Wall Static Pressure Measurements Research Articles

Effect of Dissolved Carbon Dioxide on Cavitation in a Circular Orifice

In this work, we investigate the effect of dissolved gas concentration on cavitation inception and cavitation development in a transparent sharp-edged orifice, similar to that previously analyzed by Nurick in the context of liquid injectors. The working liquid is water, and carbon dioxide is employed as a non-condensable dissolved gas. Cavitation inception points are determined for different dissolved gas concentration levels by measuring wall-static pressures just downstream of the orifice contraction and visually observing the onset of a localized (vapor) bubble cloud formation and collapse. Cavitation onset correlates with a plateau in wall-static pressure measurements as a function of a cavitation number. An increase in the amount of dissolved carbon dioxide is found to increase the cavitation number at which the onset of cavitation occurs. The transition from cloud cavitation to extended-sheet or full cavitation along the entire orifice length occurs suddenly and is shifted to higher cavitation numbers with increasing dissolved gas content. Volume flow rate measurements are performed to determine the change in the discharge coefficient with the cavitation number and dissolved gas content for the investigated cases. CFD analyses are carried out based on the cavitation model by Zwart et al. and the model by Yang et al. to account for non-condensable gases. Discharge coefficients obtained from the numerical simulations are in good agreement with experimental values, although they are slightly higher in the cavitating case. The earlier onset of fluid cavitation (i.e., cavitation inception at higher cavitation numbers) with increasing dissolved carbon dioxide content is not predicted using the employed numerical model.

Validation of Rotating Detonation Combustor Computational Fluid Dynamics Simulations for Predicting Unsteady Supersonic–Subsonic Flow Field at the Exit

Abstract Rotating detonation combustors (RDCs) have gained increased interest for integration with power-generating gas turbines due to the potential to increase thermal efficiency. The unsteady flow field exiting the RDC is fundamentally different compared to traditional swirl-stabilized combustors. Successful integration of RDC with gas turbines will depend on the ability to properly condition the unsteady flow to achieve performance levels comparable to swirl-stabilized combustors. RDC simulations require significant computational resources due to the small spatial and temporal time scales required to resolve the detonation phenomenon. Furthermore, traditional steady-state computational fluid dynamics (CFD) analyses are not possible for RDC simulations. The present study develops and validates a computationally efficient approach for predicting unsteady flow fields exiting the combustor using 2D, transient reacting CFD with periodic boundary conditions in the combustor and a downstream plenum. Validation is performed by comparing the CFD results to various experimental measurements: (i) wave speed obtained from high-speed ion probe and dynamic pressure data, (ii) average wall static pressure measurements, and (iii) time-resolved particle image velocimetry (PIV) at 100 kHz at the RDC exit. Results indicate good agreement between CFD and experiments with respect to velocity field exiting the RDC, detonation wave speed, and static pressure distribution.

Assessment of Pseudoshock Models Against Experiment in a Low-Aspect-Ratio Isolator

A highly confined shock train is investigated in a direct-connect isolator facility with a Mach 2 inflow and a constant-area low-aspect-ratio rectangular test section. High-speed schlieren imaging, wall static pressure measurements, surface oil-flow visualization, and particle image velocimetry from this isolator are synthesized into a three-dimensional schematic of the shock train structure. Against this, the prevailing pseudoshock models in the literature are assessed to evaluate the validity of their underlying assumptions. None of the prevailing pseudoshock models are found to simultaneously model the pressure and Mach number profiles, indicating a gap in the model formation and underlying assumptions when applied to the experimental isolator of interest. The presence of distortion in the isolator flowfield, such as a wall-bounded vortex, is found to skew the structure of the shock train, altering the strength and distribution of the compressive pressure gradient. It is further observed that the separated flow morphology surrounding the shock train is not monolithic, as is typically assumed, adjusting the balance of compressive forces within the shock cells. These findings lead to the conclusion that existing flux-conserved modeling approaches require modification to be effective in distorted and highly confined cases, including closure models that capture the three-dimensional distorted structure of the approach flow and its evolution along the shock train.

Heat transfer characteristics of jet impingement onto the concave surface of a cone

ABSTRACT Heat transfer coefficient measurements for jet impingement onto the concave surface of a cone with apex angles equal to 30° and 70° are reported in this study. Static pressure measurements along the cone surface are also reported which aid in explaining the results from the heat transfer measurements. The conical geometry finds application for the heating of aircraft gas turbine engine nose bullet region which could experience ice formation whenever aircraft operates at higher altitudes. The concave surface of the cone is heated by impinging hot air tapped from the compressor. The heat transfer coefficients, which are significantly different compared to flat surface impingement, and are scarcely available, are reported in this study. The influence of the jet axis coincident with the cone axis and impinging into apex region was studied first. The methodology where a thin stainless-steel foil is heated with a known heat flux by passing electric current through it and then measuring the detailed surface temperature using an infrared thermal camera, was used to compute the wall Nusselt distribution. The Nusselt number variation is presented as a function of nondimensional apex to nozzle spacing (L/D) and the nondimensional axis displacement (O/D). Three different diameter (D equal to 10 mm,14 mm and 20 mm) pipes were used for the concentric impingement case and the Nusselt number distribution was observed not depend on the injection diameter. The jet Reynolds number was varied between 25000 and 82000. The peak Nusselt number values are similar to those for flat plate impingement and do not occur at the cone apex, but at a nondimensional distance between 3.2 and 3.9 for 4.25 < L/D < 10.25 for the 30° cone. The peak Nusselt number values increase by almost 30% with decrease in the L/D from 10.2 to 3.0. Wall static pressure measurements indicated that the peak Nusselt number values occur in regions where the interaction between the incoming and outgoing streams is strongest – these regions occur at distances downstream from the stagnation region toward cone exit. The peak heat transfer coefficient location could not be obtained for the 70° cone, but the measured maximum values are higher by 30–50% compared to those for the 30° case. The jet is able to penetrate further toward the apex of the cone for 70° cone compared to 30° cone shifting the peak heat transfer locations closer to the apex for larger apex angle case. There is insignificant influence of jet diameter on Nusselt number variation for the three diameter values investigated in this study. The influence of displacing the jet axis by a distance equal to half the jet diameter in a direction perpendicular to the cone axis was also studied but detailed measurements were possible only for the 30° cone. Two semi-circular regions, one close to the impingement tube and the other away from the tube, displaying similar heat transfer coefficient behavior were observed. A peak in the Nusselt number occurs in each semi-circular region and the magnitude of the peaks are nearly same and they lie on either side of the peak location observed for the concentric impingement case when the flow geometric parameters are similar for the two cases. The measurements for the 70° cone exhibited all features of the 30° cone but the maximum values were higher by nearly 40% for this case.

Simultaneous measurement of pressure and temperature in a supersonic ejector using FBG sensors

In this work, we have demonstrated the use of fiber Bragg grating (FBG) sensors for simultaneous measurement of wall static pressure and temperature in a supersonic ejector. Supersonic ejectors are ground-based high-speed aerodynamic test facilities characterized by harsh conditions, such as high pressure and temperature gradients. An FBG-based sensor setup was developed consisting of a pressure measuring bare FBG and a specially designed pressure-insensitive FBG temperature probe that can be mounted on the wall of the supersonic ejector. The FBG temperature probe was used for temperature measurement as well as temperature compensation of the pressure measuring FBG sensor. Wall static pressure measurements in the supersonic ejector were carried out at different tank pressures and Mach number flows. The FBG pressure measurements were validated with those of standard piezoresistive-based sensor measurements. Both responses were found to match closely, with FBG sensors having a faster response time and higher pressure resolution. Fluid structure interaction simulation was carried out in Comsol Multiphysics to understand the interaction of high-speed turbulent flow with FBG sensor. The FBG strain profile due to flow-induced stress and its dependence on flow pressure was studied. A detailed analysis of the effect of preceding fiber length on FBG pressure measurement was carried out. FBG sensors, due to their miniature size, ability to withstand harsh environments and multi-parameter sensing capability, can be used in ground-based aerodynamic test facilities with minimal intrusion into the flow.

Enhancement of blowout limit in a Mach 2.92 cavity-based scramjet combustor by a gliding arc discharge

Plasma-assisted combustion (PAC) has shown excellent performance in the ignition and combustion enhancement of scramjet combustors. A multi-channel gliding arc (MCGA) plasma placed on the sidewall surface of a scramjet combustor was employed to enhance combustion near the flame blowout limit (BL) in a C2H4-fueled and cavity-based model scramjet combustor. Wall static pressure measurements, simultaneous 10 kHz CH* chemiluminescence imaging from the top and side views, optical emission spectroscopy, and discharge waveform measurements were used to elucidate the combustion physics throughout the PAC events. For the current cavity configuration and fueling condition, a general estimate of the flame BL in a rich-fuel environment was provided, in which a fuel flow rate (FFR) was 7.5 g/s. Once the FFR was exceeded, the flame was unstable and blown out, whereas the MCGA plasma demonstrated an excellent flame stabilization ability. The flame BL of the scramjet combustor in the presence of the MCGA plasma has an increase of ∼29%. Two PAC processes, including enhancement mode and re-ignition mode, were distinguished to play a vital role in extending the flame BL. The two PAC processes showed that the spark-type discharge generated by the MCGA could ignite the local fuel-rich mixture and increase the flame BL. The local ignition is dominated by the continuous accumulation of the massive flame kernels with the local fuel-rich mixture, as well as more heat and species exchanges.

Effect of Flap Deflection on Single-Expansion-Ramp Nozzles Performance at Different Pressure Ratios

Experiments have been carried out to investigate the various performance characteristics of single-expansion-ramp nozzle with straight and deflected flap conditions for two different flap lengths. Experiments were performed at different nozzle pressure ratios (NPRs) ranging between 1.5 and 9. Wall static pressure measurements were made to estimate the various performance parameters, such as surface axial force, normal force, overall axial force, pitching moment, and thrust vectoring angles. Schlieren flow visualization technique was adopted to understand the internal shock-induced flow separation and shock patterns inside the nozzle under various NPRs. Flap deflection and NPR control the shock location through area variation, and thus the pressure distribution on the walls. At low NPRs, the internal drag force generated by deflected flap is lesser than that by straight flap, whereas at high NPRs the straight flap produces significantly higher surface axial thrust over the deflected flap. Behavior of normal force and pitching moment was affected by two factors: flap length and shock location. The straight flap shows significantly higher normal force than deflected flap. Hence, the thrust vectoring was better for straight flap nozzles compared to deflected flap nozzles, over the full range of NPRs tested. On the other hand, the gross thrust force performance is better for the straight flap for low NPRs and for the deflected flap for high NPRs. Overall, at high speeds, longer deflected flap appears to give better performance considering all the parameters, but will lead to heavier nozzle.

A passive loss reduction method of square-to-square sudden expansions

Mechanical ventilation in edifices is a key contributor to the significant energy consumption of the building industry. Therefore, rational energy management in ventilation systems plays a major role in promoting the energy efficiency of buildings. The present investigation aims at the moderation of fluid mechanical losses – thus the overall energy expenditure of ventilation ducts – of a symmetric square-to-square sudden expansion, frequently found in ventilation systems. A simple passive flow control method with moderate manufacturing costs was proposed, and its efficiency was examined experimentally. The loss reducing appendages were short rectangular sheets of adjustable angle, so-called miniflaps, placed at the step edge of the sudden expansion. Flow characteristics were examined for a fully developed turbulent flow at the inlet of the sudden expansion, for an area ratio of 2.78 and a miniflap length of 0.4 step heights. The investigated Reynolds number range was Re = (0.37–1.82)·105, being representative of ventilation systems. Fluid mechanical losses were determined from wall static pressure measurements. Velocity and turbulence intensity fields were mapped with the help of a laser Doppler anemometer. The miniflaps successfully reduced the loss coefficient of the sudden expansion through quick uniformization of the velocity profile. The loss coefficient was reduced by an absolute value of 0.07–0.11, i.e. a relative value of 14%–25%, for an optimum miniflap angle of α = 12° ± 1°. Reattachment length was also reduced from 11 to 7 step heights in the wall-normal and from 14 to 13 step heights in the diagonal plane. The loss reducing capacity together with the compactness of the miniflaps and the overall reduction of the affected zone downstream of the element presents a competitive advantage of the miniflaps in the building industry, which is often characterized by limited space allowance.

Computational fluid dynamic modelling of supersonic ejectors: comparison between 2D and 3D modelling

It is known that the global performances of ejector-based systems (viz., at the “global-scale”) depend on the local flow properties within the ejector (viz., at the “local-scale”). For this reason, reliable computational fluid-dynamics (CFD) approaches, to obtain a precise and an a-priori knowledge of the local flow phenomena, are of fundamental importance to support the deployment of innovative ejector-based systems. This communication contributes to the existing discussion by presenting a numerical study of the turbulent compressible flow in a supersonic ejector. In particular, this communication focuses on a precise knowledge gap: the comparison between 2D and 3D modelling approaches as well as density-based and pressure-based solvers. The different approaches have been compared and validated against literature data consisting in entrainment ratio and wall static pressure measurements. In conclusion, this paper is intended to provide guidelines for researchers dealing with the numerical simulation of ejectors.

Micro-Ramp Control for Shock Train Structure and Oscillation

Micro-ramps for shock train control were investigated experimentally in a Mach 1.85 supersonic isolator flow using three device heights. A mechanical flap was mounted downstream and linearly choked, forcing the shock train to travel upstream of the micro-ramp from the exit. Wall static pressure measurements along the primary and corner regions obtained a more complete pressure distribution. Detailed structures of the shock train were recorded using schlieren visualization with circular, horizontal, vertical, and color knife edges. The results show that the wake and wave structures generated by the micro-ramp control the shock train. The taller the device, the stronger the wake and wave structures, and the greater the control. The wake and wave structures shorten or eliminate the Mach stem in the shock train leading edge, causing modification of shocks downstream from pseudonormal to oblique, inhibiting the thickening of the boundary layer that results from shock impingement, slowing down the contraction of the supersonic core flow, and allowing more high total pressure fluid to develop downstream. The wake structure was found to promote the oscillation and upstream motion of the shock train. The wave structure can restrain shock train oscillation and upstream propagation of the downstream adverse pressure gradient, enhancing the resistance backpressure of the flowfield downstream of the micro-ramp.

Locating the Isolator Shock-Train Leading Edge with Limited Pressure Information

Real-time detection and control of the isolator shock-train leading edge (STLE) is important to the performance of high-speed air-breathing engines, such as dual-mode scramjets. Typically, the STLE location is determined using wall static-pressure measurements, but there are often restrictions on the placement and overall number of the pressure transducers, reducing the viability and accuracy of such approaches. To address these issues, we introduce the adaptive pressure profile (APP) method for estimating the STLE location. This method does not require extensive prior characterization of the isolator or engine model. Instead, it uses real-time pressure measurements from a small number of transducers to adaptively learn the isolator pressure profile and subsequently uses this deduced profile to estimate the STLE location in a data-driven manner. The APP method works well in situations with sparse transducer placement. It produces accurate estimates when the STLE location is 1) not bounded by two or more transducers or 2) between two transducers that are several isolator duct heights apart. We demonstrate the efficacy of the APP method using simulations and experimental data from direct-connect isolator models. This validation shows that the APP method is accurate and robust for different flow regimes, transducer configurations, and model geometries.

Flow topology and loss analysis of a square-to-square sudden expansion relevant to HVAC systems: A case study

Mechanical ventilation of buildings contributes significantly to the world's overall energy consumption, therefore, it is important to consider and control the fluid mechanics losses of air duct components. The current investigation aims to investigate the Borda-Carnot sudden expansion, which is often used in ventilation systems in place of diffusers when the flow cross-section needs to be increased. While a significant amount of research has been carried out on axisymmetric sudden expansions, expansions of square or rectangular cross-sections are seldom discussed in the literature. As rectangular air ducts are widely used in ventilation systems, a claim on the research of the flow topology and the loss characteristics of rectangular expansions is raised and is assessed in the present study. In a first step, the Borda-Carnot loss was examined for uniform and fully developed inlet velocity profiles by means of semi-empirical methods. It was concluded that the classic loss formula derived for uniform inlet underestimates the losses and may cause a 2–3% error for each sudden expansion in the calculation of losses in a ventilation system. In a second step, the flow details and the loss characteristics of a square-to-square sudden expansion of an area ratio of 2.78 were examined by experimental means. The investigated Reynolds number range was (0.36–1.8)·105, being representative of ventilation systems. Detailed velocity and turbulence-related measurements were carried out with a laser Doppler anemometer. The downstream effects of the square-to-square expansion were found to extend significantly farther downstream than that of an axisymmetric expansion, resulting in a 1.5–2.5 times larger flow reattachment length and a 1.5–2 times longer path to reach a fully developed state. The elongated downstream effect gains importance when there is a need to place control or measurement devices downstream of the expansion. As revealed by wall static pressure measurements, the semi-empirical formula derived for the Borda-Carnot loss for fully developed upstream flow approximates the losses more accurately than the classic formula derived for a uniform inlet. Therefore, a methodology has been proposed herein for calculating the Borda-Carnot loss in ventilation systems, incorporating the loss formula for fully developed inflow.

Experimental investigation of shock train behavior in a supersonic isolator

For a better understanding of the shock train structure and its dynamic oscillation features, wind tunnel experiments with linear and stepwise increase backpressure were conducted with supersonic isolator flows at Mach 1.85 and 2.70. High-frequency wall static pressure measurements were performed along the primary and corner regions to capture the behavior of the primary and corner shock train. The fine structures of the shock train were recorded using schlieren visualization with circular, horizontal, vertical, and color knife edges. The pressure results show that the shock train leading shock at Mach 2.70 is more three-dimensional. The flow field exhibits the following features near the corner: the pressure fluctuation amplitude is smaller, the shock train leading shock is closer to the upstream regions, and the oscillation of the shock train leading shock can propagate a longer distance downstream. Schlieren snapshots obtained using horizontal and vertical knife edges show shock train structures with alternating distributions of the vertical and horizontal density gradients. Further application of color knife edges clearly distinguishes these regions. The power spectra analysis of a series of schlieren snapshots was performed to characterize the dominant oscillation structures in the flow field with different visualization variables, and the motion relationship between the structures in the shock train was clarified based on the coherence and phase analyses of the schlieren images. The disturbance downstream first causes the movement of the shock train leading shock, and the closer the downstream shock is to the shock train leading shock, the earlier it moves.

Spectral signal quality of fast pressure sensitive paint measurements in turbulent shock-wave/boundary layer interactions

This study presents a critical investigation of the signal quality obtained using candidate fast pressure sensitive paint measurements in realistic, spectrally complex flows. The evaluations focus on characterizing the importance of the paint layer frequency response in conjunction with the fundamental signal-to-noise ratio limitations of the imaging system. Fast fluoro-isopropyl-butyl polymer-based paint is used primarily because of an identified characteristic attenuation pattern. Complementary high-speed pressure sensitive paint and wall static pressure measurements are performed in a turbulent, separated, scooped compression ramp shock-wave/boundary layer interaction at Mach 2.5 using a sampling rate of 4 kHz. The results reveal a systematic underprediction of the RMS pressure due to the limited frequency response of the paint. A simple frequency domain correction methodology utilizing the experimental data and a diffusion-based model of the paint response are employed to compensate for these effects. A method for determining the spectral signal-to-noise ratio of the imaging system is then presented, which is found to impose a major constraint on the upper limit of the resolvable pressure frequency. The correction is observed to provide a significant improvement in the agreement between the pressure sensitive paint and transducers, but only below the noise-dominated cutoff frequency. The relative significance of these factors is then investigated for polymer–ceramic-based, 8 kHz measurements of a planar compression ramp shock-wave/boundary layer interaction in a rectangular channel.

Experimental investigations on mixing characteristics in the critical regime of a low-area ratio supersonic ejector

Ejectors have many applications in aerospace and energy conversion technologies. An ejector is a passive device that pumps a secondary fluid by energy augmentation from a primary fluid. The performance of the ejector is primarily dependent on the complex gas dynamic interactions between the primary and secondary flows in a variable area duct. The maximum mass flow rate of the ejector in the critical flow regime is limited by choking of both the primary and secondary flows. This paper focuses on experimental investigations on the mixing characteristics of the ejector having an area ratio of 2 in the critical flow regime for a range of stagnation pressure ratios varying between 5.49 and 11.12 and a primary Mach number of 1.5, 2.0, and 2.5. The gas dynamic flow field is visualized using non-invasive high-speed schlieren and Mie-scattering techniques. Mie-scattering images are used to gain insight into the mixing characteristics and estimate the non-mixed length. The optical measurements are complimented with the wall static pressure measurements. The experimental performance data are compared with two well-established analytical models. The characterization of the mixing in the critical operating regime of the ejector using optical tools is being reported here for the first time. An important outcome of the paper is the observation of a significant increment in the non-mixed length to the tune of 55% increase in the critical flow regime in comparison to the mixed flow regime.

Experimental parametric studies on the performance and mixing characteristics of a low area ratio rectangular supersonic gaseous ejector by varying the secondary flow rate

A low area ratio rectangular supersonic gaseous ejector is subjected to parametric evaluation to calculate the performance parameters like stagnation pressure ratio, compression ratio, entrainment ratio and the mixing parameter known as non-mixed length for a wide range of operating conditions by varying the secondary flow rate. The operating conditions are achieved by varying the design Mach number of the primary flow nozzle, the total pressure of the primary flow and the secondary flow rate. Air is used as the working fluid in both the primary and secondary flow. The ejector is operated in the mixed regime. Mach number ratio is used as the non-dimensionalization parameter, and fully-expanded jet height is used as the scaling variable to collapse the huge set of obtained data for parametric studies. With variation in the secondary flow rate, staging is observed in the compression ratio, entrainment ratio, and also in the non-mixed length. Variation in the entrainment ratio and non-mixed length are observed to be linear, and it scales well with the fully-expanded jet height when there is a deficit in the secondary flow. Also, when there is no secondary flow, the non-mixed length is observed to be 80% lower in comparison with the case, where the secondary flow is uncontrolled. Schlieren visualization and wall static pressure measurements supplement the findings.

Distortion reconstruction in S-ducts from wall static pressure measurements

PurposeThis paper aims to report the attempts for predicting “on-the-fly” flow distortion in the engine entrance plane of a highly curved S-duct from wall static pressure measurements. Such a technology would be indispensable to trigger active flow control devices to mitigate the intense flow separations which occur in specific flight conditions.Design/methodology/approachEvaluation of different reconstruction algorithms is performed on the basis of data extracted from a Zonal Detached Eddy Simulation (ZDES) of a well-documented S-Duct (Garnier et al., AIAA J., 2015). Contrary to RANS methods, such a hybrid approach makes unsteady distortions available, which are necessary information for reconstruction algorithm assessment.FindingsThe best reconstruction accuracy is obtained with the artificial neural network (ANN) but the improvement compared to the classical linear stochastic estimation (LSE) is minor. The different inlet distortion coefficients are not reconstructed with the same accuracy. KA2 coefficient is finally identified as the more suited for activation of the control device.Originality/valueLSE and its second-order variant (quadratic stochastic estimation [QSE]) are applied for reconstructing instantaneous stagnation pressure in the flow field. The potential improvement of an algorithm based on an ANN is also evaluated. The statistical link between the wall sensors and 40-Kulite rake sensors are carefully discussed and the accuracy of the reconstruction of the most used distortion coefficients (DC60, RDI, CDI and KA2) is quantified for each estimation technique.

Flow Characterization for a Shallow Single Serpentine Nozzle with Aft Deck

An experimental investigation is carried out to characterize the flow through a shallow single serpentine curved convergent nozzle with an aft deck under fully expanded and underexpanded conditions with nozzle pressure ratios ranging from 1.1 to 2.2 for cold flows by wall static pressure measurements and flow visualization of the exit jet. The study revealed initial deceleration of the flow due to the turning of the streamlines caused by the nozzle curvature. Asymmetric coefficient of pressure variation in the cross-stream direction across various nozzle pressure ratios was observed, which is caused by secondary flow created by wall curvature and transition. The acceleration of the core flow due to geometric transition, as well as the additional nozzle length, causes early transition to sonic and supersonic flows, resulting in the formation of shock cells at relatively lower Mach number in the absence of the aft deck, as can be seen from the schlieren images. However, the presence of the constant area extension section reduced the flow asymmetry. Furthermore, the presence of the aft deck prevented the formation of barrel shocks and is expected to prevent the formation Mach disk(s) at higher nozzle pressure ratios. Thus, the curvature, transition, and aft deck complemented each other.

Computational fluid-dynamics modeling of supersonic ejectors: Screening of turbulence modeling approaches

Abstract Ejector refrigeration has not been able to penetrate the market because of the low coefficient of performance and because of the high influence of ejector performance on the efficiency of the refrigeration system. Improving the performance of ejector refrigeration systems relies on the understanding of the fluid dynamic phenomena, which can be obtained through computational fluid-dynamics approaches. Unfortunately, no agreement has been reached yet on the closures for the turbulence modeling in the Reynolds Averaged Navier-Stokes (RANS) approach. This paper contributes to the existing discussion by presenting a numerical study of the turbulent compressible fluid in a supersonic ejector. Seven RANS turbulence closures have been compared and, in addition, the turbulence models were tested under different near-wall modeling options in order to investigate the wall treatment effect on the numerical results. The numerical results have been validated by literature data consisting in entrainment ratio and wall static pressure measurements, for different ejector geometries and operating conditions. As a result, the k-ω SST model shows better performance in terms of global and local flow phenomena predictions for the different ejector geometries and operating conditions: (a) on the global point of view, the entrainment ratio is well predicted with a maximum relative error equal to about 10%, (b) on the local point of view, the shock wave position, the pressure recovery and the wall static pressure values are well predicted. The results, taking into account previous numerical studies, suggest the use of the k-ω SST model to simulate ejector fluid dynamics.

Experimental investigation of primary and corner shock boundary layer interactions at mild back pressure ratios

Unstart of rectangular inlets occurs as a result of interactions between shock-induced separation units along the floor/ceiling, corner, and sidewalls. While a significant body of literature exists regarding the individual flow interactions at the inlet floor/ceiling (called primary separation) and sidewalls, limited efforts have focused on the mean and dynamic features of the corner separation. Experiments are conducted to investigate primary and corner shock boundary layer interactions (SBLI) with the objectives of elucidating the flow interactions that occur in the corner, and characterizing the interaction between the corner and primary separation units at mild back pressure ratios. Surface streakline flow visualization and high-frequency wall static pressure measurements are performed along the centerline and corner regions of shock-induced flow separation generated by a 12° compression ramp in a Mach 2.5 flow. Sidewall fences that extend upstream of the leading edge of the flat plate generate corner separation of adequate size to determine the mean flow structures, characterize the unsteady motions, and investigate the mechanisms that drive the unsteadiness of primary and corner SBLI. Results show that the corner and primary SBLI units differ fundamentally in both their mean and unsteady features and their response to upstream and downstream flow perturbations. These observations suggest that the two behave as independent units at this relatively low shock-induced back pressure ratio.