063] See also Figure 1-4 The structural diagram of the combined engine system shown in the embodiment of the present application is shown in FIG. Figure 1 As shown, an embodiment of the present application provides a combined engine system, including a detonation ramjet engine device 1, a detonation rocket engine device 2, and a turbojet engine device 3;

[0064] The detonation rocket engine device 2 is arranged inside the detonation ramjet engine device 1, and the turbojet engine device 3 is arranged outside the detonation ramjet engine device 1;

[0065] The detonation rocket engine device 2 and the turbojet engine device 3 are configured to work independently or in combination to form a combined mode. The detonation ramjet engine device 1 works independently, and the working modes of the combined engine system include ramjet mode, rocket mode, turbojet mode and combined mode.

[0066] In the combined engine system provided in the embodiment of the present application, the detonation rocket engine device 2 and the turbojet engine device 3 are configured to be able to work independently or in combination to form a combined mode. The detonation ramjet engine device 1 works independently. The working modes of the combined engine system in the embodiment of the present application include ramjet mode, rocket mode, turbojet mode and combined mode. The combined engine system provided in the embodiment of the present application adopts a combined mode that combines the rocket mode and the turbojet mode. The turbojet mode technology is mature. The turbojet mode can supplement the oxidizer for the rocket mode. The unique oxidizer supplementation design greatly reduces the design of the engine oxidation compartment. The aircraft engine adopts a smaller size to achieve the same flight mileage. In addition, the turbojet mode provides supplementary oxidizer for the rocket mode, which can cool the combustion chamber air film and increase the life of the electromechanical components of the detonation ramjet engine device 1 and the detonation rocket engine device 2.

[0067] Moreover, the four engine modes work together, can be switched between each other and operate independently, and combine traditional power with new power technology. The combined engine system can adapt to more application scenarios. For example, after the aircraft enters orbit, if the power unit is recovered, the turbojet mode is adopted. By changing the rotation speed, the thrust can be accurately adjusted to allow the power unit to land safely, and power can be provided for the recovery of the power unit. However, it is difficult to control the combined thrust using the detonation mode alone, and it cannot gradually change to zero thrust. The combined power mode enables the combined engine system of the present application to start at zero speed (select turbojet mode, detonation rocket mode or a combination of the two modes as needed), and can fly in the economic speed range of the turbojet engine and the economic speed range of the detonation ramjet, increase fuel economy and reduce emissions. When the turbojet module is working, it can provide part of the oxidizer for the detonation rocket, so the oxidizer required for the entire flight detonation rocket is greatly reduced.

[0068] Furthermore, the structural arrangement in which the detonation rocket engine device 2 is arranged inside the detonation ramjet engine device 1 enables the combined engine to broaden the speed range, have a smaller design size, and improve the thrust-to-weight ratio of the aircraft.

[0069] In some exemplary embodiments, Figure 1 As shown, the detonation ramjet engine device 1 includes an air intake module 11, a combustion chamber module 12, and a tail jet module 13 connected in sequence; the air intake module 11 is provided with an air intake flow channel 111 communicating with the outside; the combustion chamber module 12 is provided with an annular detonation ramjet combustion chamber 121 and a mounting cavity 122 located inside the detonation ramjet combustion chamber 121; the tail jet module 13 is provided with a tail jet flow channel 131 communicating with the outside; the air intake flow channel 111, the detonation ramjet combustion chamber 121, and the tail jet flow channel 131 are connected in sequence, and the tail jet flow channel 131 is also connected with the mounting cavity 122;

[0070] The detonation rocket engine device 2 is arranged in the installation cavity 122 , and the turbojet engine device 3 is arranged outside the tail nozzle module 13 .

[0071] In some exemplary embodiments, the air intake module 11 includes an air intake housing 112 and an air intake guide 113 , wherein the air intake housing 112 is sleeved on the outside of the air intake guide 113 , and the air intake flow channel 111 is located between the air intake housing 112 and the air intake guide 113 ;

[0072] The combustion chamber module 12 includes a combustion chamber housing 123 , which is connected to the air intake housing 112 , a mounting cavity 122 connected to the air intake guide 113 , and a detonation ramjet combustion chamber 121 located between the combustion chamber housing 123 and the mounting cavity 122 ;

[0073] The tail nozzle module 13 includes a tail nozzle shell 132 and a tail nozzle assembly 133 . The tail nozzle shell 132 is connected to the combustion chamber shell 123 . The tail nozzle assembly 133 is located inside the tail nozzle shell 132 . The internal space of the tail nozzle assembly 133 forms a tail nozzle flow channel 131 .

[0074] In some exemplary embodiments, the air intake module 11 further includes:

[0075] The air intake regulating assembly 114 is disposed between the air intake housing 112 and the air intake guide 113 . The air intake housing 112 , the air intake regulating assembly 114 , and the air intake guide 113 enclose the air intake passage 111 . The air intake regulating assembly 114 is configured to regulate the flow area of ​​the air intake passage 111 .

[0076] In some exemplary embodiments, Figure 1 As shown, the air intake adjustment assembly 114 includes an adjustment member 1141, an adjustment guide rail 1142 and a first driving member 1143;

[0077] Among them, the adjusting member 1141 is slidably connected to the air intake housing 112; the adjusting guide rail 1142 cooperates with the adjusting member 1141 and extends along the axial direction of the detonation ramjet engine device 1; the first driving member 1143 is connected to the adjusting member 1141, and is configured to drive the adjusting member 1141 to move along the adjusting guide rail 1142, so that the gap width between the adjusting member 1141 and the air intake guide member 113 changes, so as to adjust the flow area of ​​the intake flow channel 111.

[0078] In the embodiment of the present application, the adjustment member 1141 is an annular member, and four first drive members 1143 are provided. The four first drive members 1143 are evenly distributed along the circumference of the adjustment member 1141 and can operate synchronously, pushing and pulling the adjustment member 1141 along the axial direction of the combined engine system, thereby changing the flow area of ​​the intake air passage 111.

[0079] The knock ramjet engine device 1 of the embodiment of the present application can realize opening, closing and adjustment of the air inlet flow channel 111 by simple slider shaft movement without changing the air inlet cone and the outer shell of the engine, and is suitable for flight height.

[0080] In some exemplary embodiments, as shown in Figure 1 The adjusting member 1141 includes a flow channel adjusting part 1141-1 for enclosing the air inlet flow channel 111, and a support connecting part 1141-2 connected with the adjusting guide rail 1142 and the first driving member 1143.

[0081] The first driving member 1143 and the adjusting guide rail 1142 are located between the outer side wall of the flow channel adjusting part 1141-1 and the inner side wall of the air inlet outer shell 112.

[0082] In some exemplary embodiments, as shown in Figure 1 The air inlet guide member 113 is provided in a hollow structure, and the fuel storage cavity 14 is arranged in the air inlet guide member 113.

[0083] The knock ramjet engine device 1 further comprises a fuel injection module 15 provided with a fuel inlet port in communication with the fuel storage cavity 14 and an injection port in communication with the knock ramjet combustion chamber 1, and the fuel injection module 15 is arranged to inject the fuel in the fuel storage cavity 14 into the knock ramjet combustion chamber 1.

[0084] In some exemplary embodiments, as shown in Figure 1-2 The tail jet module 13 further comprises:

[0085] The tail jet adjusting assembly 134 is connected with the tail jet pipe assembly 133 and is arranged to adjust the flow area of the tail jet flow channel 131.

[0086] In some exemplary embodiments, as shown in Figure 1-2 The tail jet pipe assembly 133 comprises a plurality of guide vanes 1331 arranged along the circumference of the tail jet flow channel 131 and enclosing the tail jet flow channel 131; along the exhaust direction, the tail jet flow channel 131 comprises a converging section 1311 and a diverging section 1312 connected in sequence, and the connecting part of the converging section 1311 and the diverging section 1312 forms a throat part 1313.

[0087] The tail jet adjusting assembly 134 comprises a plurality of second driving members 1341 arranged outside the tail jet pipe assembly 133, and the plurality of second driving members 1341 are connected with the plurality of guide vanes 1331 one by one to drive the plurality of guide vanes 1331 to contract radially inward or expand radially outward at the position of the throat part 1313 along the tail jet flow channel 131, so that the flow area of the throat part 1313 is adjusted.

[0088] The tail jet adjustment assembly 134 further includes a plurality of fixing members 1342 , and the plurality of fixing members 1342 are connected to the plurality of second driving members 1341 in a one-to-one correspondence.

[0089] The flow area of ​​the tail jet flow channel 131 of the detonation ramjet engine device 1 of the embodiment of the present application can be adjusted. By radially raising and lowering the guide vanes 1331 of the tail nozzle assembly 133, the diameter of the throat portion 1313 of the tail jet flow channel 131 can be changed, thereby adjusting the expansion ratio.

[0090] like Figure 2 As shown, in some exemplary embodiments, the guide vanes 1331 include outer guide vanes 1332 and inner guide vanes 1333 , and the outer guide vanes 1332 and the inner guide vanes 1333 overlap to form the tail jet flow channel 131 .

[0091] Optionally, the second driving member 1341 is configured as a cylinder that electromagnetically drives the telescopic rod. The fixing member 1342 is used to fix and support the second driving member 1341 and the guide vane 1331.

[0092] In some exemplary embodiments, Figure 1 As shown, the combined engine system of the present application further includes:

[0093] The oxidizer replenishing channel 4 connects the turbojet engine device 3 and the detonation rocket engine device 2 . The oxidizer replenishing channel 4 is configured to utilize the turbojet engine device 3 to replenish the oxidizer to the detonation rocket engine device 2 .

[0094] Specifically, the turbojet engine device 3 has mature technology and a unique oxidizer replenishment design, which greatly reduces the design of the oxidation chamber of the aerospace engine, allowing the aircraft to achieve the same flight mileage with a smaller size.

[0095] In addition, the oxidizer replenishment channel 4 in the turbojet mode can provide a cooling solution for the detonation mode and the ramjet mode, thereby improving the reliability of components of the spacecraft combination engine.

[0096] The combined engine system described in the embodiment of the present application is provided with an oxidizer replenishment channel 4, and the rocket mode, turbojet mode and combined mode can be switched in combination, and then combined with the ramjet mode. The combined engine system of the present application combines traditional power with new power technology and can adapt to more application scenarios.

[0097] For example, after the spacecraft enters orbit, if the power unit is to be recovered, the turbojet mode is used. By changing the rotational speed, the thrust can be precisely adjusted to allow the power unit to land safely, and power can be provided for the recovery of the power unit. However, it is difficult to control the combined thrust using the detonation mode alone, and it cannot gradually change to zero thrust.

[0098] The combined power mode enables the combined engine system of the present application to start at zero speed (select turbojet mode, detonation rocket mode or a combination of the two modes as needed), and can fly in the economic speed range of the turbojet engine and the economic speed range of the detonation ramjet, increasing fuel economy and reducing emissions. When the turbojet module is working, it can provide part of the oxidizer for the detonation rocket, so the oxidizer required for the entire flight of the detonation rocket is greatly reduced.

[0099] In some exemplary embodiments, Figure 1 As shown, the turbojet engine device 3 includes a compressor and a combustion chamber (not shown in the figure); the detonation rocket engine device 2 includes a supply module 21; the air outlet of the compressor is connected to the combustion chamber and the input end 41 of the oxidizer replenishing channel 4; the output end 42 of the oxidizer replenishing channel 4 is connected to the supply module 21.

[0100] In some exemplary embodiments, Figure 1 As shown, the oxidant supplement channel 4 includes a first channel 43, a second channel 44 and a third channel 45 which are connected in sequence;

[0101] The first channel 43 is at least partially embedded in the combustion chamber housing 123 , the second channel 44 extends inward from the combustion chamber housing 123 and into the mounting cavity 122 , and the third channel 45 is located in the mounting cavity 122 .

[0102] In some exemplary embodiments, Figure 3-4 As shown, the turbojet engine device 3 includes a plurality of turbojet engines 31, and the plurality of turbojet engines 31 are arranged at intervals along the circumference of the tail jet casing 132;

[0103] There are multiple oxidant supply channels 4 , and the multiple oxidant supply channels 4 are connected to the multiple turbojet engines 3 in a one-to-one correspondence.

[0104] In an exemplary embodiment of the combined engine system, Figure 1 、 3 As shown, the number of the first channels 43 of the oxidant supplement channel 4 is set to 6, and the number of the turbojet engines 31 is set to 6. The 6 first channels 43 are evenly distributed axially along the combustion chamber casing 123, and the 6 first channels 43 are set in one-to-one correspondence with the 6 turbojet engines 31.

[0105] It is worth noting that the turbojet engine assembly 3 further includes an explosive bolt module (not shown in the figure), which is used to separate the multiple turbojet engines 31 from the outer surface of the tail jet casing 132. The turbojet engine assembly 3 also includes a fixed casing, and the multiple turbojet engines 31 are fixed by their respective fixed casings.

[0106] In some exemplary embodiments, Figure 1As shown, the detonation rocket engine device 2 includes a plurality of detonation rocket engines 22, and the plurality of detonation rocket engines 22 are arranged circumferentially along the installation cavity 122;

[0107] Along the exhaust direction, the detonation rocket engine 22 is arranged to be tilted outward relative to the axial direction of the detonation ramjet engine device 1.

[0108] Specifically, the detonation rocket engine device 2 is composed of a plurality of detonation rocket engines 22, which can realize accurate and efficient control of thrust, and achieve attitude adjustment of the aircraft through thrust control in different directions.

[0109] Specifically, the detonation rocket engine 22 is provided with a detonation rocket nozzle, and high-temperature and high-pressure gas is ejected from the detonation rocket nozzle to generate thrust, and the detonation rocket nozzle opens toward the tail nozzle module 13 .

[0110] In an exemplary embodiment, the detonation rocket engines 22 are set to four, and the four detonation rocket engines 22 can work independently, or can be set to arbitrarily select several of the detonation rocket engines 22 to work in combination.

[0111] An embodiment of the present application provides an aerospace vehicle, comprising a combined engine system as described in any one of the above exemplary embodiments.

[0112] The aerospace vehicle provided in the embodiments of the present application includes a combined engine system as described in any one of the above exemplary embodiments, and thus has the technical features and advantages of the combined engine system described in the above exemplary embodiments, which will not be described in detail here.

[0113] The present application provides a flight control method for performing a flight mission using an aerospace vehicle as described in the above exemplary embodiments, specifically comprising:

[0114] Based on the first flight condition being met, controlling the combined engine system to operate in a turbojet mode, a rocket mode, or a combined mode; wherein both the detonation rocket engine device 2 and the turbojet engine device 3 in the combined mode provide thrust;

[0115] Based on the second flight condition being met, the combined engine system is controlled to operate in a combined mode, wherein the detonation rocket engine device 2 provides thrust, and the turbojet engine device 3 supplements the detonation rocket engine device 2 with oxidizer;

[0116] Based on the satisfaction of the third flight condition, controlling the combined engine system to operate in the ramjet mode;

[0117] Based on meeting the fourth flight condition, the combined engine system is controlled to operate in rocket mode.

[0118] In an exemplary embodiment, the first flight condition includes: an aerial altitude less than or equal to 25 kilometers, or a flight speed between 0 Ma and 1.6 Ma;

[0119] The second flight condition includes: the flight speed is between 1.6Ma and 2.5Ma;

[0120] The third flight condition includes: the aviation altitude is less than 100 kilometers and the flight speed is greater than or equal to 2.5Ma;

[0121] The fourth flight condition includes: the aviation altitude is greater than 100 kilometers.

[0122] In some exemplary embodiments, the flight control method further includes:

[0123] In response to the zero-speed start command, the turbojet engine device 3 and / or the detonation rocket engine device 2 are started.

[0124] The present application provides a control device, including a processor and a memory storing a computer program. When the processor executes the computer program, the steps of the flight control method as described in any of the above exemplary embodiments are implemented.

[0125] Unless otherwise specified or limited, the terms "mounted," "connected," "connect," "fixed," and the like should be interpreted broadly. For example, they may refer to fixed or detachable connections, or integration; mechanical or electrical connections; direct or indirect connections through an intermediate medium; and internal communication between two components or interaction between two components, unless otherwise specified. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on specific circumstances.

[0126] In the description of this specification, the reference terms "one embodiment", "some embodiments", "example", "specific example", or "some examples" mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in this specification and features of different embodiments or examples without contradiction.

[0127] Although the embodiments of the present invention have been shown and described above, it will be understood that the above embodiments are illustrative and are not to be construed as limitations on the present invention. A person skilled in the art may change, modify, replace and modify the above embodiments within the scope of the present invention.

Claims

1. A combined engine system, characterized in that: include: Detonation ramjet engine units, detonation rocket engine units and turbojet engine units; The detonation rocket engine device is arranged inside the detonation ramjet engine device, and the turbojet engine device is arranged outside the detonation ramjet engine device; The detonation rocket engine device and the turbojet engine device are configured to be able to operate independently or in combination to form a combined mode. The detonation ramjet engine device operates independently, and the operating modes of the combined engine system include ramjet mode, rocket mode, turbojet mode and combined mode. The combined engine system further comprises an oxidizer replenishment passage connecting the turbojet engine device and the detonation rocket engine device, wherein the oxidizer replenishment passage is configured to replenish oxidizer to the detonation rocket engine device using the turbojet engine device; The turbojet engine device includes a compressor and a combustion chamber; the detonation rocket engine device includes a supply module; the air outlet of the compressor is connected to the combustion chamber and the input end of the oxidant replenishment channel; the output end of the oxidant replenishment channel is connected to the supply module.

2. The combined engine system according to claim 1, characterized in that: The detonation ramjet engine device includes an air intake module, a combustion chamber module, and a tail jet module connected in sequence; the air intake module is provided with an air intake flow passage communicating with the outside; the combustion chamber module is provided with an annular detonation ramjet combustion chamber and a mounting cavity located inside the detonation ramjet combustion chamber; the tail jet module is provided with a tail jet flow passage communicating with the outside, the air intake flow passage, the detonation ramjet combustion chamber, and the tail jet flow passage are connected in sequence, and the tail jet flow passage is also connected with the mounting cavity; The detonation rocket engine device is arranged in the installation cavity, and the turbojet engine device is arranged outside the tail nozzle module.

3. The combined engine system according to claim 2, characterized in that: The air intake module includes an air intake housing and an air intake guide, wherein the air intake housing is sleeved on the outside of the air intake guide, and the air intake flow channel is located between the air intake housing and the air intake guide; The combustion chamber module includes a combustion chamber shell, the combustion chamber shell is connected to the air intake shell, the installation cavity is connected to the air intake guide member, and the detonation ramjet combustion chamber is located between the combustion chamber shell and the installation cavity; The tail nozzle module includes a tail nozzle shell and a tail nozzle pipe assembly. The tail nozzle shell is connected to the combustion chamber shell. The tail nozzle pipe assembly is located inside the tail nozzle shell. The internal space of the tail nozzle pipe assembly forms the tail nozzle flow channel.

4. The combined engine system according to claim 3, characterized in that: The air intake module further includes: The air intake regulating assembly is arranged between the air intake housing and the air intake guide. The air intake housing, the air intake regulating assembly and the air intake guide together form the air intake flow channel. The air intake regulating assembly is configured to regulate the flow area of ​​the air intake flow channel.

5. The combined engine system according to claim 4, characterized in that: The air intake adjustment component includes: an adjusting member, slidably connected to the air intake housing; an adjusting guide rail, cooperating with the adjusting member and extending along the axial direction of the detonation ramjet engine device; and The first driving member is connected to the adjusting member and is configured to drive the adjusting member to move along the adjusting guide rail to change the width of the gap between the adjusting member and the air intake guide member to adjust the flow area of ​​the air intake channel.

6. The combined engine system according to claim 5, characterized in that: The adjusting member includes a flow channel adjusting portion for enclosing the intake flow channel, and a supporting connection portion connected to the adjusting guide rail and the first driving member; The first driving member and the adjusting guide rail are located between the outer side wall of the flow channel adjusting portion and the inner side wall of the air intake housing.

7. The combined engine system according to any one of claims 3 to 6, characterized in that: The air intake guide is configured as a hollow structure, and a fuel storage cavity is provided in the air intake guide; The detonation ramjet engine device also includes a fuel injection module, which is provided with a feed port connected to the fuel storage chamber and an injection port connected to the detonation ramjet combustion chamber. The fuel injection module is configured to inject fuel in the fuel storage chamber into the detonation ramjet combustion chamber.

8. The combined engine system according to any one of claims 3 to 6, characterized in that: The tail nozzle module also includes: The tail jet regulating assembly is connected to the tail jet pipe assembly and is configured to regulate the flow area of ​​the tail jet flow channel.

9. The combined engine system according to claim 8, characterized in that: The tail nozzle assembly includes a plurality of guide vanes, which are arranged along the circumference of the tail jet flow channel and enclose the tail jet flow channel; along the exhaust direction, the tail jet flow channel includes a contraction section and an expansion section that are connected in sequence, and the connection between the contraction section and the expansion section forms a throat portion; The tail jet adjustment assembly includes a plurality of second driving members provided on the outside of the tail jet pipe assembly, the plurality of second driving members being connected to the plurality of guide vanes in a one-to-one correspondence to drive the plurality of guide vanes located at the throat portion to contract inwardly or expand outwardly along the radial direction of the tail jet flow channel, so that the flow area of ​​the throat portion is adjusted; The tail jet adjustment assembly further includes a plurality of fixing members, and the plurality of fixing members are connected to the plurality of second driving members in a one-to-one correspondence.

10. The combined engine system according to claim 3, characterized in that: The oxidant supplement channel includes a first channel, a second channel and a third channel that are connected in sequence; The first channel is at least partially embedded in the combustion chamber shell, the second channel extends inward from the combustion chamber shell and into the installation cavity, and the third channel is located in the installation cavity.

11. The combined engine system according to claim 3, characterized in that: The turbojet engine device comprises a plurality of turbojet engines, and the plurality of turbojet engines are arranged at intervals along the circumference of the tail jet casing; There are multiple oxidant replenishing channels, and the multiple oxidant replenishing channels are connected to the multiple turbojet engines in a one-to-one correspondence.

12. The combined engine system according to claim 2 or 9, characterized in that: The detonation rocket engine device includes a plurality of detonation rocket engines, and the plurality of detonation rocket engines are arranged along the circumference of the installation cavity; Along the exhaust direction, the detonation rocket engine is arranged to be tilted outward relative to the axial direction of the detonation ramjet engine device.

13. An aerospace vehicle, characterized in that: The invention comprises a combined engine system according to any one of claims 1 to 12.

14. A flight control method, characterized in that: Performing a flight mission using the aerospace vehicle of claim 13, comprising: Based on the first flight condition being satisfied, controlling the combined engine system to operate in a turbojet mode, a rocket mode, or a combined mode; wherein both the detonation rocket engine device and the turbojet engine device in the combined mode provide thrust; Based on the second flight condition being satisfied, controlling the combined engine system to operate in a combined mode, wherein the detonation rocket engine assembly provides thrust and the turbojet engine assembly supplies oxidizer to the detonation rocket engine assembly; Based on the satisfaction of the third flight condition, controlling the combined engine system to operate in a ramjet mode; Based on satisfying the fourth flight condition, the combined engine system is controlled to operate in rocket mode.

15. The flight control method according to claim 14, characterized in that: The first flight condition includes: an aviation altitude less than or equal to 25 kilometers, or a flight speed between 0 Ma and 1.6 Ma; The second flight condition includes: a flight speed between 1.6 Ma and 2.5 Ma; The third flight condition includes: the aviation altitude is less than 100 kilometers and the flight speed is greater than or equal to 2.5 Ma; The fourth flight condition includes: the aviation altitude is greater than 100 kilometers.

16. The flight control method according to claim 14 or 15, characterized in that: Also includes: In response to a zero-speed start command, the turbojet engine device and / or the detonation rocket engine device is started.

17. A control device, characterized in that: The system comprises a processor and a memory storing a computer program, wherein when the processor executes the computer program, the steps of the flight control method according to any one of claims 14 to 16 are implemented.

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