Wall Static Pressure Research Articles

Investigation of flow characteristics of various-aspect-ratio rectangular nozzles with an aft deck

This study presents an experimental and numerical investigation to characterize the plume pattern of a high-aspect-ratio rectangular convergent/divergent nozzle with an aft deck in under-expanded conditions. The function of an aft deck is to shield the infrared signal of an exhaust plume at its strongest intensity located at the immediate downstream region of the nozzle exit. However, this practice may cause undesirable plume deflection, which needs to be reduced as much as possible. The nozzle pressure ratios ranged from 2 to 4, and the effect of the nozzle exit aspect ratio was examined using wall static pressure measurements and schlieren visualization for cold flows. The experimental setup involved a 3D-printed aft deck nozzle made of acrylonitrile butadiene styrene material, which underwent surface smoothing using acetone vapor. Numerical simulations were conducted using the commercial STARCCM+ software to analyze static pressure ratio variations at the aft deck. The investigation revealed that a nozzle pressure ratio of 3 induced a downward plume deflection at aspect ratio values of 6.77 and 7.54, while an increased aspect ratio of 8.35 resulted in the horizontal ejection of the plume. Moreover, at an aspect ratio of 8.35, the plume was ejected horizontally for nozzle pressure ratios ranging from 2 to 4. At a nozzle pressure ratio of 4, the flow separated from the deck without reattaching, and the plume moved horizontally with minimal deflection. The findings suggest that a combination of a high aspect ratio and sufficiently high nozzle pressure ratio can effectively reduce plume deflection.

Experimental DMD Analysis of Combustion Modes and Instabilities in a Scramjet Combustor

The combustion characteristics of a gaseous hydrogen fueled scramjet combustor were experimentally investigated through single-shot micro-pulse detonation engine (μPDE) operations across a range of fuel injection pressures and corresponding equivalence ratios. This study utilized a lab-scale direct-connect scramjet combustor integrated with a μPDE, designed to simulate flight conditions at Mach numbers between 4.0 and 5.0 at altitudes of 20 to 25 km. Static wall pressure was measured using 16 pressure probes from the isolator to the scramjet combustor at a sampling rate of 500 Hz. Additionally, Z-fold type Schlieren and flame luminosity were captured using a high-speed camera in a test section equipped with a quartz visualization window. Experiments were conducted under five conditions with equivalence ratios ranging from 0.10 to 0.55, each performed three times to ensure repeatability. Fast Fourier transform (FFT) and dynamic mode decomposition (DMD) analyses were applied to the pressure and image data obtained. Although auto-ignition was not achieved in all cases, combustion persisted while fuel was injected following the single-shot μPDE ignition. This allowed for the identification of distinct combustion modes according to equivalence ratio, by observing the dominant flame types and the shock structure within the flow of the combustor. These combustion modes showed different types of instability.

Analysis of HEM applicability for a subcritical flashing ejector at low motive pressure

Validated CO2 ejector models are essential for developing high-performance refrigeration and heat pump cycles. This study focuses on assessing the Homogeneous Equilibrium Model’s applicability to simulate a CO2 flashing ejector at a reduced pressure of 0.47. The model was implemented in FLUENT, integrating a user-defined real gas model. Simulation results with different boundary condition options were compared to experimental data. The analysis was carried out to evaluate the predictive capabilities of the model and assess the experimental data quality. The results indicate that the developed model accurately estimated the motive mass flow rate, with a maximum relative error of 5.7%, showing better performance than previously reported data. The entrained flow rate, assuming double choking operation, was significantly higher than the experimental measurement, and the CFD-predicted wall static pressure underestimated the experimental profile, suggesting single-choked ejector operation. In contrast, the outflow density was better predicted under the same assumption, with an average error of 8.6%. Nevertheless, the simulated temperature profiles showed good agreement with the experimental data, especially when using the experimental entrained mass flow rate as a boundary condition.

Effects of Perforated Plates on Shock Structure Alteration for NACA0012 Cascade Configurations

To alleviate the shock boundary layer interaction adverse effects, various active or passive flow control strategies have been investigated in the literature. This research sheds light on the behavior of perforated plates as passive flow control techniques applied to NACA0012 airfoils in cascade configurations. Two identical perforated plates with shallow cavities underneath are accommodated on the upper and lower surfaces of each airfoil in the cascade arrangement. Six different cascade arrangements, including a baseline configuration with no control applied, are additively manufactured, with different perforated plate orifice sizes in the range of 0.5–1.2 mm. A high-speed wind tunnel with Schlieren optical diagnosis and wall static pressure taps is used to investigate the changes in the shock waves pattern triggered by the perforated plates. Steady 3D density-based numerical simulations in Ansys FLUENT are conducted for further analysis and validation. In the cascade configuration, the perforated plates alter the shock structure, and the strong normal shock wave is replaced by a weaker X-type shock structure. Eventually, a 1% penalty in overall total pressure loss is induced by the perforated plates because of the negative loss balance between the reduced shock losses and the enhanced viscous losses. Further studies on perforated plate geometrical features are needed to improve this outcome in a cascade arrangement.

Experimental investigation and numerical analysis on high-efficiency shock vectoring control serpentine nozzles

The high-efficiency Shock Vectoring Control Serpentine Nozzle (SVCSN) takes into account both thrust vectoring and infrared stealth, and significantly improves the comprehensive performance of the aero-engines through an additional auxiliary duct. In this paper, the schlieren photographs at the exit of the high-efficiency SVCSN and the wall static pressure distributions were obtained by experiments, and the numerical results were used to enrich the thrust vectoring characteristics. The effects of the auxiliary injection were analyzed first to reveal the advantages of the high-efficiency SVCSN compared to the conventional SVCSN. Then, the aerodynamic parameters and the structural parameters of the high-efficiency SVCSN were investigated, including the Nozzle Pressure Ratio (NPR), the Secondary flow Pressure Ratio (SPR), the secondary flow relative area and the secondary flow injection angle. Finally, the coupling performance of the high-efficiency SVCSN is studied by using the approximate modeling technology. Results show that the auxiliary injection increases the range between the two shock legs of the “λ” shock wave induced by the secondary flow, then causes the separation zone and high-pressure boss of the down wall to expand upstream, and finally results in a prominent increase in the thrust vectoring performance. The thrust vectoring angle and Vectoring Efficiency (VE) of the high-efficiency SVCSN are about 61.6% and 75.7%, respectively, higher than those of the conventional SVCSN at NPR = 6. The effects of the NPR and the SPR on the thrust vectoring performance of the high-efficiency SVCSN are coupled with each other. A larger NPR matched with a smaller SPR shows better thrust vectoring performance. The maximum fluctuations in thrust vectoring angle and VE caused by the NPR and SPR are about 22% and 64%. The VE decreases monotonously with the increase of the secondary flow relative area. Smaller secondary flow injection angle shows better thrust vector performance, and the thrust vectoring angle and VE of the secondary flow injection angle of 90° are about 20% higher than those of the secondary flow injection angle of 110° at NPR = 6. Therefore, the secondary flow relative area of 0.06 and the secondary flow injection angle of 90° are recommended.

Numerical investigations on flow and heat transfer characteristics of a high-enthalpy double-cone in thermal and chemical non-equilibrium for hypersonic propulsion

There are strong interests in hypersonic propulsion communities in predicting the flow and heat transfer characteristics over a hypersonic vehicle. However, this is quite challenging due to the emergence of thermal and chemical out-of-equilibrium effects as prominent in a high-temperature flow. In this work, we conduct 2D numerical investigations on the hypersonic double cone flow in a surrounding of the stagnation enthalpy of 15.23 MJ/kg by solving the laminar Navier–Stokes equation systems with the vibrational non-equilibrium processes and seven species air reactions. We then conduct the sensitivities analyses on the flow and heat transfer features, as different transport models and different thermochemical models are applied. It is shown that the aerodynamic heat is more sensitive to the transport models than the flow, and the transport coefficients computed by the Gupta-Yos model are more suitable for predicting the double cone flow at the high enthalpy. On the other hand, the flow field of the high enthalpy double cone is more sensitive to the thermochemical models than the heat transfer on the wall excluding the peak position. Behind the detached shock, chemical reactions are found to enhance thermal or vibrational out-of-equilibrium effects. However, vibrational non-equilibrium effects are revealed to weaken the degree of molecule dissociation reactions. Two flow separations occur in the flow field computed by using the thermal non-equilibrium gas models, but the flows of thermal equilibrium gas models separate only once. The maximum values of wall parameters without chemical reactions are found to deviate greatly from the experimental measurements. The thermochemical out-of-equilibrium gas model can predict the separation region size closer to the experiment than the two reported numerical results. The wall static pressure and heat flux are also shown to well match the experimental results. The present work should shed lights on better understanding thermochemical non-equilibrium effects of complex flow phenomenon like the shock wave/boundary layer interaction in hypersonic propulsion with high enthalpy surroundings.

Structural optimization design of ice abrasive water jet nozzle based on multi-objective algorithm

To enhance the cleaning efficiency of the ice abrasive water jet (IAWJ), the nozzle in the jet structure was optimized. The cleaning effect of nozzle with different structural design is different. The design variables include the geometric parameters of the nozzle's structure and the flow process parameters of the jet. The maximum wall static pressure, outlet velocity and wash width are the optimization objectives. Based on the Optimal Latin Hypercube Sampling (OLHS) method, 50 sets of sample capacity were designed for simulation calculation. The parameter functions associated with the three objectives were fitted by MATLAB. The Competitive mechanism multi-objective particle swarm optimizer (CMOPSO) algorithm was used to optimize the nozzle to improve the cleaning effect, which was verified with the test results. The inlet pressure and outlet diameter are critical factors influencing the cleaning effect of the IAWJ. The maximum wall static pressure produced by the optimized nozzle is about 4.5% higher than that of the ordinary cosine nozzle. The cleaning effect of the optimized nozzle is 10% higher than that of the ordinary cosine nozzle, which is verified by the experiments. The optimization of the nozzle serves as a valuable reference for the nozzle structure design of the cleaning machine.

Mixing enhancement of transverse jets in supersonic crossflow using an actively controlled novel fluidic oscillator

This study investigates the fluid dynamics and mixing characteristics of an oscillating sonic jet injected into a supersonic cross flow of Mach 2.1 using experimental and computational techniques. The oscillating jet is produced by a novel fluidic oscillator, which consists of a primary rectangular duct that expands into an outer duct with sudden expansion. Control jets are injected in the lateral direction from the side walls of the sudden expansion in an out-of-phase manner to oscillate the injected jet in the spanwise direction of the crossflow. Experimental and numerical investigations based on wall static pressure and mass fraction fluctuations, respectively, revealed that the injected jet oscillation frequency matches the control jet frequency. The iso-surface of lambda-2 criterion showed the presence of various dominant vortex structures, such as counter-rotating vortex pairs, horseshoe vortex, sidewall vortices, and trailing vortices. Helicity contour plots showed that the streamwise vortices oscillate in the spanwise direction with the control strategy and promote the spread of the injected jet in the spanwise direction. The spatiotemporal reconstruction (z–t plot) of the density gradients at a particular streamwise location revealed that the bow shock produced by the interaction of the injected jet and the crossflow oscillates with the actuation of the control strategy. The power spectral density of the z–t plot revealed that the shock wave oscillation frequency matches the control jet frequency. The oscillating jet produced by the control strategy showed significant mixing enhancement in supersonic crossflow compared to a simple rectangular injection.

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.

Mode identification and decomposition analysis of self-excited thermodynamic oscillations in hypersonic inlet/isolator of a scramjet

The self-excited pressure and velocity oscillations caused by the downstream back pressure in a hypersonic inlet/isolator directly affect the operational stability of a scramjet. In this work, we conduct the two-dimensional unsteady RANS (Reynolds averaged Navier–Stokes) simulation to predict the flow field in the Mach 7 inlet/isolator with 147 times the freestream static pressure. With the model validated, the time evolution process of the instantaneous flow fields in the isolator is analyzed first. It is shown that the structure of the shock train exhibits an up-down asymmetry. The flow oscillations as observed near the upper wall are more complex than those near the lower wall. Furthermore, the dominant frequency of the wall pressure oscillations is approximately 520 Hz. However, the self-excited oscillations in the isolator flow are dominated with approximately 510 Hz mode but coupled with multiple harmonic higher frequencies. To identify and decompose the modes of these oscillations in the unsteady flow of the isolator, different mode decomposition technologies are applied to further examine the mechanism of such oscillations. By conducting Proper Orthogonal Decomposition (POD) analysis on the x-axis velocity and Mach number fields, it is shown that the first two modes consist of more than 77% of the total oscillations’ kinetic energy. Additionally, the oscillations of the quadrilateral separation zone and low-speed zone near the upper wall largely determine the dominant frequency of the overall unsteady flow field. As Dynamic Mode Decomposition (DMD) investigation is conducted on the x-axis velocity and Mach number fields, it is found that the identified first mode is almost the same as that identified by using POD. It is also found that both DMD and POD can well predict the viscous flows such as the separation zone, shear layer, and low-speed zone. Finally, the dominant frequencies predicted by the POD and DMD are revealed to be very close to the dominant frequency of the wall static pressure oscillations. The movements of multiple shock waves near the upper wall are the major incentive for the complicated change in the entropy generation rate. In general, the self-excited thermodynamic oscillations phenomenon as observed in the hypersonic inlet/isolator in a scramjet could be well analyzed by decomposing and identifying the dominant modes, which contribute mostly to the total energy of such oscillations.

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.

Structural uncertainty quantification of Reynolds-Averaged Navier–Stokes closures for various shock-wave/boundary layer interaction flows

Accurate prediction of Shock-Wave/Boundary Layer Interaction (SWBLI) flows has been a persistent challenge for linear eddy viscosity models. A major limitation lies in the isotropic representation of the Reynolds stress, as assumed under the Boussinesq approximation. Recent studies have shown promise in improving the prediction capability for incompressible separation flows by perturbing the Reynolds-stress anisotropy tensor. However, it remains uncertain whether this approach is effective for SWBLI flows, which involve compressibility and discontinuity. To address this issue, this study systematically quantifies the structural uncertainty of the anisotropy for oblique SWBLI flows. The eigenspace perturbation method is applied to perturb the anisotropy tensor predicted by the Menter Shear–Stress Transport (SST) model and reveal the impacts of anisotropy on the prediction of quantities of interest, such as separation and reattachment positions, wall static pressure, skin friction, and heat flux. The results demonstrate the potential and reveal the challenges of eigenspace perturbation in improving the SST model. Furthermore, a detailed analysis of turbulent characteristics is performed to identify the source of uncertainty. The findings indicate that eigenspace perturbation primarily affects turbulent shear stress, while the prediction error of the SST model is more related to turbulent kinetic energy.

Effect of Diffuser Vane Height and Position on Diffuser Wall Static Pressures in a Centrifugal Compressor

The present paper reports experimental investigations on the effect of diffuser vane height and its position on diffuser wall static pressures of a low speed centrifugal compressor. Apart from the standard types of diffusers, namely, vaneless, vane and low solidity vane diffusers, a partial vane diffuser, whose height is less than the diffuser width is also tested. The vane height of the partial vane diffuser is varied from 0.3, 0.4 and 0.5 times the diffuser width in vane and low solidity vane diffuser configurations. In addition, the effect of vane position is examined by fixing the partial vanes to the hub, shroud or hub and shroud. It is found that there is an optimum height for the diffuser vane height, which is found to be 0.3 times the diffuser width. The effect of position of the partial vanes on the hub or shroud on the diffuser wall static pressures is found to be negligible. However when partial vanes are fixed on the hub and shroud staggered at half spacing, the diffuser performance based on wall static pressures is improved substantially.

Experimental Study on the Ignition Characteristics of Scramjet Combustor with Tandem Cavities Using Micro-Pulse Detonation Engine

This experimental investigation focused on the ignition and combustion characteristics of a tandem cavity-based scramjet combustor with side-by-side identical cavities. This study utilized the Pusan National University-direct connect scramjet combustor (PNU-DCSC), which was capable of simulating flight conditions at Mach number 4.0–5.0 and altitudes of 20–25 km using the vitiated air heater (VAH). The combustion tests were conducted under off-design point conditions corresponding to low inlet enthalpy. It is a condition in which self-ignition does not occur, and a micro pulse detonation engine (μPDE) ignitor is used. The results revealed that as the injection pressure of the gaseous hydrogen fuel (GH2) and the corresponding equivalence ratio increased, the combustion mode transitioned from the cavity-shear layer flame to the jet-wake flame. Furthermore, the measured wall static pressure profiles along the isolator and scramjet combustor indicated that the region of elevated pressure distribution caused by the shock train expanded upstream with higher equivalence ratios. When ignited from the secondary cavity, the combustion area did not extend to the primary cavity at lower equivalence ratios, while it expanded upstream faster with higher equivalence ratios. Therefore, the combustion characteristics of the tandem cavity were found to vary based on the overall equivalence ratio of the main fuel (GH2) and ignition position.

On wall pressure fluctuations in conical shock wave/turbulent boundary layer interaction

The structure and the frequency spectra of wall pressure fluctuations beneath a planar turbulent boundary layer interacting with a conical shock wave at Mach number $M_\infty =2.05$ and Reynolds number $\textit {Re}_\theta \approx 630$ (based on the upstream boundary layer momentum thickness) are examined to elucidate the effects of pressure gradient and flow separation on the characteristics of the wall pressure fluctuations, by exploiting a direct numerical simulation database. Upstream of the interaction, in the zero pressure gradient region, wall pressure statistics compare well with canonical compressible boundary layers in terms of fluctuation intensities and frequency spectra. Across the main interaction zone (APG1), the root-mean-square of wall pressure fluctuations becomes very large (corresponding to approximately 173.3 dB), with maximum increase approximately 12.7 dB from the incoming level. In the second adverse pressure gradient zone (APG2), the root-mean-square of wall pressure fluctuations attains a second peak (corresponding to $164.7$ dB), with an increase of 8.4 dB from the upstream level. Both the APG1 and APG2 regions feature a substantial fraction of flow reversal events, which are, however, scattered and interspersed with regions of attached flow. The wall pressure power spectral density exhibits a broadband and energetic low-frequency component associated with the global unsteadiness of the separation bubble/conical shock system. Analysis of the two-point correlations and wavenumber/frequency spectra of wall pressure fluctuations further suggests that the typical eddies become more elongated along the spanwise direction, as the flow in the separated region tends to escape the centreline, and the convection velocity is significantly reduced.

Experimental investigation of a novel variable geometry radial ejector

This work focuses on improving ejector performance in variable conditions by developing a radial flow variable geometry radial ejector (VGRE). The research objective was to demonstrate the performance of the improved VGRE, which can operate more effectively under changing operating conditions compared to a fixed geometry design. The improved VGRE was developed and experimentally tested to enable adjustment of the nozzle and ejector throat areas to suit different ejector operating conditions in various applications. Radial ejectors have a smaller size than axial ejectors, which is a significant advantage in some applications. Results showed that the improved VGRE achieved higher entrainment ratio and critical compression ratio compared to a fixed geometry design and that optimal performance was obtained with nozzle and duct throat separations of 0.5 mm and 3.0 mm, respectively. The results also indicated that the wall static pressure in the throat region of the improved VGRE was similar to that observed in conventional axial ejectors and showed an average improvement of 107 % in entrainment ratio and 76 % in critical compression ratio relative to previous studies of air ejectors. The findings suggest that the improved VGRE concept could be applied to a wide range of applications in the future.

Effect of wall roughness on secondary flow choking in supersonic air ejector with adjustable motive nozzle

Wall roughness is present to a certain extent in all engineering devices and systems. However, it is rarely considered during numerical simulations of supersonic ejectors. In this study, the effect of wall roughness on choking within a supersonic ejector with an adjustable motive nozzle was investigated in detail. Wall roughness was studied through numerical simulations. Characteristic curves showed that roughness had almost no influence on secondary flow choking when choking occured geometrically before the mixing chamber. By contrast, when choking occurred within the mixing chamber, its phenomenology was significantly altered by wall roughness. In addition, the results confirmed the ability of an ejector to switch from the on-design working regime to the off-design one when wall roughness is considered. The numerical results were also compared with available experimental data. Accounting for wall roughness in the numerical analysis significantly improved the agreement between the experimental and numerical data, especially in terms of wall static pressure distributions. This study sheds light on secondary flow choking within a supersonic ejector considering wall roughness.

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.

Effect of free boundary on the performance of single expansion nozzle

Abstract Effect of free boundary on the flow field characteristics of a single expansion nozzle has been studied experimentally and computationally. The single expansion nozzles studied in this investigation are a convergent-divergent nozzle of rectangular cross-section with convergent-divergent wall on one side and a flat wall on the opposite side, and an identical convergent-divergent ramp with its top open to atmosphere. The studies have been carried out at nozzle pressure ratios 2, 3, 4 and 5. The results show that the single expansion nozzle with wall boundary is able to deliver the flow with Mach number around 1.5, at nozzle pressure ratios of 4 and 5 even with the single expansion. The nozzle with free boundary, the wall static pressure is appreciably lower than that of nozzle with closed boundary and the exit Mach number is 1.5 for NPR 4 and 1.75 for NPR 5. That is, at the exit, the single expansion nozzle with free boundary delivers higher Mach number compared to single expansion nozzle with wall boundary for NPR 5.

Numerical analysis for jet flow field on cleaning the outer wall of the double-layer injection needle

To avoid cross-contamination between different samples, improve the cleaning effect of the outer wall of the inner needle of the digital PCR detector, and reduce the cost of cleaning solution, the cleaning jet was simulated by FLUENT software using VOF gas-liquid two-phase flow model and PISO algorithm. In a certain range, the inlet velocity, cleaning solution viscosity, interfacial tension, and contact angle are changed to simulate the theoretical cleaning of the outer wall of the injection probe and the selection of various parameters under the best cleaning conditions is determined. The results show that the increase of inlet velocity has no significant effect on the increase of wall static pressure and the decrease of velocity increases; the increase of viscosity will greatly reduce the flow rate; higher interfacial tension requires higher inlet velocity to form a continuous jet; increasing the contact angle will not significantly reduce the static pressure and flow rate, and the residual sample is easier to clean. When the inlet velocity of 200mm /s is selected, the fluorine oil with a viscosity of 0.00124Paꞏs and interfacial tension of 16.2dyn/cm is used and the wall is coated with Teflon, the cleaning effect, and economy are excellent.