Serpentine Inlet Research Articles

Distortion elimination for serpentine duct at various Mach numbers using co-flow jet active flow control

This paper numerically investigates Co-flow Jet (CFJ) active flow control (AFC) for eliminating the AGARD M2129 serpentine inlet distortion with throat Mach number (Mth) varying from 0.42 to 0.79. The configuration was optimized at design Mth of 0.79 in a prior study. The current objective is to investigate the CFJ S-duct performance at off-design Mth with the fixed geometry. The high-order FASIP in-house code is used to conduct 3D Reynolds Averaged Navier-Stokes (RANS) simulation with the one-equation Spalart-Allmaras (S-A) turbulence model. The CFD simulation is well validated with the experiment of AGARD test cases in the studied Mach numbers. Results show that CFJ virtually eliminates the flow distortion for all the studied Mth. However, the off design Mth have lower energy efficiency. At the design Mth, CFJ requires the lowest power coefficient (Pc) to eliminate the flow separation and distortion. The exergy analysis indicates that CFJ S-duct at the design Mth has the highest system energy efficiency with the EIPR (ratio of exergy increase to the power required) exceeding 1. At the off-design Mth, EIPR is lower than 1 because the shifts of separation-onset spot and adverse-pressure-gradient region make CFJ injection and suction no longer at the most efficient location. This leads to the rising of the CFJ power coefficient even though the actual power is less. Nevertheless, the overall system is still very efficient at off-design Mth for the CFJ S-duct with the CFJ energy expenditure recovered by 72%, 90%, 96% for Mth of 0.42, 0.55, and 0.69, respectively.

Overview of Selected Serpentine Duct Simulations Using Computational Fluid Dynamics

Computational fluid dynamics (CFD) was used to perform simulations of multiple serpentine inlet ducts for comparison to experimental data. The simulations included two baseline duct geometries configured to match their respective wind tunnel tests. Each configuration was tested and simulated with and without flow control vanes. Some of the simulations included the geometric detail of the instrumentation probes at the aerodynamic interface plane. This paper highlights selected issues that impact robustness and accuracy. The topics include far-field boundary conditions and various methods of modeling the instrumentation probes. This paper summarizes discussion topics and recommendations for CFD practitioners preparing to simulate similar cases.

Design and Numerical Study of Serpentine Duct using Flow Controllers for Flow Characteristics Enhancement

In the current work, computational study has been carrying out on the both baseline serpentine duct to know the flow behaviour and baseline with vortex generator to know the effect of vortex generators on the flow properties of the duct. A set of vortex generators are chosen for the study, by keeping all other parameters constant. Here in current work, for modelling the duct and vortex generator CATIA V5 is used; to discretizing the geometry ICEM CFD is used. For the analysis, the computational fluid dynamics tool FLUENT is used, SST k-ù model available in the FLUENT is used for the computations, and the results of current work will be validated with experimental results and shows that they are converged. Use of vortex generator in the current work proved extremely effective in enforcing the active flow control. Results obtained in this work are to study the pressure recovery at the engine face. The results ensure that the current work on the baseline model of serpentine inlet duct and model with vortex generator comparison is well established in case of pressure recovery and decrease in engine face distortion. After comparing the result of both baseline duct and duct with vortex generator we are expecting at least 25-35% of pressure recovery at the engine face. In the present study significant improvements characterized by an improved pressure recovery and decrease in distortion are observed over the baseline model of serpentine duct for vortex generator model. This reduction in pressure loss and distortion enhances the engine performance of thrust and also improves the overall performance of the engine.

Aerodynamic Response of a Serpentine Inlet to Horizontal Periodic Gusts

Gust is a common atmospheric turbulence phenomenon encountered by aircraft and is one major cause of several undesired instability problems. Although the response of aircraft to the incoming gust has been widely investigated within the subject of external-flow aerodynamics in the past decades, little attention is paid to its effects on the internal flow within aircraft engines. In this paper, a newly implemented Field Velocity Method (FVM) in OpenFOAM is used to simulate the flow field and aerodynamic responses of a serpentine inlet exposed to non-stationary horizontal sinusoidal gusts. Validations are performed on the results obtained based on the baseline Computational Fluid Dynamics (CFD) solver and the gust modeling method. Finally, the flow field and aerodynamic characteristics of the serpentine inlet under horizontal sinusoidal gust conditions are comprehensively investigated. It is found that the gusts not only significantly change the flow structure but also play an unfavorable role in the total pressure distortion of the serpentine inlet. This finding shows the necessity to consider gust effects when designing and evaluating the performance of aircraft engines.

Structural design of a morphing serpentine inlet using a multi-material topology optimization methodology

A promising avenue for development in the aerospace industry involves relocation of engines from conventional locations beneath the wings, to within the aircraft fuselage. Aircraft with these so-called embedded engines have the potential to increase engine efficiency, but necessitate the use of a serpentine engine inlet duct (S-duct) to provide the propulsion system with air. The optimal shape of an aircraft S-duct varies with flight condition, incentivizing the use of morphing systems to vary inlet parameters during flight: allowing for continual supply of the optimal airflow level. Design of a morphing system therefore requires collaboration across multiple disciplines, including aerodynamic analysis and structural analysis. This work presents a methodology for the structural optimization of morphing systems utilizing aerodynamic shape optimization results as inputs, in order to assess the relationship between morphing performance and structural stiffness. The methodology is implemented on a baseline morphing S-duct model for which shape optimization has been previously conducted. Structural optimization is conducted using a gradient-based multi-material topology optimization software with multi-phase penalization. While the conclusions of this work indicate that the impact of multiple material optimization in the S-duct case study is minimal, the methodology does provide non-intuitive designs capable of supporting morphing. At the expense of structural stiffness, the methodology is shown to increase morphing performance through the generation of compliant mechanisms. A parameter study conducted on the S-duct model successfully proves the ability of the methodology to assess the trade-offs between structural and morphing performance. By emphasizing morphing performance, mass reductions from 1.75 kg to 0.201 kg were observed at the expense of a 94% reduction in fatigue lifecycle.

Dynamic characteristics analysis of hammer shock in serpentine inlet

To analyze the dynamic characteristics of hammer shocks caused by engine surge, a serpentine inlet with a front fuselage is simulated under conditions of subsonic inflow and three flight angles based on the improved delayed detached eddy simulation method. An unsteady back pressure boundary condition in an aerodynamic interface plane is used to simulate the overpressure during engine surge. There is a certain angle between the normal line of the hammer shock and the centerline, which is approximately equal to the corresponding flight angle. The inlet wall pressures are above 2.3 times of the free-stream static pressure, and even local transient pressures reach more than 3 times. Complex flow field structures are generated behind the shock, which are affected by the centrifugal force and lateral pressure gradient. The velocities and intensities of the hammer shocks and the pressure distributions on the wall and the shapes of the flow field structures are greatly affected by the flight angle. In a large yaw angle, the hammer shock velocity is the fastest, and the internal airflow and load are the most severe. In particular, it is necessary to consider the influence of the ultra-high dynamic loads in the opposite direction at the two bends in a short period of time.

Closed-loop flow control of an ultra-compact serpentine inlet based on nondimensional model

Ultra-compact serpentine inlet faces serve inlet-engine compatibility issues due to flow distortion. To ensure inlet-engine compatibility over a wide range of Mach number, novel active flow control techniques with the ability of being opened or adjusted as needed draw many attentions in recent years. In this paper, a feedback control system was developed based on the method of microjet blowing. The proposed system includes a pressure adjusting valve to adjust the control effort, a dynamic pressure sensor to sense the inlet distortion intensity, a signal processing instrument to calculate the Root-Mean-Squared (RMS) pressure, and a controller to implement feedback control. To achieve high quality closed-loop controls at dynamic conditions, a novel nondimensional feedback method was developed. The advantage of this nondimensional method was validated at both off-design and arbitrarily changing Mach number conditions. With a sectional PI control law, the RMS control error reduced more than 56% at arbitrary changing conditions. Works in this paper also showed that the dynamics of this nondimensional system can be simplified as a stable second-order overdamped system.

Numerical and Experimental Investigation of a Serpentine Inlet Constructed with a Polynomial

The investigation of the serpentine inlet is particularly important due to the high stealth requirements of the next generation unmanned aerial vehicle (UAV). The paper presents a design method of generating a serpentine inlet controlled by a center line which is a polynomial. The serpentine inlet is designed considering the positions of the airborne equipment, the stealth and aerodynamic performances. The study includes numerical simulations based on the Reynolds-averaged Navier-Stokes simulation approach and experiments in wind tunnel. The aerodynamic performances of the serpentine inlet at different inflow Mach numbers, mass flow ratios and angles of attack are obtained. Moreover, the flow structures are studied at several critical conditions and the performance and distortion of the designed serpentine inlet are acceptable without any flow control.

Flow Characteristics of an Ultracompact Serpentine Inlet with an Internal Bump

AbstractAn ultracompact serpentine inlet, which contains a compact diffuser as short as 2.3 times the exit diameter, is experimentally and numerically studied. An internal bump is introduced to red...

Microjet flow control in an ultra-compact serpentine inlet

Microjets are used to control the internal flow to improve the performance of an ultra-compact serpentine inlet. A highly offset serpentine inlet with length-to-diameter ratio of 2.5 is designed and static tests are conducted to analyze the internal flow characteristics in terms of pressure recovery, distortion and flow separation. Flow separation is encountered in the second S-turn, and two strong counter-rotating vortices are formed at the aerodynamic interface plane (AIP) face which occupy a quarter of the outlet area and result in severe pressure loss and distortion. A flow control model employing a row of microjets in the second turn is designed based on the internal flow characteristics and simplified CFD simulations. Flow control tests are conducted to verify the control effectiveness and understand the characteristics as a function of inlet throat Mach number, injection mass flow ratio, jet Mach number and momentum coefficient. At all test Mach numbers, microjet flow control (MFC) effectively improves the recovery and reduces the distortion intensity. Between inlet throat Mach number 0.2 and 0.5, the strong flow separation in the second S-turn is suppressed at an optimum jet flow ratio of less than 0.65%, resulting in a maximum improvement of 4% for pressure recovery coefficient and a maximum decrease of 75% for circumferential distortion intensity at cruise. However, in order to suppress the flow separation, the injection rate should retain in an effective range. When the injection rate is higher than this range, the flow is degraded and the distortion contour is changed from 90° circumferential distortion pattern to 180° circumferential distortion pattern. Detailed data analysis shows that this optimum flow ratio depends on inlet throat Mach number and the momentum coefficient affects the control effectiveness in a dual stepping manner.

Effects of Boundary-Layer Ingestion on the Aero-Acoustics of Transonic Fan Rotors

The use of boundary-layer-ingesting, embedded propulsion systems can result in inlet flow distortions where the interaction of the boundary-layer vorticity and the inlet lip causes horseshoe vortex formation and the ingestion of streamwise vortices into the inlet. A previously-developed body-force-based fan modeling approach was used to assess the change in fan rotor shock noise generation and propagation in a boundary-layer-ingesting, serpentine inlet. This approach is employed here in a parametric study to assess the effects of inlet geometry parameters (offset-to-diameter ratio and downstream-to-upstream area ratio) on flow distortion and rotor shock noise. Mechanisms related to the vortical inlet structures were found to govern changes in the rotor shock noise generation and propagation. The vortex whose circulation is in the opposite direction to the fan rotation (counter-swirling vortex) increases incidence angles on the fan blades near the tip, enhancing noise generation. The vortex with circulation in the direction of fan rotation (co-swirling vortex) creates a region of subsonic relative flow near the blade tip radius that decreases the sound power propagated to the far-field. The parametric study revealed that the overall sound power level at the fan leading edge is set by the ingested streamwise circulation, and that for inlet designs in which the streamwise vortices are displaced away from the duct wall, the sound power at the upstream inlet plane increased by as much as 9 dB. By comparing the far-field noise results obtained to those for a conventional inlet, it is deduced that the changes in rotor shock noise are predominantly due to the ingestion of streamwise vorticity.

Active Flow Control of a Boundary Layer-Ingesting Serpentine Inlet Diffuser

The effectiveness of active flow control for the reduction of flow distortion in boundary layer-ingesting serpentine inlets was examined. The intended purpose of the flow control was to redistribute the ingested low-momentum boundary-layer fluid around the periphery of the diffuser to reduce the flow distortion at the engine face. A variety of different blowing jet and suction hole configurations were analyzed using computational fluid dynamics, and a subset of these configurations was experimentally validated. Boundary-layer suction alone proved to be ineffective in reducing flow distortion, but suction did enhance the distortion reduction capability of blowing flow control when blowing mass flow rates were greater than 1% of the total inlet mass flow rate. To reduce the demand for bleed air from the engine compressor, it is suggested to augment blowing flow with the flow entrained from the suction holes using an ejector-pump concept. Application of the notional ejector-pump model resulted in a maximum decrease in engine-face distortion of 75% as measured by the parameter, as compared with a 28% reduction with blowing alone. Investigations also showed that the most practically effective flow control configurations maintained jet individuality to directly counter duct secondary flows while not combining to form large-scale vortices.

Shock Propagation and MPT Noise From a Transonic Rotor in Nonuniform Flow

One of the major challenges in high-speed fan stages used in compact, embedded propulsion systems is inlet distortion noise. A body-force-based approach for the prediction of multiple-pure-tone (MPT) noise was previously introduced and validated. In this paper, it is employed with the objective of quantifying the effects of nonuniform flow on the generation and propagation of MPT noise. First-of-their-kind back-to-back coupled aero-acoustic computations were carried out using the new approach for conventional and serpentine inlets. Both inlets delivered flow to the same NASA/GE R4 fan rotor at equal corrected mass flow rates. Although the source strength at the fan is increased by 38 dB in sound power level due to the nonuniform inflow, far-field noise for the serpentine inlet duct is increased on average by only 3.1 dBA overall sound pressure level in the forward arc. This is due to the redistribution of acoustic energy to frequencies below 11 times the shaft frequency and the apparent cut-off of tones at higher frequencies including blade-passing tones. The circumferential extent of the inlet swirl distortion at the fan was found to be two blade pitches, or 1/11th of the circumference, suggesting a relationship between the circumferential extent of the inlet distortion and the apparent cut-off frequency perceived in the far field. A first-principles-based model of the generation of shock waves from a transonic rotor in nonuniform flow showed that the effects of nonuniform flow on acoustic wave propagation, which cannot be captured by the simplified model, are more dominant than those of inlet flow distortion on source noise. It demonstrated that nonlinear, coupled aerodynamic and aero-acoustic computations, such as those presented in this paper, are necessary to assess the propagation through nonuniform mean flow. A parametric study of serpentine inlet designs is underway to quantify these propagation effects.

Numerical and experimental investigation of a serpentine inlet duct

This article presents a numerical and experimental investigation of the flow inside an ultra-compact, serpentine inlet duct. The numerical analysis used two flow solvers: FLUENT®, a commercial code, and UNS3D, an in-house code. The flow was modelled using the Reynolds-averaged Navier-Stokes equations. The turbulence effects were modelled by using the shear-stress transport k–ω model. The numerical investigation was compared against experimental data obtained in an open-circuit, low-speed wind tunnel in the Fluid Dynamics Laboratory at Texas A&M University. The numerical simulations and experimental testing were performed to reveal the separation points and the strong secondary flow phenomena within the inlet. UNS3D overpredicted the location of the first separation point by 9 mm and the location of the second separation point by 1 mm, while the area-averaged pressure loss coefficient was 5% higher than in the experiment. The numerical results of UNS3D agreed better with the experiment than those of FLUENT.

Thermal performance analysis of a closed wet cooling tower

A detailed model was developed and employed to examine the thermal performance of a closed wet cooling tower. The model is capable of predicting the variation of air thermodynamic properties, sprayed and serpentine water temperature as well as heat transfer rates exchanged between air and falling water stream inside the indirect wet cooling tower. The reliability of simulations was tested against experimental data obtained from the literature. A parametric study was conducted to evaluate the thermal behaviour of the indirect cooling tower under various air mass flowrates, serpentine water mass flowrates and inlet temperatures. The results of the theoretical investigation revealed an increase in cooling capacity and percentage loss of sprayed water due to evaporation, with increasing air mass flowrate. On the other hand, the increase of serpentine water mass flowrate resulted in slight increase in the overall temperature reduction of serpentine water. The effect of variable serpentine water inlet temperature on thermal performance of the indirect wet cooling tower was insignificant compared to other cases.

Serpentine Inlet Performance Enhancement Using Vortex Generator Based Flow Control

In order to provide the line of-sight blockage of the engine face for an advanced Uninhabited Combat Air Vehicle(UCAV), a highly curved serpentine inlet is proposed and experimentally studied. Based on the static pressure distribut ion measurement along the wall, the flow separation is found at the top wall of the second S duct for the baseline inlet design, which yields a high flow distortion at the exit plane. To improve the flow uniformity, a single array of vortex generators (VGs) is employed within the inlet. In this experimental study, the effects of mass flow ratio, free stream Mach number, angle of attack and yaw on the performance of a serpentine inlet instrumented with VGs are obtained. Results indicate: (1) Compared with the baseline serpentine design without flow control the application of the VGs promotes the mixing of core flow and the low momentum flow in the boundary layer and thus prevents the flow separation. Under the design condition, the exit flow distortion (▪) decreases from 11. 7% to 2.3% by using the VGs. (2) With the descent of the free stream Mach number the total pressure loss decreases. How ever, the circular total pressure distortion increases. When the angle of attack rises from - 4° to 8°, the total pressure recovery and the circular total pressure distortion both go down. In addition, with the increase of yaw the total pressure recovery is fairly constant, while the circular total pressure distortion ascends gradually. (3) When Ma0=0.6-0.8, α= −4°-8° and β= 0°-6°, the total pressure recovery varies between 0.936 and 0.961, the circular total pressure distortion coefficient varies between 1.4% and 5.4% and the synthesis distortion coefficient has a ranges from 3.8% to 7.0%. The experimental results confirm the excellent performance of the newly designed serpentine inlet incorporating VGs.

Effectiveness of a Serpentine Inlet Duct Flow Control Technique at Design and Off-Design Simulated Flight Conditions

AbstractAn experimental investigation was conducted in a static ground test facility to determine the effectiveness of a serpentine inlet duct active flow control technique for two simulated flight conditions. The experiments used a scaled model of a compact, diffusing, serpentine, engine inlet duct developed by Lockheed Martin with a flow control technique using air injection through microjets at 1% of the inlet mass flow rate. The experimental results, in the form of total pressure measurements at the exit of the inlet, were used to predict the stability of a compression system through a parallel compressor model. The inlet duct was tested at cruise condition and angle of attack flight cases to determine the change in inlet performance due to flow control at different flight conditions. The experiments were run at an inlet throat Mach number of 0.55 and a resulting Reynolds number, based on the hydraulic diameter at the inlet throat, of 1.76*105. For both of the flight conditions tested, the flow control technique was found to reduce inlet distortion at the exit of the inlet by as much as 70% while increasing total pressure recovery by as much as 2%. The inlet total pressure profile was input in a parallel compressor model to predict the changes in stability margin of a compression system due to flow control for design and off-design flight conditions. Without flow control, both cases show a reduction in stability margin of 70%. With the addition of flow control, each case was able to recover a significant portion (up to 55%) of the undistorted stability margin. This flow control technique has improved the operating range of a compression system as compared to the same inlet duct without flow control.

Flow Separation Within the Engine Inlet of an Uninhabited Combat Air Vehicle (UCAV)

This paper discusses the structure of the flow within the engine inlet of an uninhabited combat air vehicle (UCAV). The UCAV features a top-mounted, serpentine inlet leading to an engine buried within the fuselage. The performance of the inlet is found to depend strongly on a flow separation that occurs within the inlet. Both the time-averaged and the unsteady structure of this separation is studied, and an argument relating the inlet performance to the behavior of this separation is suggested. The results presented in this paper also suggest that there are considerable aerodynamic limitations to further shortening or narrowing of the inlet. Since there are substantial, system level benefits from using a smaller inlet, the case for separated flow control therefore appears clear.