A method for reducing transonic drag in a parallel TBCC power ramjet channel using low-temperature combustion

Abstract

A method for reducing transonic drag in a ramjet combustion chamber of a parallel TBCC power ramjet channel using low-temperature combustion belongs to the field of combined propulsion system design for hypersonic vehicles. In the ramjet combustion chamber of the parallel TBCC power system, an activated fuel injection structure is arranged, forced low-temperature ignition is carried out using a high-energy spark plug, and relatively stable low-temperature combustion is formed under the action of the existing flame stabilizer in the combustion chamber. Through stable low-temperature combustion, the exhaust velocity at the throat of the downstream tail nozzle of the ramjet combustion chamber under transonic conditions can be significantly increased by using the heat release acceleration effect in the subsonic airflow, so as to achieve the purpose of weakening the transonic flow blockage in the ramjet channel and significantly reducing the resistance of the ramjet channel. Reasonable adjustment and control of the low-temperature combustion position and intensity ensure that the back pressure in the combustion chamber does not significantly change the overflow of the upstream inlet. The present invention can significantly reduce the resistance of the ramjet channel without imposing higher requirements on the parallel TBCC power system.

Description

Technical Field

[0001] The present invention belongs to the fields of fluid mechanics flow control and the design of combined propulsion systems for hypersonic aircraft, and relates to an active flow control technology that can effectively reduce the resistance of the ramjet channel of a parallel TBCC (Turbine-Based Combined Cycle) power system, and is a method for reducing transonic drag in the ramjet channel of a parallel TBCC power system based on low-temperature combustion. Background Art

[0002] With the rapid development of hypersonic technology in recent years, hypersonic aircraft have now become an important direction and strategic high point in the development of future aviation technology. The horizontal takeoff and landing hypersonic aircraft based on air-breathing combined power have the capabilities of horizontal takeoff and landing on a conventional runway, reusability, and hypersonic cruise in the near space (altitude greater than 20 km, speed greater than Ma5.0+). Relying on the advantages of speed and altitude, it can break through the existing air defense system and achieve long-range strategic deterrence and strikes, and is hailed as another "revolution" in the history of world aviation after propeller and jet aircraft.

[0003] The horizontal takeoff and landing hypersonic aircraft has a speed range spanning subsonic, transonic, supersonic, and hypersonic speeds, which requires the aircraft's aerodynamic layout to be designed in a wide speed range and full envelope matching. However, the flight dynamic pressure of the hypersonic aircraft during takeoff, landing, transonic and other low-speed states is more than 10 times different from that in the cruise state. To meet the design requirements of high lift-drag ratio in cruise and low wave drag in supersonic flight, a large sweep and small aspect ratio wing design is generally adopted. To meet the overall loading requirements, the cross-sectional area of the fuselage exceeds 1.5 times that of traditional supersonic aircraft, and the air-breathing power system occupies most of the fuselage space, resulting in an irregular area rule distribution, small optimization space, and a significant increase in volume wave drag at transonic speed, and there are serious high / low-speed lift-drag matching design contradiction problems. In addition, Ma5.0+ hypersonic aircraft generally need to adopt a parallel TBCC power system scheme. At transonic speed, the ramjet engine in the parallel flow path, as the power for the high-speed section, is in a flow congestion state and cannot effectively pass through the flow. There is a strong bow shock wave system similar to blunt body flow around the upstream of the engine inlet, and the pressure behind the wave increases sharply. Research shows that under transonic flight conditions, the internal resistance of the ramjet flow path accounts for more than 20% of the total aircraft resistance, resulting in a sharp increase in the zero-lift resistance of the aircraft and further exacerbating the transonic thrust-drag contradiction.

[0004] The complex internal flow of the air-breathing propulsion system is highly coupled with the external flow of the aircraft, resulting in the traditional aerodynamic drag reduction design and flow control methods of supersonic aircraft facing difficulties. The aerodynamic efficiency and fuel consumption level of hypersonic aircraft at transonic speed have become one of the key bottlenecks restricting the increase of the aircraft's cruise Mach number and range. Therefore, the demand for the introduction of transonic drag reduction technology for the ramjet channel of a parallel TBCC power system in hypersonic aircraft is very clear and urgent. Summary of the Invention

[0005] The present invention proposes a method for reducing drag in the transonic regime of a parallel TBCC power ramjet channel using low-temperature combustion. By forming relatively stable low-temperature combustion in the combustion chamber of the ramjet channel of the parallel TBCC power system, the heat release acceleration effect in the subsonic airflow can be utilized to significantly increase the exhaust velocity of the throat of the tail nozzle downstream of the combustion chamber in the transonic regime, so as to achieve the purpose of weakening the transonic flow blockage in the ramjet channel and thus significantly reducing the drag of the ramjet channel. At the same time, by reasonably adjusting and controlling the position and intensity of the low-temperature combustion, the back pressure in the combustion chamber does not significantly change the overflow of the upstream inlet, and the present invention provides an efficient and controllable active flow control drag reduction technology for the parallel TBCC power system of hypersonic aircraft.

[0006] To achieve the above technical objectives, the following steps are required:

[0007] Step 1: Set an activated fuel injection structure in the combustion chamber of the ramjet channel of the parallel TBCC power system, and inject the activated fuel perpendicularly into the central flow region of the combustion chamber in the direction of the air inflow.

[0008] Furthermore, the activated fuel injection structure is set downstream of the diffuser section of the ramjet channel of the parallel TBCC power system and at the position of the combustion chamber inlet section, so that the activated fuel can form a well-mixed two-phase mixture using the high-speed air flow at the combustion chamber inlet in the transonic regime and the vortex structure induced by the existing flame stabilizer.

[0009] Step 2: Use a high-energy spark plug for forced low-temperature ignition, and form relatively stable low-temperature combustion under the action of the existing flame stabilizer in the combustion chamber.

[0010] Furthermore, the ignition energy of the high-energy spark plug does not exceed 12 mJ.

[0011] Step 3: According to actual needs, the activated fuel can be selected based on the reactivity of the fuel itself, and the setting position and quantity of the activated fuel injection structure can be adjusted. By adjusting and controlling the position and intensity of the low-temperature combustion, as well as the flame stabilization structure, the effective control of the low-temperature combustion intensity and heat release amount can be achieved, avoiding the generation of relatively concentrated heat release in the combustion chamber resulting in excessive inlet back pressure, thus significantly changing the overflow state of the upstream inlet, and further interfering with the external flow in the TBCC power inlet and the aircraft forebody region.

[0012] According to the Rayleigh criterion, the heat release acceleration effect in the subsonic airflow can be used to increase the exhaust velocity of the throat of the tail nozzle downstream of the combustion chamber in the transonic regime, so as to achieve the purpose of weakening the transonic flow blockage in the ramjet channel and thus reducing the drag of the ramjet channel.

[0013] The advantages of the present invention are as follows:

[0014] (1) High drag reduction efficiency. Under typical transonic conditions in a parallel TBCC power ramjet channel, based on the low-temperature combustion heat release of activated fuel, the total drag of the ramjet internal flow channel can be reduced by 20% - 40% under transonic conditions, effectively weakening the transonic thrust-drag contradiction of hypersonic aircraft.

[0015] (2) Low interference with the engine system. In the present invention, the activated fuel is directly laterally injected through the inner wall of the combustion chamber, and the high-speed air flow at the entrance of the combustion chamber and the vortex structure induced by the existing flame stabilizer are used to form a well-mixed two-phase mixture gas, without changing the existing combustion organizational structure of the ramjet channel; at the same time, by reasonably adjusting and controlling the position and intensity of low-temperature combustion, the back pressure formed by low-temperature combustion does not significantly change the over-flow of the upstream inlet, so it does not interfere with the external flow of the TBCC power inlet and the aircraft forebody area. Description of the Drawings

[0016] Figure 1 Schematic diagram of the activated fuel injection structure in a transonic drag reduction method for a parallel TBCC power ramjet channel using low-temperature combustion provided by the present invention;

[0017] Figure 2 Schematic diagram of the temperature rise in the combustion chamber of the ramjet channel when low-temperature combustion is applied under typical working conditions;

[0018] Figure 3 Schematic diagram of an application example of the low-temperature combustion drag reduction method;

[0019] Figure 4 Drag comparison of the parallel TBCC power ramjet channel before and after low-temperature combustion;

[0020] Figure 5 Flow structure at the inlet of the ramjet channel of the parallel TBCC power ramjet channel under the application of low-temperature combustion. Detailed Embodiment

[0021] The present invention proposes a transonic drag reduction method for a parallel TBCC power ramjet channel using low-temperature combustion. By setting an activated fuel injection structure in the combustion chamber of the ramjet channel of the parallel TBCC power system, forced low-temperature ignition is carried out using a high-energy spark plug, and relatively stable low-temperature combustion is formed under the action of the existing flame stabilizer in the combustion chamber. The subsonic airflow in the combustion chamber is greatly increased in the downstream nozzle throat exhaust velocity under the low-temperature heat release acceleration effect, so as to achieve the purpose of weakening the transonic flow blockage of the ramjet channel and significantly reducing the drag of the ramjet channel. At the same time, by reasonably adjusting and controlling the position and intensity of low-temperature combustion, the back pressure in the combustion chamber does not significantly change the over-flow of the upstream inlet. This method can greatly reduce the drag of the ramjet channel without imposing higher requirements on the parallel TBCC power system.

[0022] AsFigure 1 As shown, the activated fuel is injected into the combustion chamber at the sonic level perpendicular to the incoming air flow direction through the injection structure on the side wall of the inlet of the ramjet channel combustion chamber of the parallel TBCC power system, and a well-mixed two-phase mixture is formed by using the shear of the high-speed compressible air flow at the inlet of the combustion chamber under the transonic condition and the vortex structure induced by the existing flame stabilizer. The high-energy spark plug installed on the inner wall of the combustion chamber then ignites, igniting the two-phase mixture and further forming a relatively stable low-temperature combustion under the action of the existing flame stabilizer in the combustion chamber.

[0023] As Figure 2 shows the measurement results of the temperature rise after stable combustion is established in the combustion chamber under typical working conditions. Different from the combustion of traditional ramjets, under the transonic condition, a high-speed compressible low-enthalpy air flow will be formed in the ramjet channel of the parallel TBCC power system. The heat release rate and heat release amount of the activated fuel combustion are significantly different from those in the conventional environment, showing a relatively mild low-temperature combustion state, and the temperature rise in the combustion chamber does not exceed 200 - 300K. According to the Rayleigh criterion, the low-enthalpy subsonic air flow in the combustion chamber will significantly increase the absolute velocity passing through the throat of the tail nozzle under the heat release acceleration effect, thereby achieving the purpose of weakening the transonic flow blockage in the ramjet channel and significantly reducing the resistance of the ramjet channel. In addition, considering that adding a certain amount of heat in the constant cross-section combustion chamber will cause a temperature rise and also an increase in the static pressure in the combustion chamber, which may lead to the blockage and overflow of the air flow in the inlet duct upstream of the combustion chamber. Therefore, it is necessary to reasonably adjust and optimize the fuel's own reactivity, fuel injection position, flame stabilization structure, etc., in order to effectively control the low-temperature combustion intensity and heat release amount, so that the back pressure in the combustion chamber does not significantly change the overflow state of the upstream inlet duct and avoid interfering with the external flow in the TBCC power inlet and the aircraft forebody area.

[0024] A method for reducing transonic drag in the ramjet channel of a parallel TBCC power using low-temperature combustion provided by the present invention includes the following steps:

[0025] Step 1: Design an activated fuel injection structure as Figure 1 shown in the combustion chamber of the ramjet channel of the parallel TBCC power system, so that the fuel is injected into the central flow area of the combustion chamber perpendicular to the incoming air flow direction.

[0026] The selection of the activated fuel injection structure follows the following principles: It is arranged as much as possible downstream of the diffuser section of the ramjet channel of the parallel TBCC power system and at the inlet cross-section position of the combustion chamber. The activated fuel can form a well-mixed two-phase mixture by using the shear of the high-speed compressible air flow at the inlet of the combustion chamber under the transonic condition and the vortex structure induced by the existing flame stabilizer, which is beneficial to the formation of the initial fire kernel and flame propagation in the low-enthalpy environment.

[0027] Step 2: Use a high-energy spark plug for forced low-temperature ignition, and form relatively stable low-temperature combustion under the action of the existing flame stabilizer in the combustion chamber. According to the Rayleigh criterion, the subsonic airflow in the ramjet combustion chamber can significantly increase the exhaust velocity at the throat of the downstream nozzle of the combustion chamber under the supersonic condition after the heat release acceleration effect, thereby weakening the supersonic flow blockage in the ramjet channel and significantly reducing the ramjet channel resistance.

[0028] Step 3: According to the design requirements, optimize and determine the fuel's own reactivity, fuel injection position, flame stabilization structure, etc. The goal is to effectively control the low-temperature combustion intensity and heat release amount, and avoid the relatively concentrated heat release in the combustion chamber from causing too high an inlet duct backpressure, which significantly changes the overflow state of the upstream inlet duct.

[0029] Example: Figure 3 An application example of reducing drag under supersonic conditions by low-temperature combustion in the ramjet combustion chamber of a typical parallel TBCC power system is given. The simulated flight Mach number condition is 0.9 Mach, the total temperature is 287 K, and the total pressure is 1.2 standard atmospheric pressures at sea level; the activated fuel injection conditions are: the injection ports are arranged on the inner wall of the combustion chamber, perpendicular to the oncoming flow direction, at the entrance section of the straight section of the combustion chamber, the temperature is room temperature, and the equivalence ratio is 0.3. In addition, the fuel is a liquid working medium, and the nozzle diameter is 0.4 mm. Figure 2 This is the combustion chamber temperature rise response curve measured by the thermocouple at the combustion chamber outlet when organizing low-temperature combustion under this condition. It can be seen that the combustion has the characteristics of deflagration, and the temperature at the combustion chamber outlet shows a sudden jump when combustion occurs. However, the total temperature rise is about 200 K, which is significantly different from the temperature rise characteristics in the combustion chamber of a conventional ramjet engine and belongs to typical low-temperature combustion. Figure 4 This is the curve of the model resistance varying with time measured by the balance at the same moment. The results show that after the low-temperature combustion occurs, the model resistance shows the same step characteristic as the combustion chamber temperature rise, and a reduction of about 40% occurs. During the experiment, synchronous BOS measurement was carried out at the inlet of the ramjet channel inlet. The results are as Figure 5 shown. It can be seen that the BOS results did not capture obvious changes in the wave system structure at the inlet of the inlet duct, indicating that the combined design methods such as the activated fuel and injection position used in the experiment effectively controlled the low-temperature combustion intensity and heat release amount (temperature rise), so that the backpressure in the combustion chamber did not change the overflow state of the upstream inlet duct, and avoided the interference of the downstream low-temperature combustion on the external flow of the TBCC power inlet and the aircraft forebody area.

Claims

1. A method for reducing supersonic drag in a parallel TBCC power ramjet channel using low-temperature combustion, characterized in that, the steps are as follows: Step 1: Set an activated fuel injection structure in the combustion chamber of the ramjet channel of the parallel TBCC power system, and inject the activated fuel perpendicularly into the central flow region of the combustion chamber in the direction of the air inflow; Step 2: Use a high-energy spark plug for forced low-temperature ignition, and form relatively stable low-temperature combustion under the action of the existing flame stabilizer in the combustion chamber; Step 3: According to actual needs, select the activated fuel based on the reactivity of the fuel itself, adjust the setting position and quantity of the activated fuel injection structure, and control the low-temperature combustion position, intensity, and flame stabilization structure to effectively control the low-temperature combustion intensity and heat release amount, avoid the generation of relatively concentrated heat release in the combustion chamber resulting in excessive inlet backpressure, thereby significantly changing the overflow state of the upstream inlet, and further interfering with the external flow in the TBCC power inlet and aircraft forebody regions.

2. A method for reducing supersonic drag in a parallel TBCC power ramjet channel using low-temperature combustion according to claim 1, characterized in that, the activated fuel injection structure is arranged downstream of the diffuser section and at the inlet cross-section position of the combustion chamber of the ramjet channel of the parallel TBCC power system, so that the activated fuel can form a fully mixed two-phase mixture gas by using the high-speed air flow at the inlet of the combustion chamber under supersonic conditions and the vortex structure induced by the existing flame stabilizer.

3. A method for reducing supersonic drag in a parallel TBCC power ramjet channel using low-temperature combustion according to claim 1 or 2, characterized in that, the ignition energy of the high-energy spark plug does not exceed 12 mJ.

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