Total Pressure Recovery Coefficient Research Articles

Study on the Influence of Working Characteristics of Centripetal Pump Based on VOF/Mixture Model

Aviation fuel contamination can seriously affect aircraft flight safety, and the centripetal pump is the core component of aviation fuel purification equipment. The performance of centripetal pumps is highly demanded for purification equipment. The operating parameters of centripetal pumps significantly affect the internal flow characteristics, which affects the performance of centripetal pumps. However, the flow characteristics of a centripetal pump influenced by the operating parameters have not yet been elaborated upon. In this research, a three-dimensional numerical simulation of the air-fuel two-phase flow field inside a centripetal pump was carried out using the VOF/Mixture model to investigate the effects of three relatively independent physical quantities, namely, fuel flow, outlet fuel discharge pressure, and rotational speed, on the operating characteristics of the centripetal pump. The flow law inside the flow channel of a centripetal pump was analyzed based on a rotating fluid pressure model, the free liquid surface radius of air-fuel, and the total pressure recovery coefficient. It was found that centripetal pumps have a steady working state and an unsteady working state. In a steady working state, the proportion of separated zones in the flow channel of the centripetal pump is small, the flow coefficient C of the flow channel is greater than 1, and the total pressure recovery coefficient of the centripetal pump is high. In an unsteady working state, the separation zone in the flow channel of the centripetal pump accounts for a large proportion, the flow channel flow coefficient C is less than 1, and the total pressure recovery coefficient of the centripetal pump is low. An unsteady working state can easily occur in small flow, high-speed conditions. By analyzing the working state and flow characteristics of the centripetal pump, the mechanism of the influence of the flow, outlet fuel discharge pressure, and rotational speed on the working state of the centripetal pump is revealed, which provides a basis for the stable operation of the centripetal pump.

Investigation on the Influence of Film Cooling Structure on the Flow and Heat Transfer Characteristics of Axisymmetric Plug Nozzle

Some numerical simulations were used to study the effects of blowing ratio (0.25–0.5), hole diameters(0.65~1 mm), and hole inclination angle (30~60°) of the film cooling structure on the cooling and aerodynamic characteristics of the axisymmetric plug nozzle under the condition of transonic. The results showed that the bow shocks appear near the film holes on the plug wall in the supersonic region, which cause the wall cooling effectiveness of the plug to decrease. Compared with the baseline plug, the blowing ratio ranged from 0.25 to 0.5, the wall average temperature of the rear plug decreased by 34.4~48.1%, the thrust coefficient and total pressure recovery coefficient decreased by 0.31~0.61% and 0.52~0.93%, respectively. When the perforated percentage is constant, the wall cooling effectiveness increases with the decrease of the hole diameters. The increase in the inclination angle of the film holes lead to the decrease in cooling effectiveness and aerodynamic performance. This is because the penetration ability of the cooling air to the mainstream is enhanced, and the obstruction to the mainstream boundary layer is increased, resulting in the increase of bow shock intensity near the film holes in the supersonic region.

Control-Volume-Based Exergy Method of Truncated Busemann Inlets in Off-Design Conditions

A scramjet engine consisting of several components is a highly coupled system that urgently needs a universal performance metric. Exergy is considered as a potential universal currency to assess the performance of scramjet engines. In this paper, a control-volume-based exergy method for the Reynolds-averaged Navier–Stokes solution of truncated and corrected Busemann inlets was proposed. An exergy postprocessing code was developed to achieve this method. Qualitative and quantitative analyses of exergies in the Busemann inlets were performed. A complete understanding of the evolution process of anergy and the location where anergy occurs in the inlet at various operation conditions was also obtained. The results show that the exergy destroyed in the Busemann inlet can be decomposed into shock wave anergy, viscous anergy and thermal anergy. Shock wave anergy accounts for less than 4% of the total exergy destroyed while thermal anergy and viscous anergy, in a roughly equivalent magnitude, contribute to almost all the remaining. The vast majority of inflow exergy is converted into boundary pressure work and thermal exergy. Some of the thermal exergy excluded by the computation of the total pressure recovery coefficient belongs to the available energy, as this partial energy will be further converted into useful work in combustion chambers.

Numerical study on aerodynamic performance of an intake duct affected by ground effect

Abstract This paper numerically studied the aerodynamic performance of an intake duct affected by the ground effect on a mobile test bench. The simulations were conducted under no wind and headwind conditions. The time evolution of the ground effect indicates that the coherent structure of the vortex system is mainly composed of the ground vortex, the horse-shoe vortex, and the creeping vortex under no wind condition. And it is mainly composed of the ground vortex, the trailing vortex, and the creeping vortex under headwind condition. Compared to the results under no wind condition, the integral vorticity of the ground vortex is larger than that under the headwind condition. The difference of the total pressure recovery coefficient is small, and the total pressure distortion index is large. The results show that with the decrease of the velocity at the intake duct outlet, the intensity of the ground vortex decreases, and the total pressure recovery coefficient at the intake duct outlet increases.

Numerical Investigation of a Vortex Diverter Designed for Improving the Performance of the Submerged Inlet

The submerged inlet exhibits good stealth characteristics and lower drag, but it has a low total pressure recovery coefficient and high distortion rate, which limits its widespread application. This paper proposes a vortex diverter aimed at enhancing the performance of the submerged inlet and investigates the aerodynamic coupling mechanism between the vortex diverter and the submerged inlet in detail. Firstly, based on the flow field characteristics of the submerged inlet, the design principles of the vortex diverter are proposed. Then, the impact of the vortex diverter on the flow field of the submerged inlet is analyzed using the numerical method. Finally, the matching design between the vortex diverter and the submerged inlet is explored. The results show that the vortex diverter improves the average total pressure of the airflow inside the inlet by exhausting the low-energy flow from the larger radius side of the inlet, thereby suppressing flow separation and enhancing flow field uniformity. The vortex diverter improves the intake performance of the submerged inlet under different incoming flow Mach numbers, inlet exit Mach numbers, angles of attack, and small sideslip angles. The maximum increase in the total pressure recovery coefficient is 3.1099%, and the maximum reduction in the circumferential total pressure distortion is 49.5207%. Among the design parameters, the horizontal distance between the leading edge of the vortex diverter and the inlet lip has the greatest influence on the intake performance, and the best control effect is achieved when the vortex diverter is installed at the throat position. Furthermore, after installing the vortex diverter, reducing the side-edge angle of the entrance appropriately can effectively reduce the intensity of the secondary flow, thereby improving the total pressure recovery at the exit and reducing the distortion rate.

Transient Flow Evolution of a Hypersonic Inlet/Isolator with Incoming Windshear

In this paper, a novel flow perturbation model meant to investigate the effects of incoming wind shear on a hypersonic inlet/isolator is presented. This research focuses on the transient shock/boundary layer interaction and shock train flow evolution in a hypersonic inlet/isolator with an on-design Mach number of 6.0 under incoming wind shear at high altitudes, precisely at an altitude of 30 km with a magnitude speed of 80 m/s. Despite the low intensity of wind shear at high altitudes, the results reveal that wind shear significantly disrupts the inlet/isolator flowfield, affecting the shock wave/boundary layer interaction in the unthrottled state, which drives the separation bubble at the throat to move downstream and then upstream. Moreover, the flowfield behaves as a hysteresis phenomenon under the effect of wind shear, and the total pressure recovery coefficients at the throat and exit of the inlet/isolator increase by approximately 10% to 12%. Furthermore, this research focuses on investigating the impact of wind shear on the behavior of the shock train. Once the inlet/isolator is in a throttled state, wind shear severely impacts the motion of the shock train. When the downstream backpressure is 135 times the incoming pressure (p0), the shock train first moves upstream and gradually couples with a cowl shock wave/boundary layer interaction, resulting in a more significant separation at the throat, and then moves downstream and decouples from the separation bubble at the throat. However, if the downstream backpressure increases to 140 p0, the shock train enlarges the separation bubble, forcing the inlet/isolator to fall into the unstart state, and it cannot be restarted. These findings emphasize the need to consider wind shear effects in the design and operation of hypersonic inlet/isolator.

Mode transition control law analysis of ammonia MIPCC aeroengine considering inlet–compressor safety matching

Inlet–Compressor (I–C) safety matching is a key technology for successfully implementing the mode transition of a turbine-based combined-cycle (TBCC) engine. To investigate the performance variation in Ammonia mass injection pre-compressor cooling (MIPCC) aeroengines (AMAs) during the mode transition, a model considering inlet–compressor engine safety matching was developed in this study. We provide a computation for the inlet unstart margin ξ which considers the positive shock wave assumption, I–C flow conservation equation, and combined inlet characteristics to assess the degree of I–C matching. Furthermore, the AMA characteristics indicate that the total pressure recovery coefficient of the hypersonic combined inlet has significantly affects the engine performance. The restraint ξ also ensures engine safety. A control law guided by ξ and the combustor outlet temperature (ξ–COT) was proposed to enhance the specific impulse and specific thrust of the AMA. Compared with the speed and COT control laws, the specific impulse and specific thrust were maximally improved by 2.93 % and 5.47 %, respectively, when ξ = 0.2. Both were further improved by 5.91 % and 10.49 %, respectively, when ξ = 0.1. Finally, the high-thrust performance results of the AMA, with the optimized control law ξ–COT during the mode transition process, were obtained.

Aerodynamics and bird ingestion characteristics of a bulge-adjustable turboprop engine inlet

Turboprop aircraft plays an important role in the field of regional airliners and military transport. During take-off or landing, turboprop aircrafts frequently encounter bird flocks that can be ingested into the engine inlet, thereby posing a significant threat to flight safety. Consequently, the aerodynamic characteristics and bird ingestion characteristics of turboprop inlet are investigated in this article. An unsteady computational fluid dynamics (CFD) approach has been proposed to calculate bird ingestion characteristics in a turboprop inlet with different bulge heights under propeller interference by combining a self-developed collision-rebound model with the overset mesh method and 6-degree of freedom (6-DOF) motion method. Considering the characteristics of bird strikes, a parameterized controlled bulge of turboprop inlet with a scavenge duct has been designed. Moreover, the trajectory characteristics of the bird, bird-inlet coupling flowfield characteristics and the aerodynamic characteristics are analyzed in depth. The bulge of the inlet plays a positive role in preventing the bird object from entering the engine duct at a 23° incidence angle. The bird rebounds from the upper bulge and is discharged through the scavenge duct at MaAIP=0.45. The results show that the growth of inlet bulge height has the minimal influence on the total pressure recovery coefficient σ, with only approximately 0.3% variation within the bulge height range of interest. However, it significantly affects the total pressure distortion DC60 with an increase up to 84%.

Aerodynamic analysis of an airframe-inlet integrated full-waverider vehicle designed by osculating cone theory

Recently proposed, the airframe/inlet integrated full-waverider vehicle design methodology introduces an innovative approach to the design of air-breathing hypersonic vehicles. This methodology, compared to traditional integrated waverider design methods, leverages the features of the waverider to enhance the compression ability of the precompression surface and the lift-to-drag ratio characteristics of the vehicle. However, existing full-waverider design methods, based on cone-derived waverider theory, result in a gas cross-flow phenomenon at the inlet lip due to the conical shockwave. This paper suggests an airframe/inlet integrated full-waverider vehicle design method that utilizes the osculating cone waverider theory to improve gas intake quality. Initially, we present the method to obtain profile lines and basic flow field models at different osculating planes of the vehicle. We then validate the osculating cone integrated full-waverider vehicle design methodology through computational fluid dynamics. Although the viscous effect slightly reduces the lift-to-drag ratio of the vehicle and the total pressure recovery coefficient at the isolator exit, the shock wave remains attached to the leading-edge. Finally, a comparison with the cone-derived integrated full-waverider vehicle reveals that the osculating cone integrated full-waverider vehicle achieves similar aerodynamic performance, but with superior gas intake quality. The results indicate that the methodology proposed herein holds promise for guiding future engineering applications of integrated full-waverider vehicles.

Numerical investigation on spray and flow field characteristics of elliptical nozzle in supersonic crossflow

ABSTRACT It has been suggested that using an elliptical nozzle could enhance the atomization and mixture quality. In order to explore the elliptical liquid jet on spray and flow field characteristics in supersonic airflow, the VOF-Spray one-way coupling model was adopted to study the nozzle exit patterns, size, and velocity distribution of droplets. The research also compared the flow field characteristics in Mach 2.85 air crossflow under the application of different orifices. The numerical results showed that elliptical 4 nozzle (Aspect ratio of 4) had better flow performance than circular nozzle. The cavitation area and turbulent disturbance of elliptical 4 nozzle were mainly concentrated in the major axis direction. The elliptical 4 jet always had a lower penetration depth than circular jet. At the flow range of 0 ~ 50 mm, the penetration depth of elliptical 4 jet is roughly 12% less than that of circular jet. Another finding was that the spanwise extension range of elliptical 4 spray was larger than the circular spray. The spanwise extension range of elliptical 4 spray increased by about 19.4% at most in the flow range of 0 ~ 50 mm. Additionally, the elliptical 4 spray had the smallest Santer Mean Diameter (SMD). The elliptical 4 jet had a 2.9% smaller Santer Mean Diameter than the circular jet. Due to the greater windward surface area of the elliptical 4 orifice, the gas streamwise velocity near the bottom wall was reduced. A reverse vortex pair with distinct characteristics could be observed under the application of elliptical 4 orifice. Furthermore, compared with the circular orifice, the shock wave angle generated by elliptical 4 orifice reduced by 9.7%, and the total pressure recovery coefficient rose by 2.3%.

Control of Cowl Shock/Boundary Layer Interaction in Supersonic Inlet Based on Dynamic Vortex Generator

A shock wave/boundary layer interaction (SWBLI) is a common phenomenon in supersonic inlet flow, which can significantly degrade the aerodynamic performance of the inlet by inducing boundary layer separation. To address this issue, in this paper, we propose the use of a dynamic vortex generator to control the SWBLI in a typical supersonic inlet. The unsteady simulation method based on dynamic grid technology was employed to verify the effectiveness of the proposed method of control and investigate its mechanism. The results showed that, in a duct of finite width at the inlet, the SWBLI generated complex three-dimensional (3D) flow structures with remarkable swirling properties. At the same time, vortex pairs were generated close to the side wall as a result of its presence, and this led to the intensification of transverse flow and, in turn, the formation of a complex 3D structure of the flow of the separation bubble. The dynamic vortex generator induced oscillations of variable intensity in the vortex system in the supersonic boundary layer that enhanced the mixing between the boundary layer flow and the mainstream. Meanwhile, the unique effects of “extrusion” and “suction” in the oscillation process continued to charge the airflow, and the distribution of velocity in the boundary layer significantly improved. As the oscillation frequency of the vortex generator increased, its charging effect on low-velocity flow in the boundary layer increased, and its control effect on the flow field of the SWBLI became more pronounced. The proposed method of control reduced the length of the separation bubble by 31.76% and increased the total pressure recovery coefficient at the inlet by 6.4% compared to the values in the absence of control.

Study on Aerodynamic Design of the Front Auxiliary Inlet

Submerged inlets have been widely used in advanced aircraft due to their excellent stealth characteristics, but they also suffer from poor aerodynamic performance. To improve the aerodynamic efficiency while maintaining stealth capabilities, this paper proposes a design scheme for a front auxiliary inlet with an inlet grille. The front auxiliary inlet is connected to the main inlet to form a composite inlet system. The low-energy upstream airflow that accumulates at the inlet is guided by the front auxiliary inlet to flow into the mainstream, resulting in a stable and high-quality airflow. A certain type of cruise missile was used as the research subject, and intake systems with and without front auxiliary inlets were constructed to compare the inlet performance of the two configurations using the CFD method. Additionally, a sensitivity analysis of the main design parameters of the front auxiliary inlet was carried out. The study reveals that a reasonable design of the front auxiliary inlet can prevent low-energy airflow, which accumulates on the missile body surface, from directly entering the inlet. Moreover, the front auxiliary inlet can inject additional mechanical energy into the low-energy airflow, inhibit airflow separation, and improve the uniformity of the flow field. Under cruise conditions, the total pressure recovery coefficient of the front auxiliary inlet configuration increased by 12.39% compared to the model without a front auxiliary inlet configuration. Furthermore, the total pressure distortion index was reduced by 47.24%.

Surrogate-based optimization on bump for shock wave/boundary layer interaction control

Supersonic and hypersonic vehicles have complex shock wave/boundary layer interaction problems during flight, and it adversely affect the aerodynamic performance of the vehicle. In this paper, the bump passive control approach is introduced to control the shock wave/boundary layer interaction, and its configuration is optimized with the minimum separation area and the maximum total pressure recovery coefficient. The Kriging model is established by means of the Latin hypercube experimental design approach to obtain the sample points. The Non-dominated Sorting Genetic Algorithm-II (NSGA-II) algorithm is utilized to optimize the parameters of the bump configuration, and the flow field structures of the uncontrolled condition, the special basic condition and the optimized condition are compared and analyzed. The obtained results show that the surrogate-based optimized model shows better performances than the basic condition. In the parameter setting range considered in this study, the optimized Case 3c with the middle position in the optimization front owns a larger total pressure recovery coefficient and a smaller separation area, with its separation area and total pressure recovery coefficient being 0.80 mm2 and 67.39% respectively.

Numerical Investigation of Combustion Mechanism with Multi-Position Injection in a Dual-Mode Combustor

To improve the flame propagation, combustion stability, and uniformity of the temperature field, multi-position injection is applied in a dual-mode combustor by controlling heat release in different locations. Using the chemical reaction of the finite rate combustion model and the detailed reaction mechanism of hydrogen combustion as described by Jachimowski, the influence of different multi-position injection patterns in a dual-mode combustor is analyzed. The one-equation Large Eddy Simulation (LES) turbulence model was chosen to define the sublattice turbulent viscous terms in a three-dimensional scramjet model. Based on a combustion chamber, the effect of the injection equivalent ratio (0.35–0.70), the relative position of the nozzle holes, and the injection pressure on the combustion process and flow field characteristics are analyzed with multi-position injection. The combustion efficiency, total pressure recovery coefficients, and pressure distribution under different operation conditions are compared. We observed that the combustion intensity increases and the upstream combustion shock string distance becomes greater with increased equivalent ratios. When the global equivalent ratio of multi-position injection remains unchanged, the arrangement of nozzles with the small injection spacing, i.e., two injection holes arranged face to face on the upper and lower walls, or the setting of multiple injection holes with the same pressure, can effectively increase the stability rate of the combustion flow field. In addition, the combustion efficiency at the outlet and the internal pressure of the combustion chamber in the stable state are also improved, relative to the increased total pressure loss.

Prepositive Synergistic Bulge Design for Improving Aerodynamic Performance of Submerged Inlet

A submerged inlet has good stealth characteristics and a low external drag, but it also has the disadvantage of low internal flow efficiency. In view of this, a new efficiency enhancement method based on the prepositive synergistic bulge of the inlet’s anterior lip is proposed. Taking the submerged inlet of an aircraft as the baseline configuration, a miniature bulge with a square bottom and an outer convex form is designed in front of the inlet’s anterior lip. Through the convex shape of the bulge, part of the low-energy boundary layer airflow is diverted away from the inlet’s entrance, so that the airflow greatly reduces the flow separation after entering the inlet, and the internal flow performance of the entire submerged inlet is improved. Taking the flow field of an aircraft in the classic cruise state as an example, the simulation analysis results show that the flow field characteristics of the entire submerged inlet are obviously improved after adding the synergistic bulge. The total pressure recovery coefficient of the new inlet configuration increased by 1.36%, the total pressure distortion index decreased by 10.86%, and the body drag only increased by 0.37% compared with the baseline case. According to calculations of synergistic bulge inlet configurations with different design parameters, the effect of this configuration is relatively stable, whereby the aspect ratio of the bulge has the greatest impact on the performance, and its value should not be less than 0.75. In addition to the advantages of not requiring additional components or occupying space and being easy to manufacture, the method of adding a synergistic bulge can improve the aerodynamic performance of the baseline inlet under most cruise flight conditions, and its additional drag is small, which gives it a wide applicability range.

The determination of physical dimensions of a hypersonic three-stage compression system considering panel vibration effects

The panel vibration of air intake system that excitated by fluid-structural interaction, engine thrust variation and noise has decisive influences on the performance of an air-breathing hypersonic vehicle. The design of such type of intake system should contain the determination of configurational parameters and physical dimensions by considering both flow and structural vibrations, and a clear design path is still absent. In this paper, an intake system with three-stage external compression ramps is designed under Mach number of 6 and flight altitude of 24 km, and the configurational parameters are determined based on the shock-on-lip criterion, Oswatitsch criterion and viscous effects modification. The mode vibration characteristics of external compression ramps with panel thicknesses of 0.002 to 0.03 m are analyzed, and appropriate natural frequencies are obtained by constraint conditions of fixed four side boundaries with an additional and more realistic constraint of strengthened internal boundary. The flow field of the intake system with forced panel vibrations of compression ramps is simulated, and the effects of vibration amplitude of 0 to 0.11 m and frequency of 10 to 250 Hz on the intake performance, i.e., total pressure recovery coefficient and mass flow coefficient are clarified. The vibration effects of hypersonic flow on the compression ramp structure and intake performance are studied by fluid-structural coupling simulations. Finally, the available physical dimensions of external compression ramps are determined and a design path of intake system is developed by considering both the structural and intake performance.

Effect of film cooling structures on the cooling characteristics of plugs for two-dimensional (2-D) plug nozzle

Based on a turbulence model experimental validation, some numerical simulations were conducted to investigate the effect of plug film cooling on cooling and aerodynamic characteristics for a 2-D plug nozzle. In this study, two cooling schemes namely the half-covering and full-covering film cooling were designed and involved. For each cooling configuration, the effects of perforated percentage (ranged from 0.5% to 2.0%) and non-dimensional inlet total pressure (ranged from 1.02 to 1.20) of coolant on the cooling and aerodynamic performances are discussed separately. The results confirm that an obvious reduction of plug temperature is achieved under both cooling schemes. While the full-covering film cooling configuration performs better, especially at the plug trailing edge where shock waves occurred. Within the scope of the investigation, a small perforated percentage of 0.5% is favorable due to its superior ability to improve the pressure difference between mainstream and coolant; inversely, a big perforated percentage leads to the coolant flow resistance increasing, which may cause the mainstream to flow back at the plug trailing edge. Rising the inlet total pressure of coolant could enhance the cooling performance but also deteriorate the aerodynamic performance of nozzle. Specifically, compared with the case without cooling configuration, the non-dimensional inlet total pressure of coolant ranged from 1.02 to 1.20, the surface temperature of plug decreased by 20% ～ 45%; the total pressure recovery coefficient decreased by 0.22% ～ 1.26%.

Numerical Investigation of Aerodynamic and Electromagnetic Performances for S-Duct Caret Intake with Boundary-Layer Bleed System

This study presents the numerical results for the aerodynamic and electromagnetic performances of an S-duct caret intake. Using the multilevel fast multipole method (MLFMM) to solve Maxwell equations, the current on the intake surface is calculated, and the radar cross-section (RCS) is analyzed. Moreover, the intake flow field is numerically investigated using the SST k–ω turbulence model to solve the Reynolds-averaged Navier–Stokes equations. Compared to a straight intake, for an S-duct caret intake, the average RCS is lower by 7.65 dB, and the maximum RCS difference value caused by the blade rotation is lower by 6.75 dB. However, the flow capacity deteriorates when the total pressure recovery coefficient decreases by 0.004. Based on the analysis of the aerodynamic and electromagnetic characteristics of different intakes, a double S-duct intake is designed. Compared to a traditional S-duct intake, for the novel intake after model parameter modification, the average RCS is lower by 0.05 dB, and the total pressure distortion (TPD) is lower by 0.18. The analysis of the effects of different boundary-layer bleed systems shows that the symmetrical layout adversely affects the aerodynamic and electromagnetic performances of the S-duct intake, but the unilateral partial layout is beneficial, whose TPD is lower by 0.04 and average RCS is higher by −2.17 dB compared to a straight intake.

Numerical Simulation and Wind Tunnel Test of a Variable Geometry Auxiliary Inlet for a Wide-Body Aircraft Environmental Control System

A variable geometry auxiliary inlet for a wide-body aircraft environmental control system with moveable deflectors operating in a large mass flow rate range is studied through numerical simulation and wind tunnel tests, which yields a design method for the variable geometry auxiliary inlet with high performance. The characteristics of the flow field are studied by numerical simulation. The results show that the favorable pressure gradient and the roll-up vortices are the major impetus that inhales the incoming flow into the inlet. The law of regulation and the performance variation under different conditions are obtained by wind tunnel test. The flow coefficient increases first but then decreases with the increase in the inlet opening, and the pressure rise ratio and total pressure recovery coefficient increase first and then decrease with the increase in the mass flow rate. In general, under the condition of a high Mach number (Ma > 0.4), the inlet opening of this test configuration should not exceed 50%. The deflectors can maintain the normal work of the environmental control system by moving properly to control the mass flow rate of the auxiliary inlet.

Influences of microjet pressure and number of microjets on the control of shock wave/boundary layer interaction

Boundary layer separation is a common phenomenon in supersonic/hypersonic vehicles, and can adversely affect the lift coefficient. Currently, active flow control based on microjets is used to inhibit the separation of the turbulent boundary layer. The obtained results predicted by the three-dimensional Reynolds-averaged Navier–Stokes (RANS) equations and two-equation shear stress transport (SST) k-ω both with Realizable k-ε turbulence models show that the control mechanism of the microjets method is the counter-rotating vortex pair (CVP) generated by the microjet mixes the low-energy flow within the boundary layer with the high-energy flow near the boundary layer. The size of the vortex core in this CVP is the key to controlling SWBLI. The vortex nucleus that is larger and closer to the wall enables better SWBLI control. A lower or higher ratio of microjet pressure to inflow static pressure increases the total volume of the separation bubble and produces a negative gain. When the ratio of the triangular microjet hole bottom edge length to the spacing between the holes is too small, leading to a deterioration in the control effect. The addition of microjets reduces the total pressure recovery coefficient.