Circumferential Pressure Gradient Research Articles

On the large-scale vortices in an L/D = 2.30 serpentine diffuser with internal bump

In this paper, the generation and spatial evolution of large-scale vortices in an ultra-compact serpentine diffuser (L/D = 2.30) featuring an integrated internal bump are experimentally and numerically investigated. Due to reduction of the local flow area caused by the internal bump, a favorable streamwise pressure gradient is induced first and then followed by an adverse one. Concurrently, a high-pressure region exists on the side of the diffuser as it contracts, yielding a sideward circumferential pressure gradient and then an inward one. Such circumferential pressure gradients divert the boundary layer on the windward side of the internal bump to the side and induce a spanwise U-shaped vortex on the leeward side, which constrains the development of flow separation in this region. The circumferential pressure gradients further induce two pairs of streamwise vortices located at the top and bottom of the diffuser, respectively. These vortex pairs entrain the low-energy fluids and transport them to the aerodynamic interface plane (AIP), resulting in two pairs of low-total-pressure regions and swirling flows. Moreover, in the experimentally studied range of the exit Mach number MAIP from 0.293 to 0.546, the distributions of low-total-pressure regions and swirl angles at the AIP exhibit a small variation, indicating that the internal flow field structures are insensitive to MAIP.

Dynamic separation on an accelerating prolate spheroid

Time-varying flow separation on an accelerating prolate spheroid has been studied at various angles of incidence. Instantaneous pressure and scanning stereoscopic particle image velocimetry were used to shed light on the evolution of cross-flow structures for the Reynolds number ( $Re$ ) range of $1.0\times 10^6\leq Re \leq 1.5\times 10^6$ . The movement of separation lines is examined for various model accelerations to investigate on the interplay between acceleration and flow separation. The results demonstrate that for axial accelerations, the streamwise pressure distribution in the rear part of the prolate spheroid switches from an adverse to a favourable pressure gradient. At the same time, the circumferential adverse pressure gradient present during steady motion vanishes during said accelerations. In contrast, both streamwise and circumferential adverse pressure gradients strengthen when the model is axially decelerated. These dynamic pressure distributions influence the location of the separation line, which in turn moves closer to the model meridian during accelerations while moving outwards during decelerations. The streamwise vorticity distribution and the streamwise circulation both show how the separation-line position impacts the vortex formation. A high-vorticity region near the model surface is established during acceleration. In contrast, a decelerating model leads to transport of high-vorticity fluid into the outer area of the cross-flow separation. We further assess the memory effects following the near-impulsive velocity changes. The cross-flow retains the memory of moving separation lines shortly after the acceleration. However, the separation recovers quickly to a steady state.

Investigation of Vaned-Recessed Casing Treatment in a Low-Speed Axial Flow Compressor, Part I: Time-Averaged Results

This paper investigates the effects of two modifications to a vaned recessed casing treatment. First, the shape of a circular curve was used in the top of the treated casing. Second, a fully curved guide vane was also applied. The goals of the modifications are to enhance the flow recirculation as well as to relieve the low-speed flow, which is normally accumulated within the corners of the vaned recessed casing treatment. The solid casing in addition to the two vaned recessed configurations with 23.2% and 53.5% rotor blade tip axial chord exposure have been studied numerically. The results indicated that two mechanisms are involved in the stall margin enhancement. First, the circumferential pressure gradient is reduced for both configurations. The reduction in pressure gradient largely reduces the development of tip leakage vortex and, thus, the generation of low-speed fluid is diminished. Second, the main flow/tip leakage interface moves toward downstream and the movement of interface toward the leading edge is delayed. The second configuration with a greater rotor blade tip exposure enables extra flow recirculation due to decreasing surface area and, therefore, could be superior to the application of the first casing treatment configuration. The major streamlines within the casing treatment are also discussed. The time-averaged results are presented in this paper, while the unsteady results including instantaneous flow fields, origins of the unsteadiness and frequency analysis are discussed in part II.

A Three-Dimensional Body Force Modeling of Fans in Windmilling Condition and Its Application

To investigate the aerodynamic characteristics of the fan in windmilling conditions, a new body force model with the fan rotational speed prediction model was developed. The fan rotational speed prediction model was built based on the balance of fan output torque and resistance torque. The rotational speed of the fan spool can be iteratively solved simultaneously with solving the governing equations without requiring mass flow rate or other inputs. The comparison with the experimental results shows that using the body force model can accurately predict the rotational speed of the fan spool under different operating conditions. The radial distribution of flow parameters can be obtained. Moreover, numerical simulations of the fan under different circumferential total pressure distortion inflow conditions were conducted using the body force model. The results show that, unlike the design point and non-design point at which the fan operates normally, the high radius region of the fan is in the “turbine mode” while the low radius region is in the “compressor mode” under windmilling conditions. The different effects on the longitudinal vortex in the two regions deepen and alleviate the circumferential distortion, respectively. There are strong circumferential and radial pressure gradients at the junction of the distortion-affected zone and the non-distortion-affected zone, adding additional mixing losses.

Influence of the rotation characteristics of the controllable speed casing on the flow stability of a high-load compressor stage

The controllable speed casing (CSC) has been confirmed as a novel type of casing treatment technology that is capable of designing the original fully fixed casing structure into two parts, i.e., the rotatable ring section and the stationary ring section. A range of circumferential additional forces will be imposed on the tip region of the rotor by regulating the rotation direction and rotation speed of the rotatable ring section, such that different effects will be exerted on the aerodynamic performance and flow stability at the compressor stage. As indicated by the results of this study, when the rotation direction of the CSC was opposite to that of the rotor, the relative increase of the stable operating margin changed from positive to negative when the casing rotation speed was increased and accounted for 60% of the rotor speed. When the CSC exhibited the identical rotation direction to the rotor, the stability turned out to be more significantly increased. With the decrease of circumferential pressure gradient in the tip region, the original driving force of the tip leakage flow reduces, and the flow capacity was increased. The stable operating margin was increased with the rise of the rotation speed of the CSC, and the stable operating margin was increased by up to 45.44% at casing rotation speed accounting for 100% of the rotor speed.

Adjoint-Based Aerodynamic Design Optimization and Drag Reduction Analysis of a Military Transport Aircraft Afterbody

Based on the adjoint method, the afterbody of a military transport aircraft was optimized and designed to meet engineering constraints under real flight conditions. Guidance for the key design parameters of the afterbody of the military transport aircraft is given. The vortex dynamics and boundary layer extraction methods were used to analyze the optimization results of military transport aircraft. It was found that, upstream of the vortex shedding point, the circumferential accumulation process of the vorticity is weakened. The position of the vortex shedding and the appearance of the saddle line are delayed by reducing the circumferential inverse pressure gradient and the intensity of the crossflow. The afterbody vortex system of the optimized configuration is further away from the surface. Meanwhile, the distance between the counter-rotated vortex decreases, and the upwashing speed of the vortex core is smaller. Therefore, vortex-induced drag is reduced. Finally, compared with the initial configuration, the optimized configuration has a relative drag reduction of 23.2%.

Influence of Wake Intensity on the Unsteady Flow Characteristics of the Integrated Aggressive Interturbine Duct

The interturbine transition duct (ITD), located between the high-pressure (HP) and low-pressure (LP) turbines of aeroengines, tends to be designed as an aggressive ITD integrated with wide-chord struts to meet the requirements of civil aeroengines for high bypass ratios and thrust–weight ratios. This paper presents a detailed unsteady numerical investigation of the effects of the HP rotor trailing-edge radius on the unsteady flow characteristics in the integrated aggressive interturbine transition duct (AITD), including the transport and dissipation of HP rotor wakes, the control mechanism of HP rotor wakes on flow separation and the influence of wake parameters. A sweeping rod, with a nondimensional diameter ranging from d/s = 0.056~0.143 (based on the pitch (s) of wide-chord struts at the midspan) and a reduced frequency (f) of 1.07, is used to simulate the HP rotor wake to decouple its influence from other secondary flows. Using the k-ω SST turbulence model and gamma–theta transition model, a structured grid with 6.3 million nodes can achieve similar global results. The wake in the lower part of the AITD channel dissipates rapidly because of the stretching between its own circumferential motion and the radial upward secondary flow, especially for a small d/s. Only the residual wake in the upper part can reach wide-chord struts in the case with large d/s. A sweeping rod with a large d/s can reduce the radial pressure gradient in the AITD, inhibit the internal secondary flow to a certain extent, reduce the dissipation rate of the wake, enhance its suppression effect on flow separation on a wide-chord strut, and decrease the flow loss. However, the wake can also enhance the passage vortex due to the increasing circumferential pressure gradient in the wide-chord strut channel, resulting in increasing blade profile loss. In the scope of this study, the aerodynamic gain of the wake is still not enough to compensate for its loss increment (including its own dissipation loss). Therefore, selecting a small trailing-edge radius of the HP rotor is conducive to improving the aerodynamic performance of the integrated AITD.

Cross-flow transition model predictions of hypersonic transition research vehicle

Hypersonic transition research vehicle (HyTRV) is a customized lifting body for studying hypersonic three-dimensional boundary-layer transition on a complex geometry. In this study, the transition model of an HyTRV lifting body at relatively high angles of attacks is investigated. A hypersonic cross-flow transition correlation is embedded in the γ-Re˜θt transition model on the Chant2.0 computing platform, called the C-γ-Reθ transition model. The transition model is validated in a series of typical three-dimensional aerodynamic configurations. HyTRV's flow characteristics and transition phenomenon are investigated. The model predictions are in good agreement with the stability analysis results and shockwave wind tunnel flow visualizations, exhibiting the model potential for hypersonic cross-flow transition prediction. The results of wind tunnel experiments, stability analysis, and model predictions confirm that transition regions on the upper/lower surface of HyTRV are dominated by cross-flow instabilities. With an increase in the angle of attack, the circumferential pressure gradient on the lower surface decreases and cross-flow intensity reduces; the transition onset location recesses until finally disappears. 2–3 pairs of isolated cross-flow transition zones exist on the surface of the HyTRV at a low angle of attack, which guarantees the HyTRV to be a typical validation case for cross-flow transition modeling.

An experimental study on turbine vane Leading-Edge film cooling with deposition

• Impacts of particle deposition on film cooling at turbine leading-edges are highlighted. • Early-stage deposition effects are experimentally simulated by injecting molten wax droplets. • Dynamic deposition process is documented by a transient infrared thermography technique. • Effects of deposit thickness on leading-edge film cooling are coupled with coolant blowing ratios. Deposition of particles ingested into the hot turbine component of an aircraft engine could lead to degradation of turbine cooling performance and thus reduces service lifetime. It is thereby critical to gain basic knowledge of the deposition effects on the turbine cooling, of which cooling for the vane leading edge, where hot gas stagnates, is intricate. In this work, liquid wax is air-atomized into micro droplets to simulate a particle-laden flow condition where a vane leading-edge model cooled by five rows of staggered holes is immersed. To take circumferential pressure gradient at the leading edge of a real turbine vane into consideration, a rectangular obstacle is designed ahead of the test model and is shifted towards one side to form pressure and suction sides around the leading edge. Particularly, cooling effectiveness is measured by using an infrared thermography (IR) technique at an early-stage deposition for blowing ratios of 0.5–1.6. Additionally, an array of real-time IR images of surface temperatures is gathered to document the dynamic changings of film cooling with the accumulation of deposits on the leading edge for a longer exposure. Comparisons between the deposition and non-deposition cases reveal that for the incipient deposition, the deposits are helpful to improve cooling effectiveness except for the optimum blowing ratio of 0.8 at which film cooling on the non-deposition model has the best performance. Because of the high-pressure surroundings, the hole rows on the pressure side are more sensitive to deposition, where increases of 0.05–0.15 in cooling effectiveness are found due to deposition. Moreover, dynamic deposition shows that the heavy deposits and heating effects of the warm particles could generate a reduction of cooling effectiveness up to 76% for a longer exposure. The added value of this study lies in providing a better understanding of the deposition effects on film cooling at the turbine vane leading edge, which is helpful to design a robust cooling scheme for this region to withstand the deposits.

Performance Characteristics and Energy Loss Analyses of a High-Speed Centrifugal Pump with Straight Blades

Along with the rapid growth of cutting-edge petrochemical technology and the pressing demand for efficiency improvement, evaluation of the performance characteristics of high-speed pump is becoming increasingly important. In this paper, numerical simulation is presented on the flow instability of a 16 straight-blade high-speed centrifugal pump with flow rate of 3 m3/h and rotating speed of 8500 rpm. Combined with the analysis of flow stability, the entropy production method is introduced to evaluate regions of high mechanical energy loss and its distribution at different flow rates. Results show that approximately 96% of the energy loss of the pump is produced in the volute, gap, and front and back chambers. Large energy loss is observed near the trailing edge of the blade and volute tongue, which are caused by the small region including both the high and low pressure gradients and large momentum exchange by the flow separation, respectively. Moreover, the rotor–stator interaction causes much energy loss at the wall of the volute and front and back chambers. Owing to the circumferential pressure gradient and the 90° leading edge of the straight blade, the fluid tends to form counter-rotating recirculation vortices. The large number of blades narrows the passage and limits the formation of large vortices in flow channels, thus the backflow phenomenon seems not to worsen with the rise of flow rates. Hence, the entropy production in most of the flow parts are insensitive to flow rates.

Flowfield of a helicopter submerged inlet with power output shaft

The submerged inlet is an attractive configuration for advanced helicopters due to its high stealth performance and low external drag. In this paper, a submerged inlet, integrated with a ROBIN helicopter fuselage and a simplified power output shaft, is experimentally and numerically investigated to obtain the basic flow characteristics under a freestream velocity of 23.6 m/s. The results indicate that the pylon ahead of the inlet induces a horseshoe vortex. Though the vortex is ingested into the inlet, it has little effect on the internal flows and can be neglected. When the airflow enters into the inlet, it interacts with the shaft with a large incidence angle, yielding a vortex pair. At the leeside of the shaft, the two side flows of the shaft impinge at the center plane, generating a local high-pressure region at the azimuthal angle of 180°, which forces the boundary layer to roll up a counter-rotating vortex pair. In addition, the airflow adjacent to the cowl lip accelerates rapidly, resulting in a local low-pressure region at the azimuthal angle of 0°. Therefore, the inlet duct has a strong circumferential pressure gradient, which originates from an azimuthal angle of 180° to 0° and induces a vortex pair at the azimuthal angle of 0°. The three vortex pairs are the main origins of the distortion at the duct exit plane, among which the one near the cowl lip with the azimuthal angle of 0° plays the dominant role. Additionally, as the velocity ratio increases from 3.9 to 5.5, the circumferential pressure gradient and the cowl lip vortex get intensified, which causes that the total-pressure recovery coefficient drops by 0.5% and the distortion index increases by 28%. A submerged inlet, integrated with a ROBIN helicopter fuselage and a simplified power output shaft, is experimentally and numerically investigated. Three vortex pairs, which locate at the azimuthal angle of 0°, the leeside of the shaft, and 180° of the inlet surface, are the main origins of the distortion of the inlet, among which the one near the cowl lip with the azimuthal angle of 0° plays the dominant role. As the velocity ratio increases, the circumferential pressure gradient gets intensified, leading to stronger vortex pairs.

Effect of an oscillating time-dependent pressure gradient on Dean flow: transient solution

BackgroundNavier-Stokes and continuity equations are utilized to simulate fully developed laminar Dean flow with an oscillating time-dependent pressure gradient. These equations are solved analytically with the appropriate boundary and initial conditions in terms of Laplace domain and inverted to time domain using a numerical inversion technique known as Riemann-Sum Approximation (RSA). The flow is assumed to be triggered by the applied circumferential pressure gradient (azimuthal pressure gradient) and the oscillating time-dependent pressure gradient. The influence of the various flow parameters on the flow formation are depicted graphically. Comparisons with previously established result has been made as a limit case when the frequency of the oscillation is taken as 0 (ω = 0). ResultsIt was revealed that maintaining the frequency of oscillation, the velocity and skin frictions can be made increasing functions of time. An increasing frequency of the oscillating time-dependent pressure gradient and relatively a small amount of time is desirable for a decreasing velocity and skin frictions. The fluid vorticity decreases with further distance towards the outer cylinder as time passes.ConclusionFindings confirm that increasing the frequency of oscillation weakens the fluid velocity and the drag on both walls of the cylinders.

Surface-pressure pattern of separating flows over inclined slender bodies

Surface-pressure measurements and literature review of crossflows over inclined slender bodies at high Reynolds numbers (4.2 × 106 to 17 × 106, based on body length and freestream velocity) show that the constructed pressure maps have a distinct pattern, which relates to large-scale separation. The separation envelope is defined by a pair of primary and secondary surface-bifurcation lines as described by Lee [“Longitudinal development of flow-separation lines on slender bodies in translation,” J. Fluid Mech. 837, 627–639 (2018)]. The present study shows that the separation envelope falls in a “plateau” region in the circumferential distribution of the mean pressure. In the distribution of circumferential pressure gradient, there are two local maxima, where one is located on the windward side of the separation envelope and the other is located on the leeward side. Pattern tracking identifies that the locus of maximum circumferential pressure gradient resembles the locus of separation, but is located further upstream along the slender body.

Numerical Analysis of High-Altitude Inlet Air on Boundary Layer Flow Loss in an Aero-Engine Compressor

A numerical analysis is performed to explore the high altitude and high Mach flight on the effect of wall boundary layer loss in the compressor. The accuracy for solution results by the application of the similarity criterion and parameter definition of the air inlet is compared with the existing experimental test result. The results indicate that the radial adverse pressure gradient in the rotor domain gradually increases along the span direction and decreases as flight Mach number increases; meanwhile, the circumferential adverse pressure gradient on the pressure side of the rotor blade is correspondingly larger and less than that on the suction side. In particular, the entropy increase along the streamwise shows a decreasing trend and an increasing trend inside the hub and shroud wall boundary layers, respectively. At 2.1 Ma, the entropy increase in the rotor domains enhances by 24.36–27.80% inside the shroud boundary layer, relative to the hub boundary layer; however, it decreases by 0.97–8.54% in the stator domain. With the increase in flight Mach number from 2.1 to 3.4, the average entropy increase reductions in the rotor domain decrease by 18.99–24.97% within the hub boundary layer and 5.71–8.1% within the shroud boundary layer. In the stator domain, it drops by 18.45–9.03% inside the hub boundary layer and 6.88–8.67% inside the shroud boundary layer. It was therefore found that, as Mach number increases from 2.1 to 3.4, the entropy increase reduction is larger inside the hub boundary layer than inside the shroud boundary layer.

An improved prediction model for corner stall in axial compressors with dihedral effect

Corner stall predictions are important and difficult in axial compressors. However, all of the prediction models have proved to be ineffective for advanced compressor blades, which tend to use the combined sweep and dihedral. As for the prediction parameter DL, although it effectively modeled the effects of the adverse pressure gradient and secondary flow, it failed to predict the corner stall of curved blades because the model failed to consider the intersection of the boundary layer at the corner region. In this paper, the shape factor gradient Ψ of the boundary layer at the corner region was investigated by numerically studying specially shaped expansion pipes under different adverse pressure gradients. The improved prediction parameter DJ was presented based on the model of Ψ and the circumferential pressure gradient ξ. A comparison of the critical range of the prediction parameters DL and DJ was investigated using the NACA65 cascade database, which was established by a numerical method. Then, the stall criterion was validated according to the experimental results of various test facilities with different blade geometries and experimental conditions. The results show that the improved prediction parameter is able to predict the corner separation/stall flows and is in good agreement with the experimental results for axial compressors with three-dimensional designed blades.

Sensitivity investigation of a subsonic cascade performance to geometric deviations based on statistical method

This paper has established an evaluation system for the sensitivity analysis of cascade/compressor aerodynamic performance to geometric deviations. The system includes blade profile parameterization with non-uniform rational B-spline (NURBS), uniform Latin hypercube sampling in association with Monte Carlo algorithm and statistical investigation methods. The sensitivity of a subsonic cascade performance to its geometric deviations is investigated under different typical inflow conditions. Results show that there is an approximately linear relationship between the geometric deviations and the aerodynamic parameters. The total pressure loss coefficient is the most sensitive to the leading edge profile deviation at i=2°. However the suction profile deviation and the leading edge profile deviation dominate the total pressure loss when the incidences are i = −4° and i = −8°. The blocking effect of the increased leading edge profile moves forward the separation point of the boundary layer. Therefore the scale of the separation vortex is enlarged, the total pressure loss is increased at i = 2° condition. Regarding i = −8°, the enlarged suction profile deviation reduces the circumferential static pressure gradient, resulting in an increase of the scale of separation vortex on the pressure side. The work is helpful to robust design of compressor blade profile and manufacturing & maintenance techniques of compressor blades.

Development of a One-Dimensional Model for the Prediction of Leakage Flows in Rotating Cavities Under Non-Uniform Tangential Pressure Distribution

Regenerative pumps are characterized by a low specific speed that place them between rotary positive displacement pumps and purely radial centrifugal pumps. They are interesting for many industrial applications since, for a given flow rate and a specified head, they allow for a reduced size and can operate at a lower rotational speed with respect to purely radial pumps. The complexity of the flow within regenerative machines makes the theoretical performance estimation a challenging task. The prediction of the leakage flow rate between the rotating and the static disks has the greatest impact on the prediction of global performance. All the classical approaches to the disk clearance problem assume that there is no relevant circumferential pressure gradient. In the present case, the flow develops along the tangential direction and the pressure gradient is intrinsically non-zero. The aim of the present work is to develop a reliable approach for the prediction of leakage flows in regenerative pumps. A preliminary numerical simulation on a virtual model of a regenerative pump where the disk clearance is part of the control volume has been performed for three different clearance aspect ratios. The outcome of that campaign allowed the authors to determine the behavior of the flow in the cavity and choose correctly the baseline hypotheses for a mathematical-physical method for the prediction of leakage flows. The method assumes that the flow inside of the disk clearance is two-dimensional and can be decomposed into several stream-tubes. Energy balance is performed for each tube, thus generating a system that can be solved numerically. The new methodology was tuned using data obtained from the numerical simulation. After that, the methodology was integrated into an existing one-dimensional code called DART (developed at the University of Florence in cooperation with Pierburg Pump Technology Italy S.p.A.) and the new algorithm was verified using available numerical and experimental data. It is here demonstrated that an appropriate calibration of the leakage flow model allows for an improved reliability of the one-dimensional code.

URANS investigation of the interaction between the free surface and a shallowly submerged underwater vehicle at steady drift

An axisymmetric underwater vehicle (UV) at a steady drift angle experiences the complex three-dimensional crossflow separation. This separation arises from the unfavorable circumferential pressure gradient developed from the windward side toward the leeward side. As is well known, the separated flow in the leeward side gives rise to the formation of a pair of vortices, which affects considerably the forces and moments acting on the UV. In this regard, the main purpose of the present study is to evaluate the role of the leeward vortical flow structure in the hydrodynamic behavior of a shallowly submerged UV at a moderate drift angle traveling beneath the free surface. Accordingly, the static drift tests are performed on the SUBOFF UV model using URANS equations coupled with a Reynolds stress turbulence model. The simulations are carried out in the commercial code STARCCM+ at a constant advance velocity based on Froude number equal to Fn = 0.512 over submergence depths and drift angles ranging from h = 1.1D to h = ∞ and from β = 0 to β = 18.11°, respectively. The validation of the numerical model is partially conducted by using the existing experimental data of the forces and moment acting on the totally submerged bare hull model. Significant interaction between the low-pressure region created by the leeward vortical flow structure and the free surface is observed. As a result of this interaction, the leeward vortical flow structure appears to be largely responsible for the behavior of the forces and moments exerted on a shallowly submerged UV at steady drift.

RANS feasibility study of using roughness to mimic transition strip effect on the crossflowseparation over a 6:1 prolate-spheroid

An axisymmetric body at incidence experiences the three-dimensional crossflow separation. This separation is attributed to the adverse circumferential pressure gradient. However, the separation pattern is also dependent upon the structure of the boundary layer. In this regard, utilization of transition strip devices in experiments on axisymmetric bodies may modify this structure, and consequently the crossflow separation pattern. The main objective of the present research is to mimic numerically the transition strip effect on the crossflow separation over a 6:1 prolate-spheroid up to α = 30° incidence and ReL = 4.2×106. However, to avoid direct modeling of the strip, which would increase the computational cost, an attempt was made to add roughness over the body surface. To estimate the roughness that simulates closely the transition strip effect, three different roughness values were considered. The numerical model is based on RANS and a Reynolds stress turbulence model implemented in STARCCM+. The simulations have been evaluated based on the local and global variables and validated against the available experimental data. The results demonstrate the effectiveness of using a proper roughness value to mimic the transition strip effect. They also show the importance of modeling the transition strip effect, which is normally not considered, to capture the crossflow separation pattern.

Entrance Aspect Ratio Effect on S-Duct Inlet Performance at High-Subsonic Flow

This paper reports the internal performance evaluation of S-duct diffusers with different entrance aspect ratios as part of a parametric investigation of a generic S-duct inlet. The generic S-duct diffusers studied had a rectangular entrance (aspect ratios of 1.5 and 2.0) transitioning S-duct diffuser in high-subsonic (Mach number > 0.8) flow. The test section was manufactured using rapid prototyping to facilitate the parametric investigation of the geometry. Streamwise static pressure and exit-plane total pressure were measured in a test-rig using surface pressure taps and a five-probe rotating rake, respectively. The baseline and a variant were simulated through computational fluid dynamics (CFD). The investigation indicated the presence of streamwise and circumferential pressure gradients leading to a three-dimensional flow in the S-duct diffuser and to distortion at the exit plane. The static pressure recovery increased for the diffuser with the higher aspect ratio. Total pressure losses and circumferential and radial distortions at the exit plane were higher than that of the podded nacelle type of inlet. An increase in the total pressure recovery was observed for the increase in the aspect ratio for the baseline area ratio (1.57) S-ducts, but without a clear trend for the other area ratio (1.8) ducts. The work represents the development of a database on the performance of a particular type of generic inlet. This database will be useful for predicting the performance of aero-engines and air vehicles in high-subsonic flight.