Film cooling structure of the combustion chamber of a gel scramjet engine

Abstract

The present invention discloses a film cooling structure for the combustion chamber of a gel scramjet engine, comprising an upper film cooling air bleed channel and a lower film cooling air bleed channel arranged on an outer cover and an inner casing. The upper film cooling air bleed channel and the lower film cooling air bleed channel each consist of a tapered contraction section, a development section, and an incident section connected in sequence. The air inlet of the tapered contraction section is located behind the suction port, and the tapered contraction section is used to achieve supersonic airflow pressurization. The development section is parallel to the main flow direction of the combustion chamber and is used to achieve full development of the flow before incident and heat exchange with the solid wall of the engine casing. The injection position of the incident section is located downstream of the fuel injection port to cool the inner wall of the combustion chamber. The present invention can concentrate cooling on the inner wall section that is most severely heated.

Description

Technical Field

[0001] The present invention belongs to the field of thermal protection of hypersonic aircraft, in particular to an air film cooling structure for a combustion chamber of a gel scramjet engine. Background Art

[0002] Compared to subsonic ramjets, scramjets operate at higher flight Mach numbers (Ma > 5) and have higher total inflow temperatures. The internal flow field between the combustion chamber inlet and outlet experiences a sudden temperature increase of several hundred to several thousand Kelvin. The internal supersonic combustion creates a greater thermal load on the scramjets. Thermal protection technologies applicable to scramjets fall into two categories: active and passive. Zheng Kuan's "Application and Influencing Factors of Film Cooling in Scramjets" (Harbin Institute of Technology, 2016) provides an overview of the application scenarios and characteristics of thermal protection technologies for scramjets. Passive / semi-passive thermal protection technologies utilize a special high-temperature-resistant insulation layer to absorb heat through surface radiation or ablation. Active thermal protection technologies are categorized as transpiration cooling, regenerative cooling, and film cooling. Transpiration cooling, due to the presence of porous materials, reduces structural strength and has significant limitations. Regenerative cooling uses fuel as a coolant, passing it through a regenerative cooling channel between the inner and outer walls of the combustion chamber, absorbing some of the heat from the inner wall before being injected into the combustion chamber. Film cooling involves injecting a low-temperature coolant stream through narrow slits and holes into the hotter main stream. The main stream adheres to the wall, forming a film of air, which protects the wall by preventing direct contact between the main stream and the wall. Film cooling reduces frictional resistance on the wall, and the resulting film, when involved in combustion, generates additional thrust.

[0003] In Jin Y, Xu X, Yang Q, et al. Combustion Behavior of Hydrocarbon / BoronGel-Fueled Scramjet [J]. AIAA Journal, 2022: 1-10, experimental measurements of the pressure and heat flux distribution within the combustion chamber of a gel-fueled scramjet engine were conducted. The experimental results demonstrate that gel-fueled scramjet engines have high thermal protection requirements. Due to the poor fluidity and complex rheological properties of gel fuel, it easily causes blockage when flowing through fine-structured regenerative cooling channels and small-scale film cooling channels. This makes conventional regenerative cooling and liquid film cooling schemes using fuel as a coolant unsuitable for gel-fueled scramjet engines. While passive ablative thermal protection technology offers a certain degree of reliability, its protection duration is limited due to the consumption of ablative material. Excessively thick ablative layers increase the engine's structural mass and flight cost, thereby limiting its range. Excessive consumption of the ablative layer in the combustion chamber casing causes structural changes that affect the flow field within the engine. Summary of the Invention

[0004] The present invention aims to provide a film cooling structure for the combustion chamber of a gel scramjet engine. Firstly, based on the characteristics of the engine's external flow field, an air inlet and air duct are selected and designed to pressurize incoming air as a cooling air source. Secondly, based on the heat flux density distribution characteristics of the combustion chamber after fuel injection, the air film incidence position and air film cooling configuration are selected and designed to cool the areas on the inner wall that are most severely heated.

[0005] The technical solutions for achieving the purpose of the present invention are:

[0006] A gel scramjet engine combustion chamber air film cooling structure, comprising an upper air film cooling air bleed channel and a lower air film cooling air bleed channel respectively arranged on the outer cover and the inner shell;

[0007] The upper film cooling air inlet channel and the lower film cooling air inlet channel are both composed of a tapered contraction section, a development section and an incident section connected in sequence;

[0008] The air inlet of the conical contraction section is located at the rear side of the air inlet outlet or the rear side of the suction port, and the conical contraction section is used to achieve supersonic airflow pressurization;

[0009] The development section is parallel to the main flow direction of the combustion chamber and is used to achieve full development of the flow before injection and heat exchange with the solid wall of the engine casing;

[0010] The injection position of the incident section is located downstream of the fuel injection port so as to face the inner wall of the combustion chamber.

[0011] Compared with the prior art, the present invention has the following significant advantages:

[0012] (1) In the design of the cooling configuration, a rear step air film cooling structure is used, which has a more obvious cooling effect than the inclined hole air film. The upper and lower air film cooling structures are axially symmetrical and have the same size, so that the shock waves and expansion waves generated by the upper and lower injection nozzles are reflected symmetrically on the combustion chamber wall, reducing the interference of complex wave systems.

[0013] (2) In the design of the air bleed ports, the upper and lower walls are aligned horizontally, which also ensures symmetrical reflection of the generated shock waves on the wall, simplifying the internal flow field. While meeting the air bleed pressure requirements, the air bleed position leaves enough length for the flow development section and fluid-solid heat exchange section in the air bleed channel. The position is also located at a certain distance from the fuel injection port to prevent the impact of a sudden temperature increase in the flow field after fuel ignition.

[0014] (3) In the design of the injection nozzle, measurement experiments have shown that a heat flux density peak often occurs near the engine fuel injection position, and the cooling efficiency of the film cooling decreases as it goes downstream. Therefore, the film injection position is set after the fuel injection nozzle to concentrate cooling on the inner wall section that is most severely heated.

[0015] (4) In the design of the film cooling channel, first, the air is pressurized by compressing the supersonic airflow to meet the pressure requirement of the combustion chamber. Secondly, the flow development before the film injection is stabilized, and heat is effectively exchanged with the shell wall to reduce the temperature rise caused by the compression section to increase the film cooling efficiency.

[0016] (5) In terms of the evaluation method of cooling effect, two methods are provided: isothermal wall film cooling efficiency and adiabatic wall cooling efficiency. From the perspectives of temperature wall and heat flux conduction, the heat transfer analysis of the film cooling model under the two boundary conditions is carried out. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 Schematic diagram of film cooling structure.

[0018] Figure 2 Schematic diagram of the film cooling scheme applied to the gel scramjet engine.

[0019] Figure 3 Schematic diagram of the Hyshot 2 scramjet engine configuration and two-dimensional computational model.

[0020] Figure 4 This is the pressure cloud diagram of the numerical simulation of the hyshot 2 two-dimensional flow field without film cooling (no fuel injection experimental condition).

[0021] Figure 5 Geometric diagram of the film cooling structure designed for the Hyshot 2 engine.

[0022] Figure 6 This is the two-dimensional flow field pressure cloud map after the film cooling configuration is designed.

[0023] Figure 7 This is the temperature cloud diagram of the front section of the combustion chamber before and after the design of the film cooling structure. DETAILED DESCRIPTION

[0024] The present invention will be further described below with reference to the accompanying drawings and specific embodiments.

[0025] like Figure 1 As shown in FIG. 1 , a film cooling structure for a combustion chamber of a gel scramjet engine of the present invention mainly comprises an air inlet (position) 1, cooling channel sections 2, 3, 4, 5, and an injection port (position) 6. The film cooling scheme for applying the structure to a gel scramjet engine is schematically shown in FIG. Figure 2As shown in the figure, the film cooling structure's air inlet position 1 is determined based on the internal flow field pressure distribution of the target engine, ensuring smooth air injection and ejection. The injection port position 6 is determined based on the internal flow field heat distribution of the engine, cooling the most heated areas of the combustion chamber's inner wall. For engine combustion chambers with large aspect ratios, where the upper and lower inner walls occupy the majority of the heated area, the film cooling structure designed in this invention is a two-dimensional model to address thermal protection issues.

[0026] like Figure 1 As shown in the figure, the film cooling channels inside the outer cover and on the inner shell are axially symmetrical in structure. Except for the width of the air inlet, the geometric dimensions of the upper and lower channels are consistent. Figure 2 As shown, some engines have an intake port between the intake duct and the combustion chamber inlet. This port is used to extract the boundary layer at the end of the intake duct to control its thickness, while also discharging the oblique shock wave generated by the leading edge of the engine cover so that it does not enter the combustion chamber. A weaker shock wave is generated at the trailing edge of the intake port, and the intake port also has a pressure relief effect. The combined effect of these two factors results in inconsistent pressures on the upper and lower walls of the combustion chamber inlet section. Therefore, different upper and lower intake port widths W1 and W2 are designed to adjust the compression ratio of the contraction section, maintaining approximately equal pressures at the exit sections of the upper and lower film channels and balancing the ejected airflow. When film cooling is used on an engine, W1 and W2 are set to 4 to 9 mm, with W2 being 0.2 to 0.6 mm wider than W1.

[0027] The film cooling configuration designed in this invention is a rear-step film cooling configuration commonly used in slit-injection film cooling of supersonic incoming flow. By providing a cooling gas injection channel within the slit of the rear step, the cooling gas is injected parallel to the main supersonic flow. The geometric parameters of the film cooling configuration are as follows: the rear-step slit height is S, and the lip thickness is T. The film cooling configuration is axially symmetrical on the upper and lower combustion chamber walls, with the injection position designed to be downstream of the fuel injection port. When used in an engine, S = 0.5-2.5 mm, and T = 0.25-1.5 mm.

[0028] The air film cooling air duct designed by the present invention consists of three sections, and the sections transition in a tangential form in terms of geometric structure. Figure 1As shown, the first channel 2 is a tapered convergence section, used to achieve supersonic flow boosting. Its inlet is located behind the inlet outlet or the suction port. The second channel 3 is a constant-width development section parallel to the combustion chamber's main flow direction. Its distance H from the combustion chamber wall ensures the structural strength of the combustion chamber wall after the film channel is processed. The development section length L ensures full development of the flow before injection and heat exchange with the engine casing wall. The third channels 4 and 5 are constant-width injection sections. For the designed rear-step injection configuration, the rear-step slit height S and lip thickness T are fixed. This section adjusts the film channel outlet to a position consistent with the rear-step injection configuration. The width of this section is equal to the starting width of the injected film. In practical applications, L = 20-30 mm, H = 10-15 mm, and the distance X3 from the film injection outlet to the fuel injection port in the x-direction is 30-40 mm.

[0029] The index for evaluating the film cooling effect in the present invention is the film cooling efficiency (effective temperature ratio of film cooling). This dimensionless ratio is usually expressed in two ways: adiabatic wall film cooling efficiency η aw and isothermal wall cooling efficiency η iso , corresponding to the two wall thermal boundary conditions of adiabatic and isothermal respectively. The calculation is as follows:

[0030]

[0031]

[0032] Where: Under the condition of adiabatic wall, T r∞ is the recovery temperature of the high temperature mainstream after the fuel injection position, T rc is the recovery temperature of the cooling flow ejected from the injection port 6, T aw is the measured adiabatic wall temperature on the downstream wall of film cooling. Under isothermal wall conditions, q0 is the total heat flux on the wall, q f is the heat reaching the wall after film cooling, part of the total heat transfer q0 (q0-q f ) is carried downstream by the air film.

[0033] Example 1

[0034] Due to the limited publicly available flight and ground test data for gel fuel scramjet aircraft, this paper considers the application scenarios of gel fuel and takes the HyShot 2, a hypersonic aircraft successfully tested in Australia in 2002, as the engineering background. The film cooling configuration and air bleed channels are designed on its scramjet engine to simulate real flight test conditions and evaluate the film cooling effect and the impact of engine structure changes on the internal flow field.

[0035] Hyshot 2 scramjet configuration and its two-dimensional computational model Figure 3As shown, the engine's structure is nearly integrated, with two engines symmetrically combined due to the thrust requirements of actual flight. The Hyshot 2 features an intake port between the inlet and combustion chamber. The combustion chamber has an aspect ratio of approximately 8:1, with the combined top and bottom surfaces accounting for nearly 80% of the total heated area. The fuel injection port is located on the combustion chamber floor. Therefore, the engine is simplified into a two-dimensional model, ignoring cooling on the less heated sides and focusing only on the top and bottom surfaces. This allows the problem to be analyzed using a two-dimensional film cooling model.

[0036] According to Figure 4 The pressure distribution characteristics of the engine two-dimensional model flow field (wind tunnel test condition without fuel injection, incoming flow Mach number Ma=7.4) are shown in FIG. The geometric structure of the film cooling channel and configuration designed by the present invention applied to the Hyshot 2 engine is shown in FIG. Figure 5 The detailed geometric dimensions are as follows: the length of the upper wall air bleed port from the leading edge of the outer cover is X1 = 50-80 mm, and the length of the lower wall air bleed port from the rear of the suction port is X4 = 5-10 mm. The widths of the upper and lower bleed air compression section entrances, W1 and W2, are 4-9 mm, with W2 being 0.2-0.6 mm wider than W1. The middle equal-diameter section is L = 20-30 mm long and has a height of H = 10-15 mm from the inner wall. The injection slit width of the rear-step film cooling configuration is S = 0.5-2.5 mm, and the injection lip height is T = 0.25-1.5 mm. The distance from the film injection outlet to the fuel injection port in the x-direction is X3 = 30-40 mm (the distance from the fuel injection port to the lower wall of the combustion chamber in the real Hyshot 2 engine model is X2 = 57.5 mm).

[0037] Under the same experimental conditions without fuel injection, the flow field pressure cloud diagram of the two-dimensional model after numerical simulation of the designed film cooling structure is as follows: Figure 6 As shown in the figure. The design of the contraction section enables the air film channel to achieve the expected pressure increase, without flow stagnation in the channel, and the cooling gas can be smoothly injected into the combustion chamber to be cooled. The shock wave in the combustion chamber flow field and the expansion wave formed by the upper and lower injection nozzles are reflected symmetrically on the wall, and the pressure distribution on the upper and lower walls is roughly balanced. The temperature distribution cloud diagram near the air film cooling configuration is shown in the figure. Figure 7 Figure (a) shows the temperature distribution without film cooling, and Figure (b) shows the temperature distribution after the film cooling structure is designed. Without fuel injection and without causing a sharp temperature rise, the film cooling configuration of the present invention can achieve a significant cooling effect. The temperature near the wall after film injection is reduced by 400-600K compared to the upstream uncooled section.

[0038] When gel fuel (or regular kerosene) is injected and ignited, the combustion chamber's pressure and temperature rise dramatically. This requires increasing the widths W1 and W2 of the air inlet to increase the boost level in the compression section of the channel. Under these real-world engine operating conditions, the combustion chamber walls are exposed to the even hotter heat flux from combustion, increasing their temperature to 2500K to 3000K. At this point, the film cooling structure designed in this invention significantly reduces the temperature near the walls, achieving even higher cooling efficiency.

Claims

1. A gel scramjet engine combustion chamber air film cooling structure, characterized in that: An upper film cooling air bleed channel and a lower film cooling air bleed channel are respectively arranged on the outer cover and the inner shell; The upper film cooling air inlet channel and the lower film cooling air inlet channel are both composed of a tapered contraction section, a development section and an incident section connected in sequence; The air inlet of the conical contraction section is located at the rear side of the air inlet outlet or the rear side of the suction port, and the conical contraction section is used to achieve supersonic airflow pressurization; The development section is parallel to the main flow direction of the combustion chamber and is used to achieve full development of the flow before injection and heat exchange with the solid wall of the engine casing; The injection position of the incident section is located downstream of the fuel injection port so as to face the inner wall of the combustion chamber.

2. The gel scramjet engine combustion chamber air film cooling structure according to claim 1, characterized in that: The width of the air inlet of the lower film cooling air inlet channel is greater than the width of the air inlet of the upper film cooling air inlet channel.

3. The film cooling structure of the gel scramjet engine combustion chamber according to claim 1, characterized in that: The compression ratio of the contraction section is adjusted by the width of the upper and lower air inlets.

4. The gel scramjet engine combustion chamber air film cooling structure according to claim 1, characterized in that: The width of the upper and lower air inlets is 4 to 9 mm, and the lower air inlet is 0.2 to 0.6 mm wider than the upper air inlet.

5. The gel scramjet engine combustion chamber air film cooling structure according to claim 1, characterized in that: The incident section includes a rear step injection configuration, a rear step slit height S=0.5-2.5 mm, and a lip thickness T=0.25-1.5 mm.

6. The gel scramjet engine combustion chamber air film cooling structure according to claim 1, characterized in that: The distance from the air film injection outlet to the fuel injection outlet is X3 = 30-40 mm.

7. The gel scramjet engine combustion chamber air film cooling structure according to claim 1, characterized in that: The length of the equal diameter section is L = 20 to 30 mm, and the height from the inner wall is H = 10 to 15 mm.

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