A power system for aerospace vehicle

Abstract

A spacecraft power system relates to the field of aerospace engine technology. The YRCW aerospace engine comprises an air intake system, a central RCW engine, and two side YRC engines. The YRCW aerospace engine can operate in four modes: ramjet turbine, ramjet, superramjet, and rocket. The ramjet combustor can accommodate four combustion modes. An inlet airflow separator randomly distributes the RCW and YRC airflows. The combustion gas impacts the turbine, driving the axial-flow intake compressor and centrifugal compressor to inhale and compress air. The compressed air is distributed through four bypass ducts, entering the cold air, hot exhaust gas, and vectored airflow hoods, respectively. This provides compressed cold air and cooling gas, as well as vectored injection air, for closed-loop combustion of the RCW and YRC engines. Adjusting the geometric positions of the conical ramjet intake valve, the compressed cold air, hot exhaust gas, the fuel injection plate, and the oil pressure and oil volume support channel valves ensures smooth transitions between the engine's multi-mode operating modes, stable flames, and sustained combustion, enabling the aerospace vehicle to achieve single-stage orbital flight within a full envelope.

Description

Technical Field

[0001] The present invention discloses an aerospace vehicle power system; specifically, it refers to a four-mode cycle engine propulsion system for a single-stage orbital power-fuselage integrated aircraft, belonging to the technical field of power propulsion systems for aviation and aerospace transport vehicles. Technical Background

[0002] In the context of the application of aerospace engineering technology in the world, how to expand the development space of human civilization depends on how inventors and designers conduct innovative research based on existing aerospace propulsion technology, integrate the technical advantages of aircraft propulsion systems and rocket propulsion systems, and create aerospace engines that can shuttle through the Earth's atmosphere, reach orbital speed and Earth escape speed, and adapt to low-Earth orbit, lunar and galaxy flights.

[0003] Since the 1930s, with the invention and application of aircraft engines and rocket aerospace engines and the improvement of engine technology performance, human scientists' dream of flying into space has been activated.

[0004] From the successful flight of the Wright brothers' first powered, heavier-than-air, controlled, manned aircraft in Kitty Hawk, North Carolina, on December 17, 1903, to the introduction of the 18-cylinder radial engine that powered the Wright Apache in 1928, and through the trials of World War II, aircraft engine technology has advanced significantly. With the advancement and continued development of aircraft technology, the dream of space exploration has been rekindled among human scientists. From the 1930s to the present, the technological approaches to space exploration have been divided into two major factions: the airplane approach and the rocket approach.

[0005] Aircraft-style: Single-stage orbital spaceplanes feature an integrated power-fuselage design, allowing for horizontal takeoff and landing and autonomous flight into space. The propulsion system for single-stage orbital spaceplanes combines an aircraft engine with a space rocket engine. Austrian engineer Eugen Sanger proposed the "Silverbird" concept in 1930. His advanced design principles were proven scientifically correct through decades of subsequent scientific experiments, serving as a model for subsequent inventors and designers. For example, from Silverbird to the US X-15, X-30, and SpaceShipTwo 1 and 2.

[0006] Rocketry: Space planes that use dual-stage orbital or multi-stage piggyback launches. In the early 20th century, with the continued advancement and development of aircraft engine technology, Rocketry advocates sought a different approach: using rocket-launched space planes to explore space and develop space tourism. Technologically advanced space planes using dual-stage orbital or multi-stage piggyback launches include the Soviet Buran and the US X-37B.

[0007] At present, the engine technology bottleneck and the high-temperature heat protection bottleneck of the materials of the single-stage and double-stage orbital space plane propulsion system are difficult to break through. The double-stage orbital or backpack-type multi-stage launch space planes "Buran" and "X-37B" have large launch mass, extremely small payload and high cost; the multi-stage launch procedure is complicated, the multi-stage separation technology is unstable, and the re-entry space plane faces high-temperature thermal barriers during unpowered gliding return, which fails to meet the actual application requirements of aerospace engineering technology.

[0008] The History of Space Exploration: From 1961 to 1962, astronauts from the Soviet Union and the United States successively entered space. By 1972, the Soviet and American Soyuz missions, respectively, had successfully launched manned lunar exploration, satellites, spacecraft (the space shuttle), and perfected docking technology for the International Space Station. Launch vehicle space exploration opened a window into human exploration of space civilization and has become a keystone of human history, science, and civilization. By 2020, with the successful relaunch of the United States' Constellation lunar exploration program, rocket technology had reached its peak.

[0009] The limitations of rocket launches for space exploration include single-use, minimal payloads (such as the Soyuz, Apollo, and Orion missions), high launch costs, and non-reusability. To reduce costs, Elon Musk in the United States developed a reusable rocket engine to restart lunar exploration. However, the Falcon 9 orbital and lunar-landing stage rocket weighed 9,500 tons, with a payload capacity of only six passengers, two astronauts, and 100 kilograms of lunar rocks. Therefore, space exploration and the development and utilization of space resources have been relegated to single-stage, orbital-to-spaceplanes with integrated power and fuselage. First, rocket launches are expensive due to their single-use nature; second, their enormous launch mass and minimal payload make them unsuitable as a transportation tool for space resource development.

[0010] Advantages of space planes: They are reusable, fully utilize atmospheric air, and their large propellant, passenger, and cargo compartments can serve as space transportation. After orbital refueling, the aircraft can reach the moon. However, the greatest technological challenge in aerospace engineering applications lies in the space engine.

[0011] Aerospace engines: Single-stage orbital propulsion systems with integrated power and fuselage, multi-mode cycle propulsion systems. Aerospace engines are the ultimate challenge in aerospace engineering applications. Currently, aviation and aerospace engines are separate in the global engine technology landscape, with no integrated aerospace engine.

[0012] 1. Aerospace engines: rocket engines, pulse combustion engines, ramjet engines, scramjet engines, and turbine engines.

[0013] 2. Comparison of various engines: The five types of engines each have their strengths and weaknesses, and sometimes using them together becomes a burden:

[0014] (1) Pulse combustion engines, ramjets, and scramjets are lightweight, simple in structure, and low in cost;

[0015] (2) Liquid-fuel rocket engine technology is mature and suitable for space flight;

[0016] (3) The acceleration performance of turbojet engines at Mach 0 to 3 is the basis for their widespread application, but the cost is high (the engine cost of an ordinary fighter jet is over 1 million US dollars). Turbine engines are complex in structure and heavy in weight. For example, the turbojet used in fighter jets, China's WS-7: 8 turbine blades, weight 1.19 tons, length 4.6 meters. The US F119: 11 turbine blades, weight 1.36 tons, length 4.8 meters. In terms of thermal efficiency, even a jet engine with 17 turbine blades has only 35%, while ramjet and scramjet engines can reach more than 90% at Mach 3 to 16.

[0017] (4) Ramjet and scramjet engines have high fuel specific impulse and thermal efficiency, simple structure, light weight and high speed, but they do not have static start capability and need to be supplemented by other engines. In 1934, the patent for ramjet aircraft obtained by René Reduc of France was a combination of booster rocket and ramjet engine. Recently, the unmanned ultra-high-speed long-range (40,000 km) bomber developed by the United States is powered by a combination of ramjet engine and turbojet engine. It uses turbojet engine for takeoff and ramjet engine when the speed is greater than 0.5 Mach (at this time, the turbojet engine becomes a burden). It can reach a speed of more than 10 Mach at an altitude of 30 kilometers and can rush out of the atmosphere at a speed of 16 Mach at a height of 60 kilometers. After shutting down the engine, it can glide for hundreds of kilometers. After entering the atmosphere, the ramjet engine will re-ignite and then rise again. The water-skiing flight can reach any place in the world within 2 hours. However, it is difficult to break through the bottleneck of engine technology and the bottleneck of high temperature resistance of materials.

[0018] (5) Pulse, ramjet, scramjet, and turbine engines are only suitable for use within the atmosphere, while rocket engines do not have this limitation.

[0019] Technical bottlenecks of aerospace engines and the current status of their R&D and technology:

[0020] 1. R&D Process: Research and development of aerospace engines in the global aviation and aerospace engine field began in the 1960s and has spanned 80 years. This research and development focuses on exploring the advantages and potential of scramjet engines. Russia and the United States are leading the way in this field. The US X-43 scramjet hypersonic program, launched in 1964, achieved a Mach 6.68 speed during flight testing in 1997, with a sustained burn of 10 seconds. In 1998, flight tests of the US-Russia joint scramjet hypersonic program achieved Mach 6.8 for 20 seconds and Mach 10 for 10 seconds.

[0021] In 2003, based on the test results obtained in the cooperation project with Russia, the United States restarted the X-43A hypersonic program. The flight test achieved test results of Mach 6.68, burning for 10 seconds, and Mach 9.86, burning for 3 seconds.

[0022] The above-mentioned experimental results of scramjet combustion in hypersonic airflow have proved the scientific nature and feasibility of hypersonic combustion.

[0023] 2. Four technical bottlenecks that are difficult to overcome: There are still four technical bottlenecks that are difficult to overcome when applying the experimental results to aerospace engines in aerospace engineering. These are: (1) Smooth transition and automatic switching of various modes of multi-mode cycle engines in the full envelope flight range; (2) Full mixing of air and fuel in various modes of flight; (3) Ignition at various Mach speeds during flight; and (4) Stable flame transmission and continuous combustion of the flame stabilizer.

[0024] 3. Flame stabilizer: Flame stabilization and sustained combustion have always been associated with the development of aerospace engines: The existing flame stabilizers in the field of aerospace engine technology in the world include: (1) the vertical fuel injection flame stabilizer on the wall of the scramjet engine combustion chamber invented and developed by Russia; (2) the support plate flame stabilizer; (3) the backward step flame stabilizer developed by the United States; (4) the cavity flame stabilizer developed by Russia, and the parallel and series multi-acoustic cavity flame stabilizers developed by the United States and Japan on this basis. The above-mentioned various flame stabilizers have their own strengths and weaknesses. They all cause total pressure loss to varying degrees, affect speed, increase structural mass, and have a small flame range. They are easily extinguished during high-Mach flight and are difficult to re-ignite under the impact of high-altitude hypersonic airflow.

[0025] 4. Cutting-Edge Flame Holder Technology: Experimental results from China's "Standing Oblique Detonation Pulsating Ramjet Hypersonic Combustion" and the United States' "Oblique Detonation Pulsating Ramjet Hypersonic Combustion" have solved the problem of flameout during hypersonic flight. This, in turn, presents three significant technical challenges: safe control of the flames detonated by liquid and solid propellants; ensuring the high-temperature resistance, detonation strength, and toughness of materials for high-intensity detonation shocks; and silencing the high-frequency detonation noise. Experts in the field of aerospace science and technology predict that this invention is still in the experimental stage and is far from practical application in aerospace engineering. Overcoming these three challenges will require several generations of dedicated research and development, as well as advancements in fundamental process and material technologies. Summary of the Invention

[0026] The purpose of the present invention is to break through the above four major technical bottlenecks and provide an air-breathing four-mode variable cycle engine single-stage orbital power fuselage integrated aerospace vehicle power system, referred to as the YRCW aerospace engine.

[0027] The present invention provides an aerospace vehicle power system, comprising an air intake system 1, a central engine 2, and two rear-mounted engines 3 on either side. The central engine 2 has a four-mode cyclic conversion function among ramjet turbine engine operating conditions, ramjet engine operating conditions, air-fuel rocket engine operating conditions, and liquid oxygen-fuel rocket engine operating conditions, and is referred to as the RCW engine 2. The rear-mounted engines 3 on either side have a four-mode cyclic conversion function among ramjet engine operating conditions, scramjet engine operating conditions, air-fuel rocket engine operating conditions, and liquid oxygen-fuel rocket engine operating conditions, and are referred to as the YRC engines 3.

[0028] A first air intake structure 2-1, a first diffuser 2-2, a first oil pan assembly 2-3, a first combustion chamber 2-4, a reluctance generator / motor 2-5, a turbine axial flow and centrifugal assembly disc 2-6, a second diffuser 2-7, a second oil pan assembly 2-8, a second combustion chamber 2-9 (i.e., an afterburner) and a first tail nozzle 2-10 are sequentially arranged along the axial direction of the RCW engine 2; a second air intake structure 3-1, a third diffuser 3-2, a third oil pan assembly 3-3, a third combustion chamber 3-4 and a second tail nozzle 3-5 are sequentially arranged along the axial direction of the YRC engine 3; the first air intake structure 2-1 of the RCW engine 2 is connected to one output end 1-2-2 of the air intake system 1; the second air intake structure 3-1 of the YRC engine 3 is connected to the other output end 1-3 of the air intake system 1;

[0029] The intake system 1 is a U-shaped intake duct. An S-shaped slope is provided in the middle of each of the two intake channels of the U-shaped intake duct. An airflow separation plate 1-1 is provided at the peak of each S-shaped slope. A first intake duct 1-2 and a second intake duct 1-3 are provided in parallel starting from the airflow separation plate 1-1. A first intake duct 1-2 provided on each side of the U-shaped intake duct extends to the bottom of the U-shaped duct to form an output end 1-2-2 of the intake system. The output end 1-2-2 is connected to the central intake duct piston inner sleeve 1-4. It is connected to the first intake structure 2-1 of the RCW engine 2, and a straightening cone 1-2-1 is provided at the U-shaped bottom of the intake system; a second intake duct provided on each side of the U-shaped intake duct serves as another output end 1-3 of the intake system and is respectively connected to the second intake structure 3-1 of the YRC engine 3 on both sides; the cross-sectional area of ​​the central intake duct piston inner sleeve 1-4 is 1 / 5-1 / 8 of the intake duct cross-sectional area; preferably, the cross-sectional area of ​​the central intake duct piston inner sleeve 1-4 is 1 / 7 of the intake duct cross-sectional area.

[0030] The RCW engine 2 housing at the output end of the turbine axial flow and centrifugal assembly disc 2-6 is provided with a valve mechanism, which includes a compressed cold air front delivery pipe 2-6-2-3, a compressed cold air rear delivery pipe 2-6-2-4, a compressed hot exhaust gas front delivery pipe 2-6-2-5, a compressed hot exhaust gas rear delivery pipe 2-6-2-6 and a vector airflow distribution system 4; the compressed cold air front delivery pipe 2-6-2-3 is connected to the housing of the RCW engine 2 at one end, and is connected to the turbine axial flow and centrifugal assembly disc 2-6 at the output end of the turbine axial flow and centrifugal assembly disc 2-6. The compressed cold air is obtained from the output end of the turbine axial flow and centrifugal assembly disc 2-6, and the other end is connected to the first cold air collecting cover 2-1-5 in the first air intake structure 2-1 of the RCW engine 2, providing the oxygen-containing compressed air required for the air-fuel-rocket closed combustion for the combustion chamber of the RCW engine 2; one end of the compressed cold air rear delivery pipe 2-6-2-4 is connected to the housing of the RCW engine 2, and the compressed cold air is obtained from the output end of the turbine axial flow and centrifugal assembly disc 2-6, and the other end is connected to the second cold air in the second air intake structure 3-1 of the YRC engine 3. The air collecting hood 3-1-5 is connected to provide the oxygen-containing compressed air required for the air-fuel-rocket closed combustion for the combustion chamber of the YRC engine 3; one end of the compressed hot exhaust gas front delivery pipe 2-6-2-5 is connected to the RCW engine housing, and the hot exhaust gas is obtained from the output end of the turbine axial flow and centrifugal assembly disc 2-6, and the other end is connected to the first hot exhaust gas collecting hood 2-1-6 in the first air intake structure 2-1 of the RCW engine 2. The output of the first hot exhaust gas collecting hood 2-1-6 provides the combustion chamber wall of the RCW engine 2 with oxygen. Cooling gas; one end of the compressed hot exhaust gas delivery pipe 2-6-2-6 is connected to the RCW engine casing, and obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disk 2-6, and the other end is connected to the second hot exhaust gas collecting cover 3-1-6 of the YRC engine 3 to provide cooling gas for the combustion chamber wall of the YRC engine 3; the vector airflow distribution system 4 obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disk 2-6, and provides vector jet airflow for horizontal take-off and landing, vertical take-off and landing and aerospace maneuvers of aerospace vehicles.

[0031] The present invention relates to an aerospace vehicle power system. In the RCW engine 2, the first air intake structure 2-1 includes: a central air intake duct 2-1-3, a turbine axial flow air intake duct 2-1-4, a first cold air collecting cover 2-1-5, and a first hot exhaust gas collecting cover 2-1-6. The first cold air collecting cover 2-1-5 is provided with a plurality of first conveying support channels 2-1-7. The compressed cold air of the first cold air collecting cover 2-1-5 provides the first combustion chamber 2-4 of the RCW engine 2 with fresh compressed air for closed combustion required for the air-fuel-rocket engine working condition through the plurality of first conveying support channels 2-1-7. The first hot exhaust gas collecting cover 2-1-6 is provided with a plurality of second conveying support channels 2-1-8. The compressed hot exhaust gas of the first hot exhaust gas collecting cover 2-1-6 is provided with the RCW combustion chamber wall and the turbine axial flow The centrifugal assembly disc 2-6 provides cooling airflow; the central air intake duct 2-1-3 and the turbine axial flow air intake duct 2-1-4 are provided in the first cold air collecting cover 2-1-5 to form two inner and outer air intake channels; an air intake truncated cone port 2-1-1 and a conical ram air intake valve 2-1-2 matching the air intake truncated cone port 2-1-1 are provided at the port of the central air intake duct 2-1-3, and the axial sliding of the conical ram air intake valve 2-1-2 is controlled by the hydraulic system to open or close the conical ram air intake valve 2-1-2 and the air intake truncated cone port 2-1-1; the number of the first conveying support channels 2-1-7 is selected from 4 to 18 according to the designed engine power parameters, and 12 are selected in the embodiment of the present invention; the number of the second conveying support channels (2-1-8) is selected from 3 to 12 according to the designed engine power parameters, and 6 are selected in the embodiment of the present invention.

[0032] The present invention relates to an aerospace vehicle power system, wherein the turbine axial flow and centrifugal assembly disk 2-6 in the RCW engine 2 comprises a front turbofan disk group 2-6-1 and a rear fan-out disk group 2-6-2 mounted on a reluctance generator / motor 2-5 and a main shaft, wherein the turbofan disk group 2-6-1 is composed of a plurality of turbofan disks and guide disks stacked in sequence to form an axial flow air intake compressor; the fan-out disk group 2-6-2 is composed of a first fan-out disk 2-6-2-1 and its guide disk and a second fan-out disk 2-6-2-2 and its guide disk stacked in sequence to form a centrifugal compressor; The RCW engine casing at 6-2-1 is connected to the front duct 2-6-2-3 for compressed cold air and the rear duct 2-6-2-4 for compressed cold air; the front duct 2-6-2-5 for compressed hot exhaust gas and the rear duct 2-6-2-6 for compressed hot exhaust gas are connected to the RCW engine casing at the second fan disc 2-6-2-2; each stage of the turbofan disc, fan disc and guide disc is provided with three isolation sealing rings 2-6-3-1, 2-6-3-2 and 2-6-3-3 with different diameters in the radial direction, which separate the turbofan disc, fan disc and guide disc into The four ducts are defined as the first duct 2-6-4-1, the second duct 2-6-4-2, the third duct 2-6-4-3 and the fourth duct 2-6-4-4 from the center to the circumference; the first, second and fourth ducts of the turbofan disk group 2-6-1 are provided with hollow supersonic turbofan blades and guide blades 2-6-5-1, 2-6-5-2 and 2-6-5-4, and the third duct is provided with power turbine blades and guide blades 2-6-5-3; the first and second ducts of the rear-mounted two-stage fan-off disk group 2-6-2 are provided with hollow supersonic turbofan blades, and the Duct 3 is provided with power turbine blades and guide vanes, and duct 4 is provided with centrifugal blades and guide vanes 2-6-5-5 and 2-6-5-6; the hollow supersonic turbofan blades, power turbine blades, and guide vanes in ducts 1, 2, and 3 are respectively connected to two adjacent isolation sealing rings constituting the duct, and the centrifugal blades and guide vanes in duct 4 are connected at one end to isolation sealing ring 2-6-3-3 and at the other end to the outer ring of the fan disc; the number of the turbofan discs and guide discs is selected from 3 to 12 according to the designed engine power, and 3 is selected in the embodiment of the present invention.

[0033] The turbofan disc assembly 2-6-1 is a hybrid disc consisting of three-stage axial compressor hollow supersonic turbofan blades, hollow power turbine blades and guide vanes.

[0034] Duct Airflow Distribution: When the turbofan disk diameter is 1.2 meters, the headwind airflow to ducts 1, 2, and 4 is distributed based on the cross-sectional area of ​​the predetermined duct annular diameter difference. Duct 1 receives 18-19% of the total headwind airflow, duct 2 receives 60%, and duct 4 receives 21-22%. (Annular diameter difference is the difference between the diameter of the first isolation seal ring and the turbofan base disk diameter; the difference between the diameter of the third isolation seal ring and the second isolation seal ring diameter; and so on for the other ducts.)

[0035] The deflector discs are located between each stage of the turbofan discs. The fourth outer ring of the deflector discs is the engine casing, while the deflector disc casing of the fan-off disc is the centrifugal compressor casing. The front of each stage casing houses the first turbofan disc. Each stage casing is secured to the preceding and succeeding stage casings with isolation and sealing rings, secured with screws for easy removal and maintenance.

[0036] Power turbine temperature drop: Turbofan disks are generally 3-stage, and each stage of the power turbine can reduce the gas temperature by about 90-110°C. The same 3-stage turbofan disk can reach 5-7 or even 9 stages or more according to the engine power requirements, so as to output greater power and minimize the exhaust gas temperature.

[0037] The fan-out disk: There are two levels of fan-out disks behind the multi-stage turbofan disk:

[0038] 1. First-stage fan disc 2-6-2-1: This is welded together with the hollow supersonic turbofan blades and guide vanes of the first and second ducts, the power turbine blades and guide vanes of the third duct, and the centrifugal compressor blades and guide vanes of the fourth duct, along with the isolation and sealing rings of each duct. High-pressure cold air from the fourth duct is delivered via forward delivery duct 2-6-2-3 to the central first combustion chamber, providing high-pressure fresh air for closed-loop combustion. The compressed cold air rear delivery duct 2-6-2-4 supplies the high-pressure air required for the air-fuel-rocket operation of the ramjets on both sides.

[0039] 2. Second-stage fan disc 2-6-2-2: The hollow supersonic turbofan blades and guide vanes of the first and second ducts, the power turbine blades and guide vanes of the third duct, and the centrifugal blades and guide vanes of the fourth duct centrifugal compressor are welded to the isolation sealing rings of each duct. The third duct centrifugal compressor provides high-pressure, medium-temperature cooling gas of about 450°C-500°C to the central first and third combustion chambers through the forward delivery pipe 2-6-2-5 and the rear delivery pipe 2-6-2-6.

[0040] The compressed cold air of the first duct enters the first fan-out disk 2-6-2-1 and is then introduced into the second combustion chamber by its guide disk; the compressed cold air of the second duct is mixed with the gas of the third duct and enters the second fan-out disk 2-6-2-2, and is then introduced into the vector airflow distribution air collecting hood by its guide disk; the compressed cold air of the fourth duct enters the first fan-out disk 2-6-2-1 and is then guided out by its guide disk, and all the way through the compressed cold air front delivery pipe 2-6-2-3 to provide the RCW engine combustion chamber with the oxygen-containing compressed air required for the air-fuel-rocket closed combustion. The other route provides the YRC engine's combustion chamber with oxygen-containing compressed air required for the air-fuel-rocket closed combustion through the compressed cold air rear delivery pipe 2-6-2-4; part of the compressed cold air from the 4th duct is mixed with part of the fuel gas from the 3rd duct and enters the second fan-out disk 2-6-2-2, and is discharged through its guide disk. One route provides cooling gas for the RCW engine combustion chamber wall through the compressed hot exhaust gas front delivery pipe 2-6-2-5, and the other route provides cooling gas for the YRC engine's combustion chamber wall through the compressed hot exhaust gas rear delivery pipe 2-6-2-6.

[0041] The present invention provides an aerospace vehicle power system, wherein the first oil pan assembly 2-3 is composed of an annular oil spray disc 2-3-1-3 and six annular oil quantity control pressure oil pipe channels 2-3-1-2 evenly distributed around the annular oil spray disc 2-3-1-3; one side of the annular oil spray disc 2-3-1-3 is provided with three groups of large, medium and small annular oil spray nozzles 2-3-1, and the center is provided with a silane igniter nozzle 2-3-1-1, and the oil spray nozzle 2-3-1 is a piston structure consisting of an oil control inner heat pipe and an outer sleeve.

[0042] The first combustion chamber 2-4 is provided with a ramjet burner 2-4-1, and the combustion chamber 2-4 is provided with 6 gas delivery pipes 2-4-2 to deliver the burned gas in the ramjet burner 2-4-1 to the annular gas collecting groove 2-4-3, and the gas is guided by the guide vanes in the annular gas collecting groove 2-4-3 and then sprayed to the third duct power turbine blade through the guide nozzle.

[0043] In the RCW engine, the second diffuser 2-7, the second oil pan assembly 2-8, and the second combustion chamber 2-9 form a whole; an air intake duct 2-7-1, an air intake duct piston sleeve 2-7-2, and a conical ram air intake valve 2-7-3 are arranged in front of the second diffuser 2-7; the second combustion chamber 2-9 is formed by two combustion tubes to form an inner ring combustion chamber 2-9-1, an outer ring combustion chamber 2-9-2 and a gaseous pneumatic plug 2-9-3 for stabilizing the fire source of the outer ring, and the tail nozzle bell mouth of the RCW engine tail nozzle 2-10 is connected into a whole.

[0044] The second oil pan assembly 2-8 is composed of an inner oil spray pan 2-8-1, an outer oil spray pan 2-8-2 and 6 annular oil quantity control pressure oil pipe channels 2-8-4 evenly distributed around the outer oil spray pan 2-8-2; the inner oil spray pan 2-8-1 and one side of the outer oil spray pan 2-8-2 are evenly distributed with oil nozzles 2-8-5, and a silane igniter nozzle 2-8-3 is arranged in the center of the inner oil spray pan 2-8-1. The oil nozzles 2-8-5 on the inner oil spray pan 2-8-1 and the outer oil spray pan 2-8-2 are piston structures composed of an oil control inner heat pipe and an outer sleeve.

[0045] The inner fuel injection disc 2-8-1 provides fuel for the inner ring combustion chamber 2-9-1, and the outer fuel injection disc 2-8-2 provides fuel for the outer ring combustion chamber 2-9-2. The fuel in the inner ring combustion chamber 2-9-1 is ignited by the silane igniter nozzle 2-8-3 in the center of the second combustion chamber 2-9, and its combustion gas serves as a stable fire source for the outer ring combustion chamber 2-9-2 to form a gaseous pneumatic plug 2-9-3. The expansion and contraction of the jet flow after the tail nozzle is adjusted by adjusting the amount of fuel in the inner fuel injection disc 2-8-1.

[0046] The present invention relates to an aerospace vehicle power system. In the YRC engine 3, the second air intake structure 3-1 includes an air intake truncated cone 3-1-2, a conical ram intake valve 3-1-1 and an outer duct transcendental ram intake 3-1-3, a second cold air collecting cover 3-1-5, and a second hot exhaust gas collecting cover 3-1-6; the second cold air collecting cover 3-1-5 is installed at one end of the YRC engine 3 casing, and is located at the YRC engine casing portion corresponding to the position of the third diffuser chamber 3-2; the air intake truncated cone 3-1-2 is arranged on one side of the second cold air collecting cover 3-1-5, the truncated cone ram intake valve 3-1-1 is located in the center of the second cold air collecting cover 3-1-5 and matches the air intake truncated cone port 3-1-2. The axial sliding of the conical ramjet air intake valve 3-1-1 is controlled by the hydraulic system to open or close the conical ramjet air intake valve 3-1-1 and the air intake truncated cone port 3-1-2. When closed, the YRC engine is in the working condition of an air fuel rocket engine or a liquid oxygen fuel rocket engine. When opened, the YRC engine is in the working condition of a ramjet or a scramjet. The air intake truncated cone port 3-1-2 is connected to the second air intake duct 1-3 of the air intake system 1, providing the required ramjet content for the ramjet and scramjet working conditions of the YRC engine. Oxygen fresh air, a plurality of third delivery support channels 3-1-7 are provided in the second cold air collecting cover 3-1-5, and the compressed cold air entering the second cold air collecting cover 3-1-5 through the compressed cold air delivery pipe 2-6-2-4 is connected with the third diffusion chamber 3-2 through the plurality of third delivery support channels 3-1-7, providing the third combustion chamber 3-4 of the YRC engine 3 with fresh compressed air for closed combustion required for the working condition of the air-fuel rocket engine; the second compressed hot exhaust gas collecting cover 3-1-6 is set on the YRC engine casing, located at the YRC engine casing part corresponding to the position of the injection plate 3-3-2. Position, a plurality of fourth conveying support channels 3-1-8 are provided in the second hot exhaust gas collecting hood 3-1-6, and the compressed hot exhaust gas entering the second hot exhaust gas collecting hood 3-1-6 through the conveying pipe 2-6-2-6 after the compressed hot exhaust gas provides cooling airflow for the YRC combustion chamber wall through the plurality of fourth conveying support channels 3-1-8; the number of the third conveying support channels 3-1-7 is selected from 4 to 18 according to the designed engine power parameters, and 12 are selected in the embodiment of the present invention; the number of the fourth conveying support channels (3-1-8) is selected from 3 to 12 according to the designed engine power parameters, and 6 are selected in the embodiment of the present invention.

[0047] The third combustion chamber 3-4 in the YRC engine 3 is composed of a rocket combustion tube 3-4-1, a rocket ramjet combustion tube 3-4-2, and an outer duct transcendental ramjet combustion tube 3-4-3, which respectively constitute a rocket combustion chamber, a ramjet combustion chamber, and an outer duct transcendental ramjet combustion chamber, and the three combustion tubes have different diameters; when the YRC engine is in rocket and ramjet working conditions, the outer duct 3-4-6 serves as an airflow cooling channel for the rocket ramjet combustion tube 3-4-2; in the transcendental ramjet working condition, the outer duct 3-4-6 is a mixing channel for air and fuel, and transcendence occurs at the tail end 3-4-5 of the outer duct 3-4-6, and the tail end of the transcendental ramjet combustion tube 3-4-3 is extended and connected to the bell mouth of the second tail nozzle 3-5 to form a whole.

[0048] The YRC engine 3, the third fuel injection disc 3-3 includes an integrally arranged rocket fuel injection disc 3-3-1, a ram fuel injection disc 3-3-2 and a super-ram fuel injection disc 3-3-3 arranged at the super-ram fuel intake port 3-1-3 of the outer duct, the rocket fuel injection disc 3-3-1 has a smaller diameter and is coaxially embedded in the ram fuel injection disc 3-3-2, the rocket fuel injection disc 3-3-1 and the ram fuel injection disc 3-3-2 are both provided with at least one circle of fuel injection nozzles 3-3-1-1, 3-3-2-2; a silane injection igniter 3-3-4 is provided at the center of the rocket fuel injection disc 3-3-1; the fuel injection nozzles 3-3-1-1 and 3-3-2-2 are piston structures composed of an oil-control inner heat pipe and an outer sleeve; 6 hydraulic pipes, fuel, and hot exhaust gas support channels 3-3-5 are evenly distributed around the ram fuel injection disc 3-3-2.

[0049] The rocket fuel injection disc 3-3-1 provides fuel for the rocket combustion chamber, and the ramjet fuel injection disc 3-3-2 provides fuel for the ramjet combustion chamber. When the fuel in the rocket combustion chamber is ignited by the silane jet igniter 3-3-4 located in the center of the rocket fuel injection disc 3-3-1, its combustion gas becomes a stable fire source for the ramjet combustion chamber, forming a gaseous pneumatic plug 3-4-4. By adjusting the amount of fuel in the rocket fuel injection disc 3-3-1, the expansion and contraction of the jet stream behind the tail nozzle 3-5 can be adjusted.

[0050] The super ramjet fuel injection disc 3-3-3 is provided with at least one circle of fuel injectors 3-3-3-1, and the super ramjet fuel injection disc 3-3-3 is arranged on the inner edge of the super ramjet air inlet 3-1-3 of the outer duct. When the speed node of super ramjet startup is reached, a large amount of oxygen-containing ramjet air and the fuel injected by the annular fuel injector 3-3-3-1 are fully mixed in the super ramjet outer duct 3-4-6, effectively extending the mixing distance of the fuel and air, thereby greatly improving the full mixing of the ramjet air and the fuel, and burning at the junction of the tail end of the outer duct 3-4-6 and 3-4-5, thereby improving the efficiency of complete combustion.

[0051] The present invention relates to an aerospace vehicle power system, wherein the vector airflow distribution system 4 is provided with a vector airflow distribution air collecting cover 4-1, and the vector airflow distribution air collecting cover 4-1 is provided on the outer shell of the centrifugal compressor at the second fan-out disk 2-6-2-2 of the RCW engine. After the compressed cold air of the second duct is mixed with the gas of the third duct and enters the second fan-out disk 2-6-2-2, it is introduced into the vector airflow distribution air collecting cover by its guide disk 4-1-7. The vector airflow distribution air collecting cover 4-1 is provided with an axially fixed guide sleeve throat 4-1-9 and an axially movable guide sleeve throat 4-1-10. The vector airflow distribution air collecting cover 4-1 is provided with a vector airflow distribution air collecting cover 4-1 shell. The vector airflow bend 4-1-1 is provided with a vector airflow distributor 4-1-2, which distributes the hot exhaust gas in the vector airflow distribution air hood 4-1 to the front left lower spray vector airflow duct 4-1-4, the rear left lower spray vector airflow duct 4-1-3, the front right lower spray vector airflow duct 4-1-5, and the rear right lower spray vector airflow duct 4-1-6 connected to the outlet end of the vector airflow bend 4-1-1, providing vector jet airflow for the three-point lower spray vector shutters 4-8, 4-9, 4-10, the horizontal vector nozzles 4-2, 4-3, 4-4, 4-5, and the vector nozzles 4-6 and 4-7 rotating around the horizontal direction of the left and right delta wings.

[0052] When the axially movable guide sleeve throat 4-1-10 moves forward and is sealed with the outer shell 4-1-11 of the guide disk of the second fan-out disk 2-6-2-2, and the wave peaks of the axially movable guide sleeve throat 4-1-10 and the axially fixed guide sleeve throat 4-1-9 are misaligned, the second ducted mixed hot air flow passes through the radial annular and axial curved channel 4-1-12 between the axially fixed guide sleeve throat 4-1-9 and the axially movable guide sleeve throat 4-1-10, and is sprayed toward the outer wall of the second combustion chamber to cool the combustion tube, and then is expanded and ejected through the central trumpet mouth, thereby forming vector thrust for low-altitude and low-speed cruising.

[0053] When the axially movable guide sleeve throat 4-1-10 moves backward and is sealed with the wave crest of the axially fixed guide sleeve throat 4-1-9, the second ducted mixed hot air flow enters the vector air flow distribution air collecting hood 4-1 and is then output through the vector air flow bend 4-1-1, forming a vector power source for the aircraft's take-off and landing, hovering and aerospace maneuvers.

[0054] Advantages of the present invention

[0055] Compared with the world's aerospace engine technology field, the present invention has the following outstanding advantages:

[0056] 1. The air intake system 1 of the aerospace engine (abbreviated as YRCW) of the present invention connects the central engine RCW2 and the left and right rear engines YRC3 into an integral whole, and provides compressed cold air and compressed hot exhaust gas to the central engine RCW and the rear engines YRC respectively through the turbine axial flow and centrifugal assembly disks 2-6 provided in the central engine RCW; by regulating the working state of the air intake structure of the central engine RCW and the rear engines YRC, the combustion working conditions of their combustion chambers are changed to be in a closed or open state, and at the same time, combined with the supply of compressed cold air, fuel and liquid oxygen input, the four-mode conversion of the central engine RCW to a ramjet turbine engine, a ramjet engine, an air-fuel rocket engine, and a liquid oxygen-fuel rocket engine is achieved, and the four-mode conversion of the rear engines YRC to a ramjet engine, a scramjet engine, an air-fuel-rocket engine, and a liquid oxygen-fuel-rocket engine is achieved. The YRCW aerospace engine combines the advantages of multiple engines, which are interdependent and non-burdensome. It fully utilizes the buoyancy of the air, oxygen and working medium in the atmosphere, reduces the fuel load and the weight of the aircraft, and improves the aircraft's payload ratio. It has a flight speed of more than 25 Mach, a flight altitude of more than 80,000 kilometers, and a manufacturing cost of less than US$100,000.

[0057] 2. Unlike traditional turbine engines, the RCW2's combustion chamber features a ramjet-style burner, capable of operating in air-breathing ramjet, air-fuel-rocket, ramjet, and liquid oxygen-fuel-rocket modes. The RCW engine's ramjet-style combustor features the burner at the front, with the turbine and axial-flow and centrifugal compressors integrated into a single disc at the rear. This structural advantage allows for adjustable internal engine operating mode switching. Specifically, the geometric positions of the operating mode switching valves for the intake system's piston inner sleeves 1-4, intake cone 2-1-1, ramjet-type intake valve 2-1-2, the 12 cold air support channels 2-1-7 of the cold air hood 2-1-5, the six hot exhaust support channels of the hot exhaust hood 2-1-6, and the six fuel support channels 2-8-4 are adjustable, enabling the engine to switch between various operating modes.

[0058] The traditional turbine engine has the intake compressor in the front, the combustion chamber in the middle, and the power turbine in the back, so its combustion chamber can only adapt to the turbine engine working conditions and cannot adapt to the conversion of four-mode working conditions.

[0059] 3. The high-temperature, high-pressure airflow from the ramjet or rocket-type combustion chamber combustion tubes of the turbine axial-flow and centrifugal assembly disks 2-6 in the central engine RCW impacts the turbine of the turbofan disk, driving the axial-flow intake compressor and the centrifugal compressor of the fan-out disk. The compressed airflow is distributed through four ducts, and enters the high-pressure cold air collection hood and the hot exhaust gas collection hood, respectively, to provide compressed cold air for closed combustion, medium-temperature, high-pressure cooling gas, and vectored injection gas for the RCW and YRC engines, respectively. This completes the multi-mode conversion and thrust afterburning, vectored injection, exhaust gas cooling, air compression, and air cooling functions of the RCW and YRC engines, enabling various maneuvering functions of the aircraft, including horizontal take-off and landing, vertical take-off and landing, hovering, rapid translation, high-speed U-turn, aerial cruising, space shuttle, and satellite-style patrol.

[0060] 4. The two-stage annular oil spray discs 2-8-1 and 2-8-2 provided in the RCW second combustion chamber of the YRCW engine of the present invention are stacked together with the two-stage annular combustion tubes 2-9-1 and 2-9-2 to form a whole. The two-stage annular oil spray discs 3-3-1 and 3-3-2 provided in the third combustion chamber of the YRC engine are stacked together with the two-stage annular combustion tubes 3-4-1 and 4-1-2 to form a whole. Once the mixed oil and gas in the two-stage annular combustion chamber is ignited by the silane igniter on the central support seat of the two-stage oil pan, the gas in the closed combustion of the combustion chamber acts as a stable fire source, becoming a gaseous pneumatic plug of the flame stabilizer, providing an ignition source for the engine during mode conversion.

[0061] The gaseous aerodynamic plug of the present invention is composed of closed combustion gas instead of directly using air. Therefore, it is not affected by atmospheric pressure and flight speed, has no structural mass, does not cause total pressure loss, has adjustable flame size, and burns continuously and stably. It can adapt to the different flight speed requirements of aerospace vehicles within the full flight envelope, and ensure the reliability of ignition and the continuous stability of combustion during the mode conversion of the engine under supersonic and hypersonic conditions. The present invention can adjust the position, shape, and size of the gaseous aerodynamic plug by adjusting the amount of fuel injected by the fuel injection disc. When the amount of fuel injected by the inner ring fuel injection disc of the combustion chamber is small, the tail cone formed is short, and the gas ejected from the outer ring combustion chamber shrinks toward the center. The plug-type wake flow has a high speed and is suitable for high-altitude and high-speed conditions. When the amount of fuel injected by the inner ring fuel injection disc is large, the tail cone formed is long, or even trumpet-shaped, and the gas ejected from the outer ring combustion chamber expands outward, which is suitable for low-altitude and low-speed conditions.

[0062] This invention breaks through the bottleneck of detonation flameout in supersonic and hypersonic conditions in the field of aerospace engine technology. By employing a pneumatic plug design, it solves the problem of the Laval tube's throat diameter being difficult to adjust. This overcomes the shortcomings of solid plugs currently being researched in the aerospace engine community, including difficulty adjusting and cooling, heavy structural mass, high total pressure loss, and a small, easily extinguished flame.

[0063] 5. The medium-temperature and high-pressure hot exhaust gas from the vector airflow distribution hood 4-1 set in RCW2 is output through the vector airflow distributor 4-1-2, and is transported to the front three-point downward spray vector airflow shutters, the horizontal nozzles on both sides, and the tail delta wing rotating nozzle through the front transmission pipe and the left and right pipes to maintain the balance of the aircraft's horizontal take-off and landing, vertical take-off and landing, hovering in the air, steady-state flight, and aerospace maneuvers.

[0064] In today's world of aviation engines, representative aircraft with advanced vertical take-off and landing capabilities include: Sea Harrier, Yak-141, F35, helicopters and Osprey V-22. Their vertical take-off and landing and horizontal injection systems have heavy structural weight, high fuel consumption, large mechanical transmission losses, and extremely unreliable conversion between vertical injection and horizontal injection. The load capacity, range, and horizontal flight speed are all limited. Both vertical take-off and landing and hovering in the air consume a lot of fuel.

[0065] The guide vanes of the vectored airflow louvers of the present invention have a wide range of backward and forward tilt angles, require no rotating device, and feature a simple structure, low bulk, and low fuel consumption. Because the vectored airflow nozzles of the vectored airflow injection system produce no air backflow, the fresh air in the first combustion chamber of the first duct has a normal oxygen content, complete combustion, and rapid acceleration. After reaching Mach 0.2-0.4 in 5-10 seconds, the louvers can be closed, and the vectored airflow is converted to a rearward jet to cool the outer casing of the second combustion chamber, thereby generating vectored thrust for the aircraft's low-speed cruise. Once the engine reaches Mach 0.6 or higher, the ramjet operation of the first and second combustion chambers of the central RCW can be initiated.

[0066] The front, back, left, and right horizontal vector nozzles on the YRCW aerospace aircraft's wing surface can translate quickly and swiftly from side to side when hovering, and can rotate on the spot faster than helicopters that rely on tail propellers. Its various aerospace maneuverability is stronger than that of existing flight platforms.

[0067] 6. When the air is at an altitude of 60 kilometers or above and the atmosphere is close to vacuum, the central air intake duct piston inner sleeve 1-4 moves forward and closes with the top of the rectifier cone to seal the central air intake duct and prevent the hot gas of the central engine from leaking forward.

[0068] The axial sliding of the conical ram air intake valve 2-1-2 is controlled by the hydraulic system to open or close the conical ram air intake valve 2-1-2 and the intake cone port 2-1-1 to switch the combustion conditions of the combustion chamber.

[0069] When the flight speed of the aircraft is between 0 and 0.2 Mach, the hydraulic system keeps the conical ramjet valve 2-1-2 closing the intake truncated cone 2-1-1, and the incoming air flows along the outer edge of the intake truncated cone into the turbine axial flow intake duct 2-1-4. The turbine rotates at high speed, and the hollow supersonic turbofan blades, guide vanes and centrifugal blades compress the fresh cold air in the 1st, 2nd and 4th ducts, and outputs the high-pressure fresh cold air through the forward delivery pipe 2-6-2-3 of the first fan disc 2-6-2-1, providing the RCW engine combustion chamber with the oxygen-containing compressed air required for the closed combustion of air-fuel-rocket. The RCW central engine combustion chamber then forms a closed combustion air-fuel rocket working condition.

[0070] When the aircraft speed reaches 0.2 Mach or above, the cross-sectional area of ​​the central inlet piston inner sleeve 1-4 is only 0.28m 2 , is the cross-sectional area of ​​the air inlet channel 2m 2 At 1 / 7 of the original engine's mass, the airflow within the smaller central intake piston inner sleeve 1-4, driven by the intake of the turbine's axial-flow and centrifugal discs 2-6, reaches a velocity of approximately Mach 1. At this point, headwind pressure pushes the hydraulically deactivated conical ramjet valve 2-1-2 back into the first diffuser chamber, allowing airflow to enter the first diffuser chamber directly through the intake truncated cone 2-1-1. This deceleration causes a sharp increase in pressure, establishing ramjet combustion conditions in the first and second RCW combustion chambers. The design of the central intake piston inner sleeve 1-4 overcomes the principle that a ramjet engine must reach Mach 0.5 for proper operation, effectively reducing fuel consumption.

[0071] 7. The YRCW boasts a high takeoff mass-to-payload ratio. The YRCW's propulsion system integrates a single-stage YRCW engine into a single-stage, orbital power unit, maximizing its mass-to-payload ratio. The YRCW can be equipped with one or two additional ramjet-rocket engines on either side to meet transportation needs. A spaceport can be established in low-Earth orbit to provide refueling, oxygenation, and the replenishment of vital supplies for the YRCW's journey to the moon.

[0072] The moon's orbital velocity relative to Earth is Mach 2.8. When the YRCW vehicle is idling in space, the hot exhaust from the second duct is ejected through a micro-Laval tube, accelerating it to 38 m / s (12.93 km / s). It takes three days to reach a distance of 384,000 km from the moon. When approaching the moon by 4,000 to 10,000 km, the YRCW can turn around and slow down, landing on the moon in a hovering vertical landing. BRIEF DESCRIPTION OF THE DRAWINGS

[0073] Figure 1 It is a schematic diagram of the overall structure of the present invention.

[0074] Figure 2 It is a schematic diagram of the central engine structure in the present invention.

[0075] Figure 3 for Figure 2 Schematic diagram of the layout structure of the central engine's first air intake structure, first combustion chamber, turbine axial flow and centrifugal assembly disc.

[0076] Figure 4 for Figure 3 Schematic diagram of the layout structure of the first air intake structure and the first combustion chamber of the central engine.

[0077] Figure 5 This is a schematic structural diagram of the first oil pan assembly of the central engine in the present invention.

[0078] Figure 6 This is a schematic diagram of the turbine axial flow and centrifugal assembly disc structure of the central engine of the present invention.

[0079] Figure 7 It is a schematic diagram of the valve train structure of the central engine in the present invention.

[0080] Figure 8 This is a schematic diagram of the structure of the second combustion chamber of the central engine in the present invention.

[0081] Figure 9 This is a schematic structural diagram of the second oil pan assembly of the central engine in the present invention.

[0082] Figure 10 It is a schematic diagram of the structure of the rear-mounted YRC engine in the present invention.

[0083] Figure 11 This is a schematic diagram of the structure of the third fuel injection plate of the rear-mounted YRC engine in the present invention.

[0084] Figure 12 This is a schematic diagram of the structure of the Chaoran ramjet fuel injection disc in the rear-mounted YRC engine of the present invention.

[0085] Figure 13 Schematic diagram of the structure of the vector airflow distribution system in the present invention.

[0086] In the picture:

[0087] 1—Intake system;

[0088] 1-1--air flow separation plate, 1-2--first air intake duct, 1-3--second air intake duct, 1-4--central air intake duct piston inner sleeve;

[0089] 2---Central engine;

[0090] 2-1---First air intake structure, 2-2--First diffuser, 2-3--First oil pan assembly, 2-4--First combustion chamber, 2-5--Reluctance generator / motor, 2-6--Turbine axial flow and centrifugal assembly disc, 2-7--Second diffuser, 2-8--Second oil pan assembly, 2-9--Second combustion chamber, 2-10--First tail nozzle;

[0091] 3---rear engine;

[0092] 3-1--Second air intake structure, 3-2--Third diffuser, 3-3--Third fuel injection disc, 3-4--Third combustion chamber, 3-5--Second tail nozzle;

[0093] 4---Vector air flow distribution system;

[0094] 4-1--Vector air flow distribution hood, 4-2, 4-3, 4-4, 4-5--Horizontal vector nozzles, 4-6, 4-7 Rotating vector nozzles, 4-8, 4-9, 4-10--Downward spray vector shutters. DETAILED DESCRIPTION

[0095] See also Figure 1-13 The present invention provides an aerospace vehicle power system, comprising an air intake system 1, a central engine 2, and two rear-mounted engines 3 on either side. The central engine 2 has a four-mode conversion function among a ramjet, a ramjet, an air-fuel rocket engine, and a liquid oxygen-fuel rocket engine, referred to as an RCW engine. The rear-mounted engines on either side have a four-mode conversion function among a ramjet, a scramjet, an air-fuel rocket engine, and a liquid oxygen-fuel rocket engine, referred to as a YRC engine.

[0096] See also Figure 2 Along the axial direction of the RCW engine 2, a first air intake structure 2-1, a first diffuser 2-2, a first oil pan assembly 2-3, a first combustion chamber 2-4, a reluctance generator / motor 2-5, and a turbine axial flow and centrifugal assembly disc 2-6 provided on the main shaft of the reluctance generator / motor, a second diffuser 2-7, a second oil pan assembly 2-8, a second combustion chamber 2-9, and a first tail nozzle 2-10 are sequentially provided; along the axial direction of the YRC engine 3, a second air intake structure 3-1, a third diffuser 3-2, a third oil injection disc 3-3, a third combustion chamber 3-4, and a second tail nozzle 3-5 are sequentially provided;

[0097] See also Figure 1The intake system 1 is a U-shaped intake duct. An S-shaped slope is provided in the middle section of the two intake channels of the U-shaped intake duct, and an air flow separation plate 1-1 is provided at the peak of each S-shaped slope; a first intake duct 1-2 and a second intake duct 1-3 are provided in parallel starting from the air flow separation plate 1-1; a first intake duct 1-2 provided on each side of the U-shaped intake duct extends to the bottom of the U-shape and connects to form an output end 1-2-2 of the intake system, and the output end 1-2-2 is connected to the first intake structure 2-1 of the RCW engine 2 through the central intake duct piston inner sleeve 1-4, and a straightening cone 1-2-1 is provided at the U-shaped bottom of the intake system; a second intake duct 1-3 provided on each side of the U-shaped intake duct serves as another output end of the intake system and is respectively connected to the second intake structure 3-1 of the YRC engine 3 on both sides; the cross-sectional area of ​​the central intake duct piston inner sleeve 1-4 is 0.28m 2 , is the cross-sectional area of ​​the air inlet duct 2m 2 1 / 7 of

[0098] See also Figure 6 、 78. In the RCW engine, a valve mechanism is provided on the housing of the RCW engine 2 at the output end of the turbine axial flow and centrifugal assembly disc 2-6. The valve mechanism includes a compressed cold air front delivery pipe 2-6-2-3, a compressed cold air rear delivery pipe 2-6-2-4, a compressed hot exhaust gas front delivery pipe 2-6-2-5, a compressed hot exhaust gas rear delivery pipe 2-6-2-6 and a vector airflow distribution system 4; one end of the compressed cold air front delivery pipe 2-6-2-3 is connected to the housing of the RCW engine 2, and ... The compressed cold air is obtained from the output end of the centrifugal assembly disc 2-6, and the other end is connected to the first cold air collecting cover 2-1-5 in the first air intake structure 2-1 of the RCW engine 2, providing the oxygen-containing compressed air required for the air-fuel-rocket closed combustion for the combustion chamber of the RCW engine 2; one end of the compressed cold air rear delivery pipe 2-6-2-4 is connected to the housing of the RCW engine 2, and the compressed cold air is obtained from the output end of the turbine axial flow and centrifugal assembly disc 2-6, and the other end is connected to the second air intake structure 3 of the YRC engine 3 -1 is connected to the second cold air collecting cover 3-1-5 in the YRC engine 3 to provide the oxygen-containing compressed air required for the air-fuel-rocket closed combustion for the combustion chamber of the YRC engine 3; the compressed hot exhaust gas front delivery pipe 2-6-2-5 is connected to the RCW engine casing at one end, and obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disc 2-6, and the other end is connected to the first hot exhaust gas collecting cover 2-1-6 in the first air intake structure 2-1 of the RCW engine 2. The output of the first hot exhaust gas collecting cover 2-1-6 is the output of the RCW engine 2. The combustion chamber wall provides cooling gas; one end of the compressed hot exhaust gas delivery pipe 2-6-2-6 is connected to the RCW engine housing, and obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disk 2-6, and the other end is connected to the second hot exhaust gas collecting cover 3-1-6 of the YRC engine 3 to provide cooling gas for the combustion chamber wall of the YRC engine 3; the vector airflow distribution system 4 obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disk 2-6, and provides vector jet airflow for vertical take-off and landing and aerospace maneuvers of aerospace vehicles;

[0099] See also Figure 3In the RCW engine 2, the first air intake structure 2-1 includes: a central air intake duct 2-1-3, a turbine axial flow air intake duct 2-1-4, a first cold air collecting cover 2-1-5, and a first hot exhaust gas collecting cover 2-1-6. The first cold air collecting cover 2-1-5 is provided with 12 first conveying support channels 2-1-7. The compressed cold air of the first cold air collecting cover 2-1-5 provides the first combustion chamber 2-4 of the RCW engine 2 with fresh compressed air for closed combustion required for the air rocket working condition through the 12 first conveying support channels 2-1-7. The first hot exhaust gas collecting cover 2-1-6 is provided with 6 second conveying support channels 2-1-8. The compressed hot exhaust gas from the cover 2-1-6 provides cooling airflow for the RCW combustion chamber wall, turbine axial flow and centrifugal assembly disc 2-6 through six second delivery support channels 2-1-8; the central air inlet duct 2-1-3 and the turbine axial flow air inlet duct 2-1-4 are provided in the first cold air collecting cover 2-1-5, forming two inner and outer air inlet channels; an air intake truncated cone port 2-1-1 and a conical ram air inlet valve 2-1-2 matching the air intake truncated cone port 2-1-1 are provided at the port of the central air inlet duct 2-1-3, and the axial sliding of the conical ram air inlet valve 2-1-2 is controlled by a hydraulic system to realize opening or closing between the conical ram air inlet valve 2-1-2 and the air intake truncated cone port 2-1-1;

[0100] See also Figure 2 、 3, 6, 7, the turbine axial flow and centrifugal assembly disk 2-6 in the RCW engine 2 includes a front turbofan disk group 2-6-1 and a rear fan-out disk group 2-6-2 installed on the reluctance generator / motor 2-5 and the main shaft, the turbofan disk group 2-6-1 is composed of three-stage turbofan disks and guide disks stacked in sequence to form an axial flow intake compressor; the fan-out disk group 2-6-2 is composed of a first fan-out disk 2-6-2-1 and its guide disk and a second fan-out disk 2-6-2-2 and its guide disk stacked in sequence to form a centrifugal compressor; in the first fan The RCW engine casing at the fan disc 2-6-2-1 is connected to the front conveying pipe 2-6-2-3 for compressed cold air and the rear conveying pipe 2-6-2-4 for compressed cold air; the front conveying pipe 2-6-2-5 for compressed hot exhaust gas and the rear conveying pipe 2-6-2-6 for compressed hot exhaust gas are connected to the RCW engine casing at the second fan disc 2-6-2-2; each stage of the turbofan disc, fan disc and guide disc is provided with three isolation sealing rings 2-6-3-1, 2-6-3-2 and 2-6-3-3 with different diameters in the radial direction. The fan disk, fan-out disk and guide disk are radially divided into four ducts. The four ducts distributed from the center to the circumference are defined as the first duct 2-6-4-1, the second duct 2-6-4-2, the third duct 2-6-4-3 and the fourth duct 2-6-4-4. The first, second and fourth ducts of the turbofan disk group 2-6-1 are equipped with hollow supersonic turbofan blades and guide blades 2-6-5-1, 2-6-5-2 and 2-6-5-4, and the third duct is equipped with power turbine blades and guide blades 2-6-5-3. The 1st and 2nd ducts of the 2nd stage fan-out disk group 2-6-2 are provided with hollow supersonic turbofan blades, the 3rd duct is provided with power turbine blades and guide vanes, and the 4th duct is provided with centrifugal blades and guide vanes 2-6-5-5 and 2-6-5-6; the hollow supersonic turbofan blades, power turbine blades and guide vanes in the 1st, 2nd and 3rd ducts are respectively connected to the two adjacent isolation sealing rings constituting the duct, and the centrifugal blades and guide vanes in the 4th duct are connected to the isolation sealing ring 2-6-3-3 at one end, and connected to the outer ring of the fan-out disk at the other end.

[0101] The turbofan disc assembly 2-6-1 is a hybrid disc consisting of three-stage axial compressor hollow supersonic turbofan blades, hollow power turbine blades and guide vanes.

[0102] Duct Airflow Distribution: When the turbofan disk diameter is 1.2 meters, the headwind airflow to ducts 1, 2, and 4 is distributed based on the cross-sectional area of ​​the predetermined duct annular diameter difference. Duct 1 receives 18-19% of the total headwind airflow, duct 2 receives 60%, and duct 4 receives 21-22%. (Annular diameter difference is the difference between the diameter of the first isolation seal ring and the turbofan disk diameter; the difference between the diameter of the third isolation seal ring and the diameter of the second isolation seal ring; and so on for the other ducts.)

[0103] The deflector discs are located between each stage of the turbofan discs. The fourth outer ring of the deflector discs is the engine casing, while the deflector disc casing of the fan-off disc is the centrifugal compressor casing. The front of each stage casing houses the first turbofan disc. Each stage casing is secured to the preceding and succeeding stage casings with isolation and sealing rings, secured with screws for easy removal and maintenance.

[0104] Power turbine temperature drop: Turbofan disks are generally 3-stage, and each stage of the power turbine can reduce the gas temperature by about 90-110°C. The same 3-stage turbofan disk can reach 5-7 or even 9 stages or more according to the engine power requirements, so as to output greater power and minimize the exhaust gas temperature.

[0105] The fan-out disk: There are two levels of fan-out disks behind the multi-stage turbofan disk:

[0106] 1. First-stage fan disc 2-6-2-1: This is welded together with the hollow supersonic turbofan blades and guide vanes of the first and second ducts, the power turbine blades and guide vanes of the third duct, and the centrifugal compressor blades and guide vanes of the fourth duct, along with the isolation and sealing rings of each duct. High-pressure cold air from the fourth duct is delivered via forward delivery duct 2-6-2-3 to the central first combustion chamber, providing high-pressure fresh air for closed-loop combustion. The compressed cold air rear delivery duct 2-6-2-4 supplies the high-pressure air required for the air-fuel-rocket operation of the ramjets on both sides.

[0107] 2. Second-stage fan disc 2-6-2-2: The hollow supersonic turbofan blades and guide vanes of the first and second ducts, the power turbine blades and guide vanes of the third duct, and the centrifugal blades and guide vanes of the fourth duct centrifugal compressor are welded to the isolation sealing rings of each duct. The third duct centrifugal compressor provides high-pressure, medium-temperature cooling gas of about 450°C-500°C to the central first and third combustion chambers through the forward delivery pipe 2-6-2-5 and the rear delivery pipe 2-6-2-6.

[0108] The compressed cold air of the first duct enters the first fan-out disk 2-6-2-1 and is then introduced into the second combustion chamber by its guide disk; the compressed cold air of the second duct is mixed with the gas of the third duct and enters the second fan-out disk 2-6-2-2 and is then introduced into the vector airflow distribution air collecting hood by its guide disk; the compressed cold air of the fourth duct enters the first fan-out disk 2-6-2-1 and is then guided out by its guide disk, and all the way through the compressed cold air front delivery pipe 2-6-2-3 to provide the RCW engine combustion chamber with the oxygen-containing compressed air required for the air-fuel-rocket closed combustion. The other route provides the YRC engine's combustion chamber with oxygen-containing compressed air required for the air-fuel-rocket closed combustion through the compressed cold air rear delivery pipe 2-6-2-4; part of the compressed cold air from the 4th duct is mixed with part of the fuel gas from the 3rd duct and then enters the second fan-out disc 2-6-2-2 and is then discharged through its guide disc. One route provides cooling air for the RCW engine combustion chamber wall through the compressed hot exhaust gas front delivery pipe 2-6-2-5, and the other route provides cooling air for the YRC engine combustion chamber wall through the compressed hot exhaust gas rear delivery pipe 2-6-2-6.

[0109] See also Figure 5 The present invention provides an aerospace vehicle power system, wherein the first oil pan assembly 2-3 is composed of an annular oil spray disc 2-3-1-3 and six annular oil quantity control pressure oil pipe channels 2-3-1-2 evenly distributed around the annular oil spray disc 2-3-1-3; one side of the annular oil spray disc 2-3-1-3 is provided with three groups of large, medium and small annular oil spray nozzles 2-3-1, and the center is provided with a silane igniter nozzle 2-3-1-1, and the oil spray nozzle 2-3-1 is a piston structure consisting of an oil control inner heat pipe and an outer sleeve.

[0110] See also Figure 4 The first combustion chamber 2-4 is provided with a ramjet burner 2-4-1, and the combustion chamber 2-4 is provided with six gas delivery pipes 2-4-2 to deliver the burned gas in the ramjet burner 2-4-1 to the annular gas collecting groove 2-4-3, and the gas is guided by the guide vanes in the annular gas collecting groove 2-4-3 and then sprayed to the third duct power turbine blades through the guide nozzle;

[0111] See also Figure 8 In the RCW engine, the second diffuser 2-7, the second oil pan assembly 2-8, and the second combustion chamber 2-9 form an integral whole; an air intake duct 2-7-1, an air intake piston sleeve 2-7-2, and a conical ram air intake valve 2-7-3 are arranged in front of the second diffuser 2-7; the second combustion chamber 2-9 is formed by two combustion tubes to form an inner ring combustion chamber 2-9-1, an outer ring combustion chamber 2-9-2, and an outer ring gaseous pneumatic plug 2-9-3 for stabilizing the fire source, and the tail nozzle bell mouth of the RCW engine tail nozzle 2-10 is connected to form an integral whole;

[0112] See also Figure 9The second oil pan assembly 2-8 consists of an inner oil spray pan 2-8-1, an outer oil spray pan 2-8-2 and 6 annular oil quantity control pressure oil pipe channels 2-8-4 evenly distributed around the outer oil spray pan 2-8-2; the inner oil spray pan 2-8-1 and one side of the outer oil spray pan 2-8-2 are evenly distributed with oil nozzles 2-8-5, and a silane igniter nozzle 2-8-3 is set in the center of the inner oil spray pan 2-8-1. The oil nozzles 2-8-5 on the inner oil spray pan 2-8-1 and the outer oil spray pan 2-8-2 are piston structures composed of an oil control inner heat pipe and an outer sleeve;

[0113] The inner fuel injection disc 2-8-1 supplies fuel to the inner ring combustion chamber 2-9-1, and the outer fuel injection disc 2-8-2 supplies fuel to the outer ring combustion chamber 2-9-2. Once the inner ring combustion chamber 2-9-1 is ignited by the silane igniter nozzle 2-8-3 in the center of the second combustion chamber 2-9, its combustion gas serves as a stable ignition source for the outer ring combustion chamber 2-9-2, forming a gaseous pneumatic plug 2-9-3. By adjusting the amount of fuel in the inner fuel injection disc 2-8-1, the expansion and contraction of the jet stream after the tail nozzle can be adjusted.

[0114] See also Figure 10In the YRC engine 3, the second air intake structure 3-1 includes an air intake truncated cone 3-1-2, a conical ram intake valve 3-1-1 and an outer duct transcendental ram intake 3-1-3, a second cold air collecting cover 3-1-5, and a second hot exhaust gas collecting cover 3-1-6; the second cold air collecting cover 3-1-5 is installed at one end of the YRC engine 3 casing, located at the YRC engine casing portion corresponding to the position of the third diffuser chamber 3-2; the air intake truncated cone 3-1-2 is set at the second cold air collecting cover 3-1-5 On one side, the truncated cone type ram inlet valve 3-1-1 is located in the center of the second cold air collecting cover 3-1-5 and matches the truncated cone inlet port 3-1-2. The axial sliding of the truncated cone type ram inlet valve 3-1-1 is controlled by the hydraulic system to open or close the truncated cone type ram inlet valve 3-1-1 and the truncated cone inlet port 3-1-2. When closed, the YRC engine is in the working condition of an air fuel rocket engine or a liquid oxygen fuel rocket engine. When opened, the YRC engine is in the working condition of a ramjet or a scramjet. The truncated cone inlet port 3-1- 2 is connected to the second air intake duct 1-3 of the air intake system 1 to provide the required ram-oxygen-containing fresh air for the YRC engine ramjet and super ramjet working conditions. The second cold air collecting cover 3-1-5 is provided with 12 third delivery support channels 3-1-7. The compressed cold air entering the second cold air collecting cover 3-1-5 through the delivery pipe 2-6-2-4 is connected to the third diffusion chamber 3-2 through the 12 third delivery support channels 3-1-7, providing air fuel rocket launch for the third combustion chamber 3-4 of the YRC engine 3. Fresh compressed air for closed combustion required for engine operation; a second hot exhaust gas collecting hood 3-1-6 is provided on the YRC engine casing, located at the portion of the YRC engine casing corresponding to the position of the fuel injection plate 3-3-2. Six fourth delivery support channels 3-1-8 are provided in the second hot exhaust gas collecting hood 3-1-6. The compressed hot exhaust gas entering the second hot exhaust gas collecting hood 3-1-6 through the delivery pipe 2-6-2-6 after compression passes through the six fourth delivery support channels 3-1-8 to provide cooling airflow for the YRC combustion chamber wall;

[0115] The third combustion chamber 3-4 in the YRC engine 3 is composed of a rocket combustion tube 3-4-1, a rocket ramjet combustion tube 3-4-2, and a rocket super-ramjet combustion tube 3-4-3, which respectively constitute a rocket combustion chamber, a ramjet combustion chamber, and an outer duct super-ramjet combustion chamber. The three combustion tubes have different diameters. When the YRC engine is in rocket and ramjet operating conditions, the outer duct 3-4-6 serves as an airflow cooling channel for the outer wall of the rocket ramjet combustion tube 3-4-2. When in super-ramjet operating conditions, the outer duct 3-4-6 is a mixing channel for air and fuel. Super-fuel occurs at the tail end 3-4-5 of the outer duct 3-4-6, and the tail end of the super-ramjet combustion tube 3-4-3 is extended and connected to the bell mouth of the second tail nozzle 3-5 to form a whole.

[0116] See also Figure 11 The YRC engine 3, the third fuel injection disc 3-3 includes an integrally arranged rocket fuel injection disc 3-3-1, a ram fuel injection disc 3-3-2 and a super-ram fuel injection disc 3-3-3 arranged at the super-ram fuel intake 3-1-3 of the outer duct. The rocket fuel injection disc 3-3-1 has a smaller diameter and is coaxially embedded in the ram fuel injection disc 3-3-2. The rocket fuel injection disc 3-3-1 and the ram fuel injection disc 3-3-2 are both provided with at least one circle of fuel injection nozzles 3-3-1-1 and 3-3-2-2; a silane injection igniter 3-3-4 is provided at the center of the rocket fuel injection disc 3-3-1; the fuel injection nozzles 3-3-1-1 and 3-3-2-2 are piston structures composed of an oil-control inner heat pipe and an outer sleeve; 6 hydraulic pipes, fuel, and hot exhaust gas support channels 3-3-5 are uniformly distributed around the ram fuel injection disc 3-3-2;

[0117] The rocket fuel injection disc 3-3-1 provides fuel for the rocket combustion chamber, and the ramjet fuel injection disc 3-3-2 provides fuel for the ramjet combustion chamber. When the fuel in the rocket combustion chamber is ignited by the silane jet igniter 3-3-4 located in the center of the rocket fuel injection disc 3-3-1, its combustion gas serves as a stable fire source for the ramjet combustion chamber and forms a gaseous pneumatic plug 3-4-4. By adjusting the amount of fuel in the rocket fuel injection disc 3-3-1, the expansion and contraction of the jet stream after the tail nozzle 3-5 can be adjusted.

[0118] See also Figure 12 The super ramjet fuel injection disc 3-3-3 is provided with at least one circle of fuel injection nozzles 3-3-3-1. The super ramjet fuel injection disc 3-3-3 is arranged along the inner edge of the super ramjet air inlet 3-1-3 of the outer duct. When the speed node of super ramjet startup is reached, a large amount of oxygen-containing ram air and the fuel injected by the annular fuel injection nozzles 3-3-3-1 are fully mixed in the super ramjet outer duct 3-4-6, effectively extending the mixing distance of the fuel and air, thereby greatly improving the full mixing of the ramjet air and the fuel, and burning at the tail end 3-4-5 of the outer duct 3-4-6, thereby improving the efficiency of complete combustion;

[0119] See also Figure 13The vector airflow distribution system 4 includes a vector airflow distribution air collection cover 4-1, which is arranged on the outer shell of the centrifugal compressor at the second fan-out disk 2-6-2-2 of the RCW engine. The compressed cold air of the second duct is mixed with the gas of the third duct and enters the second fan-out disk 2-6-2-2, and is introduced into the vector airflow air collection cover by its guide plate. The hot exhaust gas is introduced into the vector airflow distribution air collection cover 4-1 through the guide plate 4-1-7 overlapping with the second fan-out disk 2-6-2-2. The vector airflow distribution air collection cover 4-1 is provided with an axially fixed guide sleeve throat 4-1-9 and an axially movable guide sleeve throat 4-1-10. The vector airflow distribution air collection cover A vector airflow bend 4-1-1 is provided outside the shell 4-1, and a vector airflow distributor 4-1-2 is provided on the vector airflow bend 4-1-1 to distribute the hot exhaust gas in the vector airflow distribution air collecting hood 4-1 to the front left lower jet vector airflow duct 4-1-4, the rear left lower jet vector airflow duct 4-1-3, the front right lower jet vector airflow duct 4-1-5, and the rear right lower jet vector airflow duct 4-1-6 connected to the outlet end of the vector airflow bend 4-1-1, providing a vector jet airflow source for the three-point lower jet vector shutters 4-8, 4-9, 4-10, the horizontal vector nozzles 4-2, 4-3, 4-4, 4-5, and the vector nozzles 4-6 and 4-7 rotating around the horizontal direction of the left and right delta wings;

[0120] When the axially movable guide sleeve throat 4-1-10 moves forward and is sealed with the second fan-out disc 2-6-2-2 and the outer shell 4-1-11 of the guide disc, and the wave crest of the axially movable guide sleeve throat 4-1-10 and the axially fixed guide sleeve throat 4-1-9 is misaligned, the second duct mixed hot air flow passes through the radial annular and axial curved channels 4-1-12 between the axially fixed guide sleeve throat 4-1-9 and the axially movable guide sleeve throat 4-1-10, and is sprayed toward the outer wall of the second combustion chamber to cool the combustion tube, and then is expanded and ejected through the central trumpet mouth, thereby forming vector thrust for low-altitude and low-speed cruising.

[0121] When the axially movable guide sleeve throat 4-1-10 moves backward and is sealed with the wave crest of the axially fixed guide sleeve throat 4-1-9, the second ducted mixed hot air flow enters the vector air flow distribution air collecting hood 4-1 and is then output through the vector air flow bend 4-1-1, forming a vector jet power source for aircraft take-off and landing, hovering and aerospace maneuvers.

[0122] Computer simulation experimental data of the YRCW aerospace engine of the present invention

[0123] YRCW utilizes the buoyancy of air at altitudes between 0 and 40 km, as well as the mass of oxygen and air involved in the after-expulsion. Therefore, calculations cannot be based solely on the equations for horizontal rocket launches; instead, they must be based on the formulas for action, reaction, and velocity. At altitudes between 30 and 60 km, oxygen levels are insufficient, requiring liquid oxygen supplementation. At altitudes between 60 and 100 km, air levels are negligible, requiring calculations based on the equations for horizontal rocket launches.

[0124] The computer simulation experiment of the present invention adopts the American SimuLnK simulation calculation software, and imports the action force, reaction force and speed formula and the rocket horizontal launch equation, the speed of the YRCW aerospace engine in various working conditions and the total mass data of the YRCW engine and the aircraft, fuel data, and the physical data of the earth's gravity and the atmosphere into the SimuLnK model. Then, the MATLAB algorithm is integrated into the model in SimuLnK, and the inclined velocity obtained by the predetermined elevation angle of the YRCW is used to calculate the horizontal velocity component and the vertical ascent velocity component, and the simulation data results are exported for further analysis.

[0125] Solution 1: Simulation result data:

[0126] Preset conditions: Pitch angle 0-30 degrees; medium YRCW takeoff gross weight 52 tons;

[0127] Air inlet cross-sectional area: 2m 2 ;

[0128] The equivalent of complete combustion of CH compound (kerosene) with air is: air:kerosene = 14.78:1 (equivalent to: oxygen:kerosene = 2.956:1);

[0129]

[0130]

[0131] Solution 1: Evaluation summary of simulation results:

[0132] The YRCW aerospace vehicle weighs 5-6 tons, takes off at 50-52 tons, and carries 11 tons of fuel and 17 tons of liquid oxygen. It consumes 6.4-7 tons of fuel and 7.6-7 tons of liquid oxygen to reach a low orbit at an altitude of 140 km. One orbital orbit consumes 0.5-1 ton of electrical power and 1.5-3 tons of liquid oxygen. The return flight, inverted from -24m to -23m, consumes 1.2 tons of fuel and 3.6 tons of liquid oxygen. Total fuel consumption is 8.5-9 tons, and liquid oxygen is 14-15 tons. The remaining mass is 28 tons, with 2 tons of fuel and 2 tons of liquid oxygen remaining. The estimated payload is 15-16 tons.

[0133] Analysis and optimization: In Scheme 1, although the YRCW utilizes the air buoyancy and oxygen at an altitude of 0 to 40 km and the mass of air participating in the after-spray, the oxygen in the thin air at an altitude of 30 to 60 km is insufficient and liquid oxygen needs to be supplemented. The air buoyancy and oxygen in the air are not optimally utilized. The ramjet engine is suitable for operating conditions of 1.5M to 5M, and the scramjet engine is suitable for operating conditions of 5M to 12M, and can also reach a limit of 25M, but it must operate at an altitude of 30 to 40km. Only at high speed can it inhale enough air, and the thin air can avoid thermal barriers.

[0134] Solution 2: Simulation result data:

[0135] Preset conditions: Pitch angle 0-30 degrees; medium YRCW takeoff gross weight 52 tons;

[0136] Air inlet cross-sectional area: 2m 2 ;

[0137] The equivalent of complete combustion of CH compound (kerosene) with air is: air:kerosene = 14.78:1 (equivalent to: oxygen:kerosene = 2.956:1);

[0138]

[0139]

[0140] Scheme 2: Simulation Results Evaluation Summary: Based on flight envelope optimization, the aerospace vehicle weighs 5-6 tons, takes off with a weight of 52 tons, and carries 11 tons of fuel and 11 tons of liquid oxygen. Entering a low-orbit altitude of 110 km consumes 7-8 tons of fuel and 3.5-4 tons of liquid oxygen. One orbital orbit consumes 0.5-1 ton of electrical power and 1.5-3 tons of liquid oxygen. Return flight from -24m to -23m consumes 0.7 tons of fuel and 1.9 tons of liquid oxygen. Total fuel consumption and liquid oxygen consumption are 9 tons, leaving a residual mass of 34 tons, 2 tons of fuel, and 2 tons of liquid oxygen. The estimated payload is 22-24 tons. Therefore, flight envelope optimization requires significant attention.

[0141] The working principle of the present invention is briefly described below:

[0142] The aerospace vehicle power system of the present invention is installed on the aerospace vehicle and operated according to the following steps:

[0143] Step 1: Start the YRCW engine and take off the aircraft - speed 0 Mach - 0.6 Mach

[0144] (1) Horizontal takeoff: When the YRCW engine is started for the first time, the fuel and silane in the first oil pan assembly 2-3 are pre-pressurized, the airflow distribution plates 1-1 on both sides of the air intake are adjusted to slope, covering the outer air intake 1-3, and all the airflow from the air intake 1 on both sides enters the inner air intake 1-2 of the central engine RCW. The conical ramjet inlet valve 2-1-2 of the central engine RCW and the conical ramjet inlet valve 3-1-1 of the left and right rear engines YRC are hydraulically moved forward, closing the RCW intake truncated cone port 2-1-1 and the YRC intake truncated cone port 3-1-2. The reluctance generator / motor 2-5 is in motor mode, driving the turbine axial flow and centrifugal assembly disk 2-6 to rotate, and the RCW turbine axial flow inlet 2-1 -4 is compressed by the axial flow intake of the three-stage turbine and then centrifugally compressed by the first fan-out disk 2-6-2-1 to obtain the compressed fresh cold air, which is then delivered to the first cold air collecting hood 2-1-5 through the compressed cold air front delivery pipe 2-6-2-3. At this time, the pressure of the compressed fresh cold air has reached more than 2068 kPa. The compressed cold air is then evenly introduced into the first diffusion chamber 2-2 for diffusion through the 12 first delivery support channels 2-1-7 provided in the first cold air collecting hood 2-1-5, and then enters the first combustion chamber 2-4. The inner annular oil pan of the first oil pan assembly 2-3 is sprayed with oil, and the silane igniter 2-3-1-1 ignites and burns. The RCW first combustion chamber then forms a closed combustion air-fuel-rocket working condition.

[0145] The gas generated by combustion is delivered to the annular gas collecting tank 2-4-3 through the six gas delivery pipes 2-4-2 provided in the combustion chamber 2-4. The gas is guided by the guide vanes in the annular gas collecting tank 2-4-3 and then ejected through the guide nozzle to impact the power turbine, driving the turbofan disk and the fan-out disk to rotate at high speed. At this time, the front and rear reluctance generators / motors 2-5 are in the generator state.

[0146] The closed combustion air-fuel-rocket working condition combustion gas of the first combustion chamber 2-4 impacts the turbofan disk and the fan-out disk. The hot exhaust gas after high-speed rotation and compression enters the second fan-out disk 2-6-2-2 through the third duct, and is discharged through its guide disk and enters the first hot exhaust gas collecting hood 2-1-6 and the second hot exhaust gas collecting hood 3-1-6 respectively, providing cooling gas for the combustion chamber walls of the RCW engine and the YRC engine. At this time, the medium-temperature hot exhaust gas of the second duct enters the vector airflow distribution hood 4-1 through the guide disk of the second fan-out disk 2-6-2-2. When the axially movable guide throat moves forward and is sealed with the outer shell 4-1-11 of the guide disk of the second fan-out disk 2-6-2-2, the fixed guide throat 4-1-9 and the axially movable guide throat 4-1-10 are misaligned, and the mixed hot air flow of the second duct is ejected into the channel 4-1-12 between the outer wall of the second combustion chamber 2-9 and the RCW engine casing through the radial annular and axial curved channels between the axially fixed guide throat 4-1-9 and the axially movable guide throat 4-1-10 to cool the outer wall of the combustion chamber, and then is ejected through the central trumpet expansion pressure, thereby driving the aerospace vehicle to glide.

[0147] When the axially movable guide throat 4-1-10 moves backward to align with the wave crest of the axially fixed guide throat 4-1-9, the mixed hot air from the second duct enters the vector airflow distribution hood 4-1, passes through the left and right vector airflow elbows 4-1-1, and enters the vector airflow micro-distributor 4-1-2. The vector injection system then controls the three-point downward spray vector shutters 4-8, 4-9, and 4-10 to spray air toward the ground, causing the aerospace vehicle to rise. At the same time, the second combustion chamber 2-9 ignites, propelling the aerospace vehicle to rapidly accelerate away from the runway until the aerospace vehicle reaches a flight speed of Mach 0.6. At this point, the second combustion chamber 2-9 is in a closed combustion air rocket mode.

[0148] (2) Vertical takeoff: The RCW engine is started (same as in scheme (1)). After the first combustion chamber 2-4 is ignited and burned, the axial movable guide sleeve throat 4-1-10 in the engine vector airflow system 4 is moved backward to be sealed with the wave crest of the axial fixed guide sleeve throat 4-1-9. The second ducted mixed hot air flow enters the vector airflow distribution air collecting cover 4-1, and enters the vector airflow micro-distributor 4-1-2 through the left and right vector airflow bends 4-1-1. Fine-tuning is performed to maintain the balance of the aircraft for vertical takeoff. The fine-tuned vector airflow is transported to the front three-point down-spray vector airflow through the front and left and right pipes. Vanes 4-8, 4-9, and 4-10 spray. The outer ring of the first oil pan 2-3 in the RCW central engine sprays oil, increasing the flow rate of the downjet vectored hot air. When the aircraft vertically rises to 2-5 meters, the second combustion chamber injector disc 2-8 sprays and ignites, accelerating to above Mach 0.2. The rear downjet vectored air shutters gradually close. The axially movable guide sleeve throat 4-1-10 moves forward, forming a partially sealed seal with the guide disc housing 4-1-11 of the second fan disc 2-6-2-2. A very small portion of the second ducted mixed hot air flows into the front downjet vectored air shutters 4-8 for upward-lift vectored injection. Left and right ducts deliver the air to the left and right vectored flaps for preparatory injection. The majority of the air flows through the radial annular and axially curved channels 4-1-12 between the axially fixed guide sleeve throat 4-1-9 and the axially movable guide sleeve throat 4-1-10, ejecting it into the outer wall channel of the second combustion chamber. It then expands and exits through the central bell mouth. The aircraft enters subsonic flight.

[0149] Step 2: The flight speed of the aerospace vehicle is from 0.6 Mach to 0.8 Mach.

[0150] When the flight speed of the aerospace vehicle reaches Mach 0.6 or above, the airflow distribution plates 1-1 on both sides of the air inlet are adjusted to gradually leave the ramp peak to distribute air for the ramjet working condition of the first and second combustion chambers of the central engine RCW. The conical ramjet inlet valve 2-1-2 is controlled by the hydraulic system to slide axially so that the conical ramjet inlet valve 2-1-2 retreats to expose the intake truncated cone 2-1-1. The ramjet air of the central air inlet 2-1-3 enters the first combustion chamber 2-4 and the second combustion chamber 2-9 of the RCW engine and directly burns to produce work. The first combustion chamber and the second combustion chamber of the central RCW engine are converted to ramjet working condition.

[0151] At this time, the high-speed oncoming ram fresh air of the turbine axial flow inlet 2-1-4 continues to be compressed by the turbofan disk and fan-out disk and then output according to the step 1 plan.

[0152] At this point, the conical ramjet valve 3-1-1 and the intake truncated cone 3-1-2 in the YRC engine's second intake structure 3-1 are closed. Fresh air entering from the RCW engine's turbine axial-flow intake duct 2-1-4 is compressed by the turbofan axial-flow compressor and the fan-off centrifugal compressor. It is then delivered to the second cold air hood 3-1-5 via the compressed cold air delivery pipe 2-6-2-4. It is then evenly delivered to the YRC combustion chambers 3-4-1 and 3-4-2 on both sides through the 12 high-pressure fresh cold air support channels 3-1-7. At this point, the first and second stages of the YRC engine's third combustion chambers on both sides enter a closed-cycle air-fuel-rocket mode, propelling the aerospace vehicle to a speed of Mach 0.8.

[0153] Step 3: The flight speed of the aerospace vehicle is from 0.8 Mach to 6 Mach, and the altitude is 10-30 km:

[0154] When the aerospace vehicle reaches a speed of Mach 0.8, the airflow distribution plates 1-1 on both sides of the intake duct are adjusted further away from the ramp peak to distribute air for the first and second stage combustion chambers of the YRC third combustion chambers on both sides. The airflow separation plate 1-1 in the intake system 1 is adjusted to open the second intake duct 1-3. Simultaneously, the conical ramjet inlet valve 3-1-1 and the intake truncated cone 3-1-2 in the second intake structure 3-1 of the YRC engine are controlled to be open. Ramjet fresh air entering from the second intake duct 1-3 directly enters the third diffuser 3-2 for expansion, and then enters the two-stage burners 3-4-1 and 3-4-2 of the third combustion chamber. Fuel injection ignites and combustion occurs, and the YRC engine combustion chamber enters ramjet mode. This propels the aerospace vehicle to a speed of Mach 6.

[0155] Step 4: The flight speed of the aerospace vehicle is from Mach 6 to Mach 16, and the altitude is 30-60 km.

[0156] When the flight speed of the aerospace vehicle reaches Mach 6, the airflow distribution plates 1-1 on both sides of the air intake are adjusted to be closer to the inner side of the ramp peak to distribute air for the YRC outer ducts of the engines on both sides in the super-combustion ramjet mode. The inner sleeve 1-4 of the central air intake piston moves forward to seal against the straightening cone 1-2-1 to prevent the hot gas from leaking forward from the central engine. The central RCW engine controls the conical ramjet intake valve 2-1-2 to hydraulically rush forward to close the intake truncated cone port 2-1-1. The YRC engines on both sides control the conical ramjet intake valve 3-1-2 to hydraulically rush forward to close the intake truncated cone port 3-1-1. The RCW engine uses the compressed fresh cold air from the first cold air hood 2-1-5 to appropriately supplement oxygen to maintain the first combustion chamber 2-4 and the second combustion chamber 2-9 working in the closed combustion air-fuel-rocket-turbine mode. The YRC engine uses the compressed fresh cold air from the second cold air hood 3-1-5 to appropriately supplement oxygen. Maintaining the combustion tubes 3-4-1 and 3-4-2 of the third combustion chamber working in the air-fuel-rocket-turbine working condition of closed combustion, at this time, the hypersonic airflow of the duct 3-4-6 of the ramjet inlet increases greatly, and the ramjet oil pan 3-3-3 arranged on the inner edge of the ramjet inlet 3-1-3 of the duct of the YRC engine begins to spray oil. The duct combustion chamber of the YRC third combustion chamber enters the ramjet working condition, and the ramjet occurs at the junction of the duct 3-4-6 and the tail end 3-4-5, until the flight speed of the aerospace vehicle reaches Mach 16.

[0157] Step 5: The flight speed of the aerospace vehicle is from Mach 16 to Mach 24.

[0158] When the aerospace vehicle reaches Mach 16, the central engine is properly oxygenated and operates in an air-fuel-rocket-turbine mode. The scramjets on both sides stop operating in a scramjets mode, and the ramjets on both sides gradually increase the amount of liquid oxygen, entering a liquid oxygen-rocket mode. This propels the aerospace vehicle to Mach 24 and an altitude of 140-320 kilometers.

[0159] Step 6: After the aerospace vehicle reaches a speed of Mach 24

[0160] The YRCW vehicle can reach various low-Earth orbits at speeds of Mach 24 and altitudes exceeding 100 km. During a sightseeing flight around Earth, the YRCW vehicle enters space cruise mode, with both engines completely deactivated. The central engine operates in an air-fuel-rocket-turbine mode, with only a small amount of fuel injected from the inner ring of the RCW's first injector disc 2-3. The turbine engine idles to power the generator, and the recycled hot exhaust gas provides vector propulsion for orbital maneuvers. The axially movable guide sleeve throat 4-1-10 moves rearward to align with the wave crest of the axially fixed guide sleeve throat 4-1-9, sealing it. The second duct mixes the hot air flow into the vector flow hood, maintaining the hot exhaust pressure to facilitate space maneuvers, space travel, and space debris avoidance. If the circulating hot exhaust gas pressure in the vector flow hood exceeds the limit during low-orbit cruise, the closed-circuit hot exhaust gas is released through the vector flap nozzles.

[0161] Step 7: Return to space:

[0162] When re-entering the atmosphere, a YRCW aircraft must first change its flight attitude. At an altitude of 100 km, all YRCW aircraft in orbital cruise mode must pivot, with the longitudinal axis of the fuselage forming a 30-degree angle with the trajectory. The YRCW engines on both sides are activated in a liquid oxygen-fuel-rocket mode, decelerating the inverted flight from -24M to -23M. When the speed of the inverted aircraft is less than -23M, the remaining mass of the aircraft is 28 tons, and the aircraft begins to free fall. After a free fall of 111s at an altitude of 60km, the vertical terminal speed can reach 3.3 Mach. The vector airflow jet must be turned on within 110s after the aircraft starts free fall to drive the YRCW aircraft to turn around again. When the aircraft lands at an altitude of 60km, its horizontal speed is still 22 to 23 Mach, the vertical terminal speed is 3.3 Mach, and the slant speed reaches 22.24 to 23.23M. The aircraft is maneuvered into and maintained at an elevation angle of 15 to 35 degrees through vector jet, with the belly of the aircraft acting as a resistance surface to slow down the acceleration of the free fall and the slant speed of 22.24 to 23.23 Mach. At an altitude of 40 to 30 km, the YRCW aircraft activates its ramjet engine to control its horizontal velocity, pitch speed, descent rate, and pitch attitude. This is because the air intakes at altitudes of 40 to 30 km provide the necessary air for the RCW engines. Although the air is thin, the horizontal velocity and air intake volume are high. At an altitude of 30 km, the YRCW aircraft enters a high-angle glide state of approximately 30 degrees, avoiding thermal barriers. After decelerating to below 1 meter, it corrects its course and glides to an airport for a horizontal landing. Upon landing, the tricycle's downward-facing vectored airflow shutters deploy, shortening the glide distance. During return-to-space landing, the YRCW aircraft can choose a vertical landing method in challenging weather conditions or at locations without runways.

[0163] The above description is merely a preferred embodiment of the present invention. The scope of protection of the present invention is not limited to the above embodiment. All technical solutions based on the concept of the present invention are within the scope of protection of the present invention. It should be noted that for those skilled in the art, various improvements and modifications that do not depart from the principles of the present invention should also be considered within the scope of protection of the present invention.

Claims

1. An aerospace vehicle power system, comprising an air intake system (1), a central engine (2), and two rear-mounted engines (3) on both sides, wherein the central engine (2) has a four-mode cyclic conversion function among a ramjet turbine engine operating mode, a ramjet engine operating mode, an air-fuel rocket engine operating mode, and a liquid oxygen-fuel rocket engine operating mode, and is referred to as an RCW engine (2); the two rear-mounted engines (3) on both sides have a four-mode cyclic conversion function among a ramjet engine operating mode, a scramjet engine operating mode, an air-fuel rocket engine operating mode, and a liquid oxygen-fuel rocket engine operating mode, and are referred to as YRC engines (3); characterized in that: A first air intake structure (2-1), a first diffuser (2-2), a first oil pan assembly (2-3), a first combustion chamber (2-4), a reluctance generator / motor (2-5), a turbine axial flow and centrifugal assembly disk (2-6) provided on the main shaft of the reluctance generator / motor, a second diffuser (2-7), a second oil pan assembly (2-8), a second combustion chamber (2-9), and a first tail nozzle (2-10) are sequentially provided along the axial direction of the RCW engine (2); along the Y The RC engine (3) is axially provided with a second air intake structure (3-1), a third diffuser chamber (3-2), a third fuel injection disc (3-3), a third combustion chamber (3-4), and a second tail nozzle (3-5); the first air intake structure (2-1) of the RCW engine (2) is connected to an output end (1-2-2) of the air intake system (1); the second air intake structure (3-1) of the YRC engine (3) is connected to the second air intake duct (1-3) of the air intake system (1); A valve mechanism is provided on the housing of the RCW engine (2) at the output end of the turbine axial flow and centrifugal assembly disc (2-6), the valve mechanism comprising a compressed cold air front delivery pipe (2-6-2-3), a compressed cold air rear delivery pipe (2-6-2-4), a compressed hot exhaust gas front delivery pipe (2-6-2-5), a compressed hot exhaust gas rear delivery pipe (2-6-2-6) and a vector airflow distribution system (4); the compressed cold air front delivery pipe (2-6-2-3) One end is connected to the housing of the RCW engine (2), and compressed cold air is obtained from the output end of the turbine axial flow and centrifugal assembly disk (2-6). The other end is connected to the first cold air collecting cover (2-1-5) in the first air intake structure (2-1) of the RCW engine (2), and oxygen-containing compressed air required for the air-fuel-rocket closed combustion is provided to the combustion chamber of the RCW engine (2); one end of the compressed cold air rear delivery pipe (2-6-2-4) is connected to the housing of the RCW engine (2), and compressed cold air is obtained from the output end of the turbine axial flow and centrifugal assembly disk (2-6). The other end is connected to the second cold air collecting cover (3-1-5) in the second air intake structure (3-1) of the YRC engine (3), and oxygen-containing compressed air required for the air-fuel-rocket closed combustion is provided to the combustion chamber of the YRC engine (3); one end of the compressed hot exhaust gas front delivery pipe (2-6-2-5) is connected to the RCW engine (2), and compressed cold air is obtained from the output end of the turbine axial flow and centrifugal assembly disk (2-6). The W engine casing obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disc (2-6), and the other end is connected to the first hot exhaust gas collecting cover (2-1-6) in the first air intake structure (2-1) of the RCW engine (2), and the output of the first hot exhaust gas collecting cover (2-1-6) provides cooling gas for the combustion chamber wall of the RCW engine (2); one end of the compressed hot exhaust gas delivery pipe (2-6-2-6) is connected to the RCW engine casing, obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disc (2-6), and the other end is connected to the second hot exhaust gas collecting cover (3-1-6) of the YRC engine (3), and provides cooling gas for the combustion chamber wall of the YRC engine (3); the vector airflow distribution system (4) obtains hot exhaust gas from the output end of the turbine axial flow and centrifugal assembly disc (2-6), and provides vector jet airflow for horizontal take-off and landing, vertical take-off and landing, and aerospace maneuvers of the aerospace vehicle.

2. The aerospace vehicle power system according to claim 1, characterized in that: The air intake system (1) is a U-shaped air intake duct, wherein an S-shaped slope is provided in the middle of each of the two air intake channels of the U-shaped air intake duct, and an air flow separation plate (1-1) is provided at the peak of each S-shaped slope; a first air intake duct (1-2) and a second air intake duct (1-3) are provided in parallel starting from the air flow separation plate (1-1); a first air intake duct (1-2) provided on each side of the U-shaped air intake duct extends to the bottom of the U-shaped duct and connects to form an output end (1-2-2) of the air intake system, and the output end (1-2- 2) The central intake duct piston inner sleeve (1-4) is connected to the first intake structure (2-1) of the RCW engine (2), and a rectifying cone (1-2-1) is provided at the U-shaped bottom of the intake system; a second intake duct provided on each side of the U-shaped intake duct serves as another output end (1-3) of the intake system and is respectively connected to the second intake structure (3-1) of the YRC engine (3) on both sides; the cross-sectional area of ​​the central intake duct piston inner sleeve (1-4) is 1 / 5-1 / 8 of the cross-sectional area of ​​the intake duct.

3. The aerospace vehicle power system according to claim 1, characterized in that: In the RCW engine (2), the first air intake structure (2-1) includes: a central air intake duct (2-1-3), a turbine axial flow air intake duct (2-1-4), a first cold air collecting hood (2-1-5), and a first hot exhaust gas collecting hood (2-1-6). The first cold air collecting hood (2-1-5) is provided with a plurality of conveying support channels (2-1-7). The compressed cold air of the first cold air collecting hood (2-1-5) provides the first combustion chamber (2-4) of the RCW engine (2) with fresh compressed air for closed combustion required for the air-fuel-rocket engine working condition through the plurality of first conveying support channels (2-1-7). The first hot exhaust gas collecting hood (2-1-6) is provided with a plurality of second conveying support channels (2-1-8). The compressed hot exhaust gas from the air collecting hood (2-1-6) provides cooling airflow for the RCW combustion chamber wall, the turbine axial flow and centrifugal assembly discs (2-6) through multiple conveying support channels (2-1-8); the central air inlet duct (2-1-3) and the turbine axial flow air inlet duct (2-1-4) are provided in the first cold air collecting hood (2-1-5), forming two inner and outer air inlet channels; an air intake truncated cone port (2-1-1) and a conical ramjet air inlet valve (2-1-2) matched with the air intake truncated cone port (2-1-1) are provided at the port of the central air inlet duct (2-1-3); the axial sliding of the conical ramjet air inlet valve (2-1-2) is controlled by a hydraulic system to realize opening or closing between the conical ramjet air inlet valve (2-1-2) and the conical ramjet air inlet port (2-1-1).

4. The aerospace vehicle power system according to claim 3, characterized in that: The number of the first conveying support channels (2-1-7) is 4-18, and the number of the second conveying support channels (2-1-8) is 3-12.

5. The aerospace vehicle power system according to claim 1, characterized in that: The turbine axial flow and centrifugal assembly disc (2-6) in the RCW engine (2) includes a front turbofan disc group (2-6-1) and a rear fan-out disc group (2-6-2) installed on a reluctance generator / motor (2-5) and a main shaft, wherein the front turbofan disc group (2-6-1) is composed of a plurality of turbofan discs and guide discs stacked in sequence to form an axial flow air intake compressor; the rear fan-out disc group (2-6-2) is composed of a first fan-out disc (2-6-2-1) and its guide disc and a second fan-out disc (2-6-2-2) and its guide disc stacked in sequence to form a centrifugal compressor; The RCW engine casing at the first fan disc (2-6-2-1) is connected to a front delivery pipe for compressed cold air (2-6-2-3) and a rear delivery pipe for compressed cold air (2-6-2-4); the RCW engine casing at the second fan disc (2-6-2-2) is connected to a front delivery pipe for compressed hot exhaust gas (2-6-2-5) and a rear delivery pipe for compressed hot exhaust gas (2-6-2-6); each stage of the turbofan disc, fan disc, and guide disc is provided with three isolation sealing rings (2-6-3-1, 2-6-3-2, 2-6-3-3) of different diameters in the radial direction. ), the turbofan disk, fan-out disk and guide disk are divided into four ducts along the radial direction. The four ducts distributed from the center to the circumference are defined as the 1st duct (2-6-4-1), the 2nd duct (2-6-4-2), the 3rd duct (2-6-4-3) and the 4th duct (2-6-4-4). The 1st, 2nd and 4th ducts of the front turbofan disk group (2-6-1) are equipped with hollow supersonic turbofan blades and guide blades (2-6-5-1, 2-6-5-2, 2-6-5-4), and the 3rd duct is equipped with power turbine blades and guide blades (2-6- 5-3); The 1st and 2nd ducts of the rear-mounted 2-stage fan-out disk group (2-6-2) are provided with hollow supersonic turbofan blades, the 3rd duct is provided with power turbine blades and guide vanes, and the 4th duct is provided with centrifugal blades and guide vanes (2-6-5-5, 2-6-5-6); the hollow supersonic turbofan blades, power turbine blades and guide vanes in the 1st, 2nd and 3rd ducts are respectively connected to the two adjacent isolation sealing rings constituting the duct, and the centrifugal blades and guide vanes in the 4th duct are connected to the isolation sealing ring (2-6-3-3) at one end and connected to the outer ring of the fan-out disk at the other end.

6. The aerospace vehicle power system according to claim 5, characterized in that: The number of the turbofan discs and guide discs is 3-12, forming an axial flow air intake compressor.

7. The aerospace vehicle power system according to claim 5, characterized in that: The compressed cold air of the first duct enters the first fan-out disk (2-6-2-1) and is then introduced into the second combustion chamber by its guide disk; the compressed cold air of the second duct is mixed with the combustion gas of the third duct and enters the second fan-out disk (2-6-2-2) and is then introduced into the vector gas distribution flow collecting hood by its guide disk; the compressed cold air of the fourth duct enters the first fan-out disk (2-6-2-1) and is then guided out by its guide disk, and all the way through the compressed cold air front delivery pipe (2-6-2-3) to provide the RCW engine combustion chamber with oxygen-containing compressed air required for air-fuel-rocket closed combustion. The other route provides the oxygen-containing compressed air required for the air-fuel-rocket closed combustion of the YRC engine's combustion chamber through the compressed cold air rear delivery pipe (2-6-2-4); part of the compressed cold air from the 4th duct is mixed with part of the combustion gas from the 3rd duct and enters the second fan-out disk (2-6-2-2) and then is discharged through its guide disk. One route provides cooling gas for the RCW engine combustion chamber wall through the compressed hot exhaust gas front delivery pipe (2-6-2-5), and the other route provides cooling gas for the YRC engine's combustion chamber wall through the compressed hot exhaust gas rear delivery pipe (2-6-2-6).

8. The aerospace vehicle power system according to claim 1, characterized in that: The first oil pan assembly (2-3) is composed of an annular oil spray pan (2-3-1-3) and a plurality of annular oil quantity control pressure oil pipe channels (2-3-1-2) uniformly distributed around the annular oil spray pan (2-3-1-3); three groups of large, medium and small oil spray nozzles (2-3-1) are provided on one side of the annular oil spray pan (2-3-1-3), and a silane igniter nozzle (2-3-1-1) is provided at the center; the oil spray nozzle (2-3-1) is a piston structure composed of an oil control inner heat pipe and an outer sleeve.

9. The aerospace vehicle power system according to claim 8, characterized in that: The number of annular oil quantity control pressure oil pipe channels (2-3-1-2) is 3-12.

10. The aerospace vehicle power system according to claim 1, characterized in that: A ramjet burner (2-4-1) is provided in the first combustion chamber (2-4). Six gas delivery pipes (2-4-2) are provided in the first combustion chamber (2-4) to deliver the combusted gas in the ramjet burner (2-4-1) to an annular gas collecting groove (2-4-3). The gas is guided by guide vanes in the annular gas collecting groove (2-4-3) and then sprayed toward the third duct power turbine blades through a guide nozzle.

11. The aerospace vehicle power system according to claim 1, characterized in that: In the RCW engine, the second diffuser (2-7), the second oil pan assembly (2-8), and the second combustion chamber (2-9) form a whole; an air intake duct (2-7-1), an air intake duct piston sleeve (2-7-2), and a conical ram air intake valve (2-7-3) are arranged in front of the second diffuser (2-7); the second combustion chamber (2-9) is formed by two combustion tubes to form an inner ring combustion chamber (2-9-1), an outer ring combustion chamber (2-9-2), a gaseous pneumatic plug (2-9-3) for stabilizing the fire source of the outer ring, and a tail nozzle bell mouth of the RCW engine tail nozzle (2-10), which are connected to form a whole.

12. The aerospace vehicle power system according to claim 1, characterized in that: The second oil pan assembly (2-8) is composed of an inner oil spray pan (2-8-1), an outer oil spray pan (2-8-2), and six annular oil quantity control pressure oil pipe channels (2-8-4) uniformly distributed around the outer oil spray pan (2-8-2); oil spray nozzles (2-8-5) are uniformly distributed on one side of the inner oil spray pan (2-8-1) and the outer oil spray pan (2-8-2), a silane igniter nozzle (2-8-3) is provided at the center of the inner oil spray pan (2-8-1), and the oil spray nozzles (2-8-5) on the inner oil spray pan (2-8-1) and the outer oil spray pan (2-8-2) are piston structures composed of an oil control inner heat pipe and an outer sleeve; The inner fuel injection disc (2-8-1) provides fuel for the inner ring combustion chamber (2-9-1), and the outer fuel injection disc (2-8-2) provides fuel for the outer ring combustion chamber (2-9-2). The fuel in the inner ring combustion chamber (2-9-1) is ignited by the silane igniter nozzle (2-8-3) in the center of the second combustion chamber (2-9), and its combustion gas serves as a stable fire source for the outer ring combustion chamber (2-9-2) to form a gaseous pneumatic plug (2-9-3). The expansion and contraction of the jet flow after the tail nozzle can be adjusted by adjusting the amount of fuel in the inner fuel injection disc (2-8-1).

13. The aerospace vehicle power system according to claim 1, characterized in that: In the YRC engine (3), the second air intake structure (3-1) includes an air intake truncated cone (3-1-2), a conical ramjet air intake valve (3-1-1), an external duct scramjet air intake (3-1-3), a second cold air collecting hood (3-1-5), and a second hot exhaust gas collecting hood (3-1-6); the second cold air collecting hood (3-1-5) is installed at one end of the YRC engine (3) casing, and is located at a portion of the YRC engine casing corresponding to the position of the third diffuser chamber (3-2); the air intake truncated cone (3-1-2) is arranged at the second cold air collecting hood On one side of (3-1-5), the conical ramjet inlet valve (3-1-1) is located in the center of the second cold air collecting cover (3-1-5) and matches the air intake truncated cone port (3-1-2). The axial sliding of the conical ramjet inlet valve (3-1-1) is controlled by the hydraulic system to open or close the conical ramjet inlet valve (3-1-1) and the air intake truncated cone port (3-1-2). When closed, the YRC engine is in the working condition of an air fuel rocket engine or a liquid oxygen fuel rocket engine. When opened, the YRC engine is in the working condition of a ramjet or a scramjet. The air intake truncated cone port (3 -1-2) is connected to the second air inlet duct (1-3) of the air inlet system (1) to provide the oxygen-containing gas required for the YRC engine ramjet and scramjet working conditions. The second cold air collecting hood (3-1-5) is provided with a plurality of third delivery support channels (3-1-7). The compressed cold air entering the second cold air collecting hood (3-1-5) through the delivery pipe (2-6-2-4) is connected to the third diffusion chamber (3-2) through the plurality of third delivery support channels (3-1-7) to provide air, fuel, and fire for the third combustion chamber (3-4) of the YRC engine (3). Fresh compressed air for closed combustion required for the working condition of the Arrow engine; a second compressed hot exhaust gas collecting hood (3-1-6), which is arranged on the YRC engine casing and is located at the YRC engine casing portion corresponding to the position of the injection plate (3-3-2); a plurality of fourth delivery support channels (3-1-8) are provided in the second hot exhaust gas collecting hood (3-1-6); the compressed hot exhaust gas entering the second hot exhaust gas collecting hood (3-1-6) through the delivery pipe (2-6-2-6) after the compressed hot exhaust gas provides cooling airflow for the YRC combustion chamber wall through the plurality of fourth delivery support channels (3-1-8).

14. The aerospace vehicle power system according to claim 13, characterized in that: The number of the third conveying support channels (3-1-7) is 4-18, and the number of the fourth conveying support channels (3-1-8) is 3-12.

15. The aerospace vehicle power system according to claim 13, characterized in that: The third combustion chamber (3-4) in the YRC engine (3) is composed of a rocket combustion tube (3-4-1), a rocket ramjet combustion tube (3-4-2), and a rocket scramjet combustion tube (3-4-3), which respectively constitute a rocket combustion chamber, a ramjet combustion chamber, and an outer duct scramjet combustion chamber, and the three combustion tubes have different diameters; when the YRC engine is in rocket and ramjet working conditions, the outer duct (3-4-6) serves as an air flow cooling channel for the rocket ramjet combustion tube (3-4-2); when in scramjet working conditions, the outer duct (3-4-6) is a mixing channel for air and fuel, and scramjet occurs at the tail end (3-4-5) of the outer duct (3-4-6), and the tail end of the scramjet combustion tube (3-4-3) is extended and connected to the bell mouth of the second tail nozzle (3-5) to form a whole.

16. The aerospace vehicle power system according to claim 1, characterized in that: The YRC engine (3), the third fuel injection disc (3-3) comprises an integrally arranged rocket fuel injection disc (3-3-1), a ramjet fuel injection disc (3-3-2), and a scramjet fuel injection disc (3-3-3) arranged at the scramjet air inlet (3-1-3). The rocket fuel injection disc (3-3-1) has a relatively small diameter and is coaxially embedded in the stamping fuel injection disc (3-3-2). Both the rocket fuel injection disc (3-3-1) and the stamping fuel injection disc (3-3-2) are provided with at least one circle of fuel injection nozzles (3-3-1-1, 3-3-2-2). A silane injection igniter (3-3-4) is provided at the center of the rocket fuel injection disc (3-3-1). The fuel injection nozzles (3-3-1-1, 3-3-2-2) are piston structures composed of an oil control inner heat pipe and an outer sleeve. Six hydraulic pipes, fuel, and hot exhaust gas support channels (3-3-5) are uniformly distributed around the stamping fuel injection disc (3-3-2). The rocket fuel injection disc (3-3-1) provides fuel for the rocket combustion chamber, and the ramjet fuel injection disc (3-3-2) provides fuel for the ramjet combustion chamber. When the fuel in the rocket combustion chamber is ignited by the silane injection igniter (3-3-4) located at the center of the rocket fuel injection disc (3-3-1), the combustion gas serves as a stable fire source for the ramjet combustion chamber to form a gaseous pneumatic plug (3-4-4). By adjusting the amount of fuel in the rocket fuel injection disc (3-3-1), the expansion and contraction of the rear jet flow of the tail nozzle (3-5) can be adjusted. The scramjet fuel injection disc (3-3-3) is provided with at least one ring of fuel injection nozzles (3-3-3-1), which are arranged on the inner edge of the scramjet air inlet (3-1-3) of the ring outer duct. When the speed node for scramjet startup is reached, a large amount of oxygen-containing ramjet air and fuel injected by the ring fuel injection nozzles (3-3-3-1) are fully mixed in the scramjet outer duct (3-4-6) and are completely burned at the tail end (3-4-5) of the outer duct (3-4-6).

17. The aerospace vehicle power system according to claim 5, characterized in that: The vector airflow distribution system (4) comprises a vector airflow distribution air collecting hood (4-1), wherein the vector airflow distribution air collecting hood (4-1) is arranged on the outer shell of the centrifugal compressor at the second fan-out disk (2-6-2-2) of the RCW engine. After the compressed cold air of the second duct is mixed with the combustion gas of the third duct and enters the second fan-out disk (2-6-2-2), it is introduced into the vector airflow distribution air collecting hood (4-1) by the guide disk (4-1-7) thereof. The vector airflow distribution air collecting hood (4-1) is provided with an axial fixed guide sleeve throat (4-1-9) and an axial movable guide sleeve throat (4-1-10). The vector airflow distribution air collecting hood (4-1) is provided with a vector airflow bend pipe (4-1) outside the shell. -1), a vector airflow distributor (4-1-2) is provided on the vector airflow bend (4-1-1), which distributes the hot exhaust gas in the vector airflow distribution air hood (4-1) to the front left lower jet vector airflow duct (4-1-4), the rear left lower jet vector airflow duct (4-1-3), the front right lower jet vector airflow duct (4-1-5), and the rear right lower jet vector airflow duct (4-1-6) connected to the outlet end of the vector airflow bend (4-1-1), providing a vector airflow source for the three-point lower jet vector shutter (4-8, 4-9, 4-10), the horizontal vector nozzle (4-2, 4-3, 4-4, 4-5), and the vector nozzle (4-6, 4-7) rotating around the horizontal direction of the left and right delta wings.

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