Coreless non-circular rotating detonation scramjet engine using micro PDE

Abstract

A coreless non-circular rotating detonation scramjet engine, provided by the present invention, comprises: a combustion-type air heater; a shape transition nozzle connected to the combustion-type air heater; an isolator connected to the shape transition nozzle; a combustor connected to the isolator and comprising a cavity having an obround cross section and an expansion nozzle; multiple fuel injection ports formed in a ring shape along the outer circumferential surfaces of the isolator, the cavity, and the expansion nozzle; and an igniter for igniting a compound located in the cavity.

Description

Coreless non-circular rotary detonation scramjet engine using micro PDE

[0001] The present invention relates to a scramjet engine operating on the basis of detonation, and more particularly, to a coreless non-circular rotating detonation scramjet engine.

[0002]

[0003] The scramjet engine, a supersonic combustion ramjet engine, is an advanced propulsion technology that enables hypersonic flight. In the aerospace industry, this engine can be applied to hypersonic vehicles and missiles capable of speeds exceeding Mach 5. It is attracting attention as a hypersonic vehicle engine concept that overcomes the flight speed limitations of conventional jet and ramjet engines.

[0004] A scramjet is an advanced version of the ramjet engine, designed to operate at supersonic speeds. Ramjets utilize naturally compressed air, utilizing airflow, but combustion requires slowing the air down to subsonic speeds. In contrast, scramjets burn supersonic air without slowing it down, allowing them to operate efficiently even at high speeds exceeding Mach 5.

[0005] Looking at the operating process, 1) Air intake: When a supersonic aircraft inhales air, the air naturally compresses. 2) Combustion: The compressed air and fuel are mixed, and combustion occurs under high temperature and high pressure. 3) Propulsion generation: The combusted gas is ejected at high speed, providing powerful propulsion.

[0006] The key features of the scramjet engine are as follows: 1) Structural simplicity: Scramjets do not require a rotating compressor or turbine, and are designed as a single, long tube with a cross-section such as a circle or square. 2) Utilization of atmospheric oxygen: Since scramjets use atmospheric oxygen as an oxidizer, they do not require a separate oxygen tank, making them lighter and more efficient than rockets. 3) Capable of hypersonic flight: They can achieve speeds of Mach 5 to 15 or more, giving them significant potential for military and commercial applications.

[0007] The advantages of the scramjet engine are as follows: 1) Fuel efficiency: Because it utilizes atmospheric oxygen, it eliminates the need for oxygen tanks, making it more economical than rockets. 2) Hypersonic flight: It can travel at speeds exceeding Mach 5, offering military advantages in evading air defense systems. 3) Simplified structure: It compresses air using shock waves without a turbine, resulting in a simpler structure.

[0008] The challenges facing scramjet engines are as follows: 1) High-temperature issues: Special materials and cooling technologies are needed to withstand the high temperatures encountered during ultra-high-speed flight. 2) Combustion stability: The technical challenges of maintaining stable fuel-air mixing and combustion at supersonic speeds must be addressed. 3) Initial startup: Scramjets cannot achieve supersonic speeds on their own, requiring auxiliary engines such as turbofans or rockets.

[0009] Scramjets have potential applications in a wide range of fields, including hypersonic aircraft, cruise missiles, and satellite launch vehicles. In particular, hybrid applications utilizing scramjets in the atmosphere and conventional rocket engines in space offer significant cost savings and expanded flight envelopes. Scramjet technology is currently being actively researched and developed in major countries, including the United States and Japan. South Korea has also developed core technology for a Mach 6.7-class supersonic ramjet engine.

[0010] Scramjet engines have made rapid progress over the past decade, with numerous demonstrations and flight tests reported, but there are still limitations that need to be addressed.

[0011] A key feature of scramjet combustion engines is that they inhale air at hypersonic flight speeds, which limits the time required for fuel / air mixing and combustion due to the supersonic flow velocity inside the combustor. Although the flow residence time inside the combustor is very short, on the order of 1 ms, the time required for fuel / air mixing and combustion is much longer, resulting in lower combustion efficiency compared to other propulsion engines. Currently, the supersonic combustion of scramjet engines relies on turbulent mixing due to the combustion performance of a longitudinally elongated diffusion flame, and the combustion efficiency of the engine remains at around 0.7. Therefore, due to the low combustion efficiency, a sufficiently long combustor is required, which increases the overall system weight and makes systematic integration difficult.

[0012] Furthermore, as the combustor length required for supersonic combustion increases, operational instability within scramjet engines can arise due to unsteady combustion conditions within the supersonic flow field, pressure gradients arising from various causes, and boundary layer interactions. Furthermore, the elongated combustor geometry can also lead to longitudinal combustion instability.

[0013] In order to solve the above-described problem, the purpose of the present invention is to provide a shape design and technique that can effectively control combustion (location and amount of heat generation) within a scramjet engine, improve combustion efficiency by compensating for combustion instability without disturbing the supersonic flow field.

[0014]

[0015] In order to achieve the above-mentioned object, the coreless non-circular rotary detonation scramjet engine of the present invention proposes a method for controlling the flame position and heat amount to increase combustion efficiency and prevent and control combustion instability, and proposes a method for generating and maintaining rotary detonation by rotating a detonation wave, which is a combustion method combining a shock wave and combustion, in a cavity, which is a combustion region for flame stabilization (flameholding) in a supersonic combustor, to obtain a short combustion region and high combustion efficiency.

[0016] An embodiment of the present invention for implementing this is a coreless non-circular rotating detonation scramjet engine experimental device, which may include a combustion air heater, a shape-shifting nozzle connected to the combustion air heater, an isolation portion connected to the shape-shifting nozzle, a combustor connected to the isolation portion and including a cavity having an obround cross-section and an expansion nozzle, a plurality of fuel inlets formed in a ring shape along the outer periphery of the isolation portion, the cavity, and the expansion nozzle, and an igniter for igniting a compound positioned in the cavity. At the same time, the cross-sectional shapes of the isolation portion and the supersonic combustor are not necessarily limited to an obround shape, and may have various shapes such as an ellipse, a rounded-square (a square with rounded corners), a triangle, and a polygon.

[0017] Additionally, the isolation section, the cavity and the expansion nozzle can be separated from each other so that the relative positions, expansion angles and positions of the fuel inlets can be changed, and each can be formed in an upper and lower coupled modular manner.

[0018] Furthermore, a small pulse detonation engine (micro PDE) can be used as the igniter.

[0019]

[0020] According to the present invention, the combustion efficiency of a scramjet engine can be dramatically increased by using rotary detonation.

[0021] In addition, by generating a swirling flow through rotational detonation in a supersonic flow field within a combustor, it is possible to achieve the effects of improving combustion performance of a swirl burner in a subsonic combustor such as a gas turbine swirl burner, and reducing internal drag, heat load, and engine weight due to shortened flame length.

[0022] Furthermore, structurally, there is only a cavity that maintains rotational detonation throughout the supersonic combustor, with no separate core. In a scramjet engine, the presence of a core exposed to supersonic flow complicates the structural design due to extremely high thermal loads, and the weight of the core increases the overall weight. However, the present invention overcomes these issues by enabling the production of various design shapes.

[0023]

[0024] Figure 1 is a conceptual diagram showing one embodiment of the present invention.

[0025] Figure 2 shows the operating range of a rotary detonation scramjet at the flight Mach number.

[0026] Figure 3 is a conceptual diagram showing the operation method of one embodiment of the present invention.

[0027] Figure 4 is a drawing showing the change in expansion angle and cavity position.

[0028] Figure 5 is a conceptual diagram of an aircraft to which a non-circular rotary detonation scramjet engine can be applied.

[0029]

[0030] The advantages and features of the present invention, as well as the techniques for achieving them, will become clearer with reference to the embodiments described below, along with the accompanying drawings. However, the present invention is not limited to the embodiments presented and can be implemented in various forms. These embodiments are provided to ensure a complete description of the present invention and to fully convey the contents of the invention to those skilled in the art.

[0031] Meanwhile, the terms used in this specification are for the purpose of describing embodiments and are not intended to limit the present invention. In this specification, singular forms also include plural forms unless specifically stated otherwise. As used in the specification, the terms "comprise" and "comprising" do not exclude the presence or addition of one or more other components, steps, operations, and / or elements with the mentioned components, steps, operations, and / or elements.

[0032] Additionally, like reference numerals appearing throughout each drawing designate like elements, and detailed descriptions of well-known features and techniques may be omitted to clarify the description of the described embodiments of the present invention.

[0033] Figure 1 is a conceptual diagram showing one embodiment of the present invention.

[0034] A coreless non-circular rotary detonation scramjet engine (100) according to the present invention may include a combustion air heater (110), a shape-shifting nozzle (120) connected to the combustion air heater, an isolation part (130) connected to the shape-shifting nozzle, a combustor (140) including a cavity (141) connected to the isolation part and having an over-round cross-section and an expansion nozzle (142), a plurality of fuel inlets (150) formed in a ring shape along the outer circumference of the isolation part (130), the cavity (141), and the expansion nozzle (142), and an igniter (160) that ignites a compound positioned in the cavity.

[0035] In addition, the isolation portion (130), cavity (141) and expansion nozzle (142) can be separated from each other so that the relative positions, expansion angles and positions of the fuel inlet (150) can be changed, and each can be formed in an upper and lower combined modular manner.

[0036] Furthermore, a micro PDE can be used as the igniter (160).

[0037] The cross-section of the combustor (140) is configured in an obround shape, and quartz (170) is installed on the sides of the isolation section (130) and the combustor (140). Through the quartz (170), direct photography of the inside of the combustion chamber (140), Schilleren photography, and flow visualization techniques can be applied, and the combustion mode and flow field can be identified. In addition to the obround shape, the cross-section shapes of the isolation section (130) and the combustor (140) can have various shapes such as an ellipse, a rounded-square (a square with rounded corners), a triangle, and a polygon.

[0038] The micro-PDE is used as an igniter (160) of the RD-Scramjet engine, and directly forms detonation inside the RD-Scramjet engine combustor (140) by imposing detonation. In addition, the micro-PDE is capable of repeated operation, so it is a device that can re-ignite even if combustion is extinguished.

[0039] The upper and lower parts of the module are interchangeable, allowing for free change of expansion angle, fuel inlet (150) location, cavity (141) location, and micro-PDE location. The micro-PDE can also be changed into various shapes. In other words, various isolation (130) and supersonic combustor (140) shapes can be configured to suit the characteristics of hypersonic aircraft.

[0040] Figures 2, 3, and 4 are conceptual diagrams showing the operational range of a rotary detonation scramjet at a flight Mach number, the operating method of one embodiment of the present invention, and drawings showing the expansion angle and the change in the position of the cavity, respectively. The description will be made with reference to the drawings.

[0041] Improving the combustion efficiency of a scramjet engine not only increases engine performance but also allows for a shorter supersonic combustor (140), facilitating integration into aircraft systems. Furthermore, if combustion instability in the supersonic combustor (140) develops, it can directly lead to the unstarting or even flameout of the hypersonic aircraft. These issues are directly related to safety and require countermeasures and solutions.

[0042] Accordingly, the present invention proposes a method for controlling flame position and heat quantity to increase combustion efficiency of a scramjet engine and prevent and control combustion instability, wherein a detonation wave, which is a combustion method combining shock waves and combustion, is generated and maintained by rotating (rotating detonation) in a cavity (141), which is a combustion area for flame stabilization (flame holding) in a supersonic combustor (160), thereby obtaining a short combustion area and high combustion efficiency.

[0043] The present invention provides a RD-Scramjet engine test device for ground test research as well as a method for operating a scramjet engine (RD-Scramjet) through rotary detonation.

[0044] The rotary detonation scramjet (RD-Scramjet) engine of the present invention uses a micro-pulse detonation engine (micro-PDE) as an igniter capable of stable / repeatable re-ignition, and includes a method of arranging a plurality of fuel inlets (150) in an isolation section (130) and a supersonic combustor (140) for controlling a combustion position and heat amount, and a combustor (140) shape and a cavity (141) shape for maintaining rotary detonation in the supersonic combustor (140). The combustor shape (140) and the cavity shape (141) include all of the combustor (140) expansion angle, the combustor (140) cross-section, the position of the cavity (141), and the size of the cavity (141).

[0045] The micro-PDE used as the igniter (160) in the present invention is capable of both single operation and repeated operation at several tens of Hz, and is composed of a fuel-oxidizer supply unit, an ignition device (spark plug), and a long detonation combustion tube. When the micro-PDE operates, it ignites the fuel-oxidizer through the ignition device, generates a detonation transition (Deflagration to Detonation, DDT) within the long combustion tube, and transmits the detonation wave to the cavity within the supersonic combustor. Since the micro-PDE can be operated at a cycle rate of several to several tens of Hz according to the user's intention to propagate the detonation wave, the rotating detonation can be continuously maintained through rapid re-ignition of the combustor using the micro-PDE in the event of possible rotational detonation extinction, loss, or collapse (detonation failure).

[0046] Typically, a cavity (141) located within a supersonic combustor (140) of a scramjet engine is typically used as a flameholder. However, in the present invention, the cavity (141) space is utilized as a rotating detonation operating region. The cavity (141) is not configured as an annular combustor with a core like a typical RDE (Rotating Detonation Engine). This allows for a reduction in the internal drag, thermal load, and engine weight of the combustor (140). In addition, although the RDE is typically designed to be circular, the region through which detonation propagates does not necessarily have to be circular and there are no significant restrictions on its cross-sectional shape. Therefore, the cavity (141), which is a space through which rotating detonation propagates, is configured as a non-circular cross-section. In the present proposal, the cross-section of the combustor (140) is presented as an obround shape as an example, but the non-circular cross-section is not limited to the obround shape. That is, when a detonation wave is transmitted from the micro-PDE to the cavity (141) space, the detonation rotates within the cavity (141) with an obround cross-section, thereby continuing combustion.

[0047] Through this, the performance can be improved in a narrow combustor (140) area in a scramjet engine by rotary detonation, and the flame length and combustor (140) length can also be reduced. In addition to improving combustion performance, it is possible to achieve the effects of reducing internal drag, heat load, and engine weight. The rotary detonation engine generates a supersonic swirl flow in a supersonic flow field by the induced velocity behind the shock wave, and has the effects of improving combustion performance, shortening the flame length, reducing internal drag, heat load, and engine weight, which are present in a subsonic combustor such as a gas turbine swirl burner, as well as the swirl burner.

[0048] Referring to Figure 2, it can be seen that while a typical scramjet engine has an extremely limited operational flight Mach number depending on its design purpose and shape, the RD-Scramjet enables a wide flight range operation that goes beyond the existing limitations through combustion using rotary detonation.

[0049] Referring to Fig. 3, the concept of both the ground test device and the RD-Scramjet engine can be explained. Conventional scramjet engines struggle to achieve complete combustion at the exit of the supersonic combustor (140) due to their low combustion efficiency. However, the RD-Scramjet engine achieves combustion through high-temperature, high-pressure detonation in the cavity, thereby significantly enhancing combustion performance at that location. The increased combustion performance / efficiency allows for a shortened combustor length, thereby reducing combustor weight. Simultaneously, the reduced combustor length reduces the likelihood or risk of longitudinal combustion instability. Each fuel inlet (150) controls the equivalence ratio and combustion mode within the combustor (140) by adjusting the fuel injection flow rate.

[0050] Referring to Fig. 4 (wherein 111 to 113 denote fuel inlets, 120 denotes a cavity, and 130 denotes an expansion nozzle), the schematic positional relationship between the isolation section and the supersonic combustion chamber can be seen. At (111) to (113) shown in the drawing, fuel inlets are arranged in a ring shape along the cross-section of the obround, and the combustion mode is controlled by adjusting the fuel injection pressure and flow rate at each position. Through modularization, each fuel inlet can be position-adjusted. In addition, the combustion mode can be controlled by adjusting the fuel flow rate by position of the fuel inlet according to the user's intention. By controlling the fuel flow rate of the fuel inlet by position, the amount of heat supplied and the location of heat generation can be controlled, thereby enabling combustion mode control, and the flight Mach number (flight region) of the aircraft can be widely expanded. Hypersonic engines using scramjet aircraft are still being researched with a focus on cruise conditions due to unexpected engine burnout, shock wave-boundary layer interaction within the isolation-combustor, and undesirable ram mode transition. In the present invention, by controlling the combustion mode by adjusting the injector arrangement, fuel injection position, and pressure described above, combustion conditions and combustion mode can be controlled according to the acceleration, deceleration, turn, ascent, and descent of the aircraft, thereby helping the flight performance of the aircraft. The cavity is a region where rotational detonation is directly maintained. The cavity interacts with the supersonic flow field to create a recirculation region within the cavity, which assists in mixing the fuel and air. The mixture region formed within the cavity is ignited by the micro-PDE and combustion proceeds by rotational detonation.

[0051] For the expansion section of a supersonic combustor, the upper and lower expansion angles can be adjusted according to the flight Mach number and altitude. Furthermore, the position of the cavity and fuel injector in the supersonic combustor can be adjusted.

[0052] The present invention enables increased combustion efficiency in hypersonic scramjet engines. This allows for the implementation of various complex acoustic excitations tailored to the aircraft's flight conditions (suction flow characteristics).

[0053] Figure 5 is a schematic diagram of an aircraft with a coreless non-circular rotary detonation scramjet engine. This schematic diagram includes schematic diagrams of the isolation section and combustor of the non-circular rotary detonation scramjet engine, and a cross-sectional schematic diagram of the non-circular rotary detonation scramjet engine. Previously, research focused on scramjet engines with a rectangular cross-section air intake, isolation section, supersonic combustor, and nozzle, such as the X-43 or X-51 Waverider, to enable two-dimensional prediction and facilitate academic approaches. However, such a rectangular combustor design is not suitable for the design of aircraft that must carry actual passengers, cargo, or payloads. The rotary detonation scramjet engine of the present invention is free from the constraints of a rectangular cross-section combustor, and allows the cross-sections of the air intake, isolation section, and supersonic combustor to be designed according to the design of an actual hypersonic aircraft fuselage. At the same time, the combustor length can be reduced, thereby achieving the effect of reducing engine weight.

[0054] Conventional scramjet engines, due to their low combustion efficiency, are designed with long supersonic combustors to ensure sufficient combustion. This long structure can also lead to longitudinal combustion instability in the long flow path. Compared to existing technologies, the present invention maintains rotating detonation in a cavity located within the supersonic combustor, significantly increasing combustion efficiency through explosive combustion (detonative combustion). Furthermore, the shorter combustor length significantly contributes to engine weight reduction and offers advantages in system integration. The shortened combustor length is also expected to suppress longitudinal combustion instability. Furthermore, unlike conventional RDE-based propulsion systems with a central body, the RDE design eliminates drag loss and thermal load issues by not interfering with supersonic flow. Hypersonic vehicles are significantly affected by combustor drag, but the removal of the central body allows for relative freedom from this issue.

[0055] The RD-scramjet engine proposed in this study is not limited to circular or rectangular cross-sections, offering flexibility in system integration. The ignition method using micro-PDEs generates repetitive detonation at a rapid rate, thereby addressing flame loss and failure to maintain rotating detonation due to combustion instability. Furthermore, it can be utilized as a combustion control tool in engines.

Claims

1. Combustion air heater; A shape-changing nozzle connected to the above combustion air heater; An isolation section connected to the above-mentioned shape transfer nozzle; A combustor comprising a cavity having an over-round cross-section and an expanding nozzle connected to the above isolation section; A plurality of fuel injection ports formed in a ring shape along the outer surface of the above isolation section, cavity, and expansion nozzle; and An igniter for igniting a compound located in the above cavity; Coreless non-circular rotary detonation scramjet engine including 2. In paragraph 1, A coreless non-circular rotary detonation scramjet engine characterized in that the above isolation section, cavity and expansion nozzle can be separated from each other so that the relative positions, expansion angles and fuel inlet positions can be changed, and each is formed in an upper and lower coupled modular manner.

3. In Article 1, The above igniter is characterized by being a micro PDE and is a coreless non-circular rotary detonation scramjet engine operating on the basis of rotary detonation operating in a cavity.

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