Numerical Study on Flow and Cooling Characteristics for Supersonic Film Cooling

E JECTORS have a wide variety of applications: thrust augmentation of jet engines, vacuuming systems, and thermocompressor of the desalination plant, to name a few [1–7]. Depending on the applications, different configurations of ejectors have been in use. Interestingly, however, most of the past works on ejectors reported central injection ejectors, where primary flow is injected along the centerline of the secondaryflow.When ejectors are used for pumping chemical lasers, central injection type ejectors cannot be used because the passage of primary flow is exposed to the secondary flow of hot burnt gas with temperatures well beyond 1200 K [8,9]. By injecting the primary flow annularly, direct contact between the hot secondary flow and primary flow passage can be avoided.Annular injection of the primaryflow is also used in a rocket based combined cycle engine, where the high-momentum of the secondary flow can be maintained by removing the protrusion of the primary flow passage [10,11]. As the annular injection ejectors make up an important part of the systems described above, the study of the annular injection ejectors have been device specific, which can explain the lack of the literature on ejectors with this injection arrangement. In the present study, we investigated the effect of shape of the primary nozzle and configuration of the flow passage downstream of the primary nozzle exit on the performances of an isolated annular injection ejector, namely, static pressure of the secondary pressure and the primary stagnation pressure at the starting and unstarting conditions. By doing so, we intend to understand the performance characteristics and provide a baseline data for ejector sizing. Figure 1 is a typical performance curve of an annular injection supersonic ejector. The normalized primary stagnation pressure was plotted against the normalized secondary pressure. As the stagnation pressure of the primary flow increases, the forepart of the diverging section of the primary nozzle becomes supersonic and the aftpart becomes subsonic with a normal shock demarcating these two flow regions. As a result, subsonicmixing occurs between the primary and secondary flows in the entire mixing chamber. This condition corresponds to region (1) in Fig. 1. As the stagnation pressure of the primary flow increases further, the shock wave is pushed outside of the primary nozzle. Therefore, supersonic mixing takes place in part of the mixing chamber as shown in region (2). When the primary stagnation pressure increases beyond the starting pressure in region (3), the whole mixing chamber is filled with supersonic primary flow, and the shock is swallowed by the second throat. At this condition, the design static pressure of the secondary flow is

Key features that drive the operation of a supersonic ejector are the complex gasdynamic interactions of the primary and secondary flows within a variable area duct and the phenomenon of compressible turbulent mixing between them, which have to be understood at a fundamental level. An experimental study has been carried out on the mixing characteristics of a two dimensional supersonic ejector with a supersonic primary flow (air) of Mach number 2.48 and the secondary flow (subsonic) which is induced from the ambient. The non-mixed length, which is the length within the ejector for which the primary and secondary flow remain visually distinct is used to characterize the mixing in the ejector. The operating pressures, flow rates and wall static pressures along the ejector have been measured. Two flow visualization tools have been implemented—time resolved schlieren and laser scattering flow visualization. An important contribution has been the development of in-house image processing algorithms on the MATLAB platform to detect the non-mixed length from the schlieren and laser scattering images. The ratio of mass flow rates of the secondary flow to primary flow (entrainment ratio) has been varied in a range of 0.15–0.69 for two locations of the primary nozzle in the ejector duct. Representative cases have been computed using commercial CFD tool (Fluent) to supplement the experiments. Significant outcomes of the study are—the non-mixed length quantified from the flow visualization images is observed to lie within 4.5 to 5.2 times the height of the mixing duct which is confirmed by the wall static pressure profiles. The flow through the supersonic ejector in the mixed regime is explained using corroborative evidences from different diagnostic tools. A reduction of the non-mixed length by 46.7% is observed at operating conditions when the nozzle is sufficiently overexpanded. The disturbance caused to the mixing layer due to unsteady shock-boundary layer interactions within the nozzle at such conditions enhances mixing. The analysis of time resolved schlieren images have provided interesting observations on repetitive back and forth motion of the shock cells in the primary flow with a co-flowing secondary flow in the confines of the supersonic ejector. The oscillations have significant amplitudes (order of the nozzle height) at the centerline. The details of these experiments followed by the analysis of data and the inferences drawn from the results are discussed in this article.

Compressible flow past thin wings and slender bodies are studied. First, subsonic flow past thin wings are analyzed by means of potential flow theory. The kernel function method is introduced for arbitrary planforms undergoing simple harmonic oscillations. The spanwise polynomial approximation and the chordwise trigonometric function approximations which automatically satisfy the Kutta condition and inherit the leading edge singularity are used. The Doublet-Lattice method as a more general numerical approach is given for the analysis of the flow past additional surfaces like tail or store surfaces, which are not necessarily in plane with the wing surface or surfaces having spanwise deflections. Airfoil response to the arbitrary unsteady motion is also given for subsonic flows. A brief review of shock waves and Mach waves in a supersonic flow is given. Afterwards, unsteady supersonic potential flow is studied for a simple harmonically oscillating point source. First, unsteady flow past an airfoil is considered. Then, supersonic flow about thin wings are analyzed using the Mach box technique. Introduction to supersonic kernel function method is briefly provided. Arbitrary unsteady motion of an airfoil in a supersonic flow is presented. Slender body theory is introduced to analyze the cross flow past bodies of considerable fineness where the cross flow is shown to be approximately incompressible under certain conditions. In connection with the slender body approximation, the Munk’s airship theory is utilized to predict the stability derivatives of missile like bodies.

Discharge header design inside a reactor pool for flow stability in a research reactor

Three-dimensional viscous flow computations are presented for 90 deg injection angle jets in subsonic and supersonic cross flow. Comparisons with experimental data include jet centerline and vortex trajectories for the subsonic cross flow, and surface pressure measurement for the supersonic crossflow case. The vortices induced in the jet/freestream interaction are computed and illustrated. The vortices persist in subsonic flow and die out quickly in supersonic flow. The structure of the shocks in the unconfined supersonic flow is illustrated.

It is pointed out that there are two separate mechanisms for upstream influence through the boundary layer in supersonic flow, and that one of these (that involving separation) operates also in subsonic flow. A quantitative theory of subsonic flow up a step is given to illustrate this. The main differences between the subsonic and supersonic flows are as follows: (i) The boundaries of dead-air regions are nearly straight in supersonic flow but are usually highly curved in subsonic flow. (ii) Separation (whether of the laminar or turbulent layer) occurs at a much lower pressure coefficient in supersonic flow; this is only slightly due to the fact that the fluid nearest the wall is then lighter and so more easily brought to rest; it is due much more to the relative suddenness of the pressure rise ahead of the dead-air region. (iii) However, for a given pressure coefficient in the dead-air region, the distance of upstream influence is somewhat greater in the subsonic flow, except at the highest pressures. A qualitative discussion of the second mechanism of upstream influence, in supersonic flow, is given; for a quantitative theory of this see part II (Lighthill 1953).

Characteristics of gas flow in the transition regime are presented for microchannels for both subsonic and supersonic flows. A low speed pressure driven flow and two high speed flow cases through microchannels have been documented using a hydrodynamic finite element formulation with first order slip/jump boundary conditions. The subsonic flow case uses a channel of aspect ratio 5639, while for the high speed flow the microchannel has an aspect ratio of 5. Presented subsonic flow results are benchmarked with experiment and reported numerical model. The solution profiles for supersonic micro flow cases are compared to the published DSMC results. Numerical solution of the hydrodynamic model documents excellent shock capturing capability.

High-temperature, high supersonic Mach number flows can be encountered in different innovative propulsion system configurations such as rotating detonation combustors (RDCs) and scramjets, requiring the use of film cooling for the thermal management. For this reason, a clear understanding of the performance of film cooling in supersonic flow is paramount. This experimental study investigated the performance of film cooling in a supersonic flow at Mach 1.65 comparing it with subsonic flow operations at Mach 0.3. Adiabatic effectiveness measurements were performed using pressure-sensitive paint. This technique also provided pressure field maps, useful to enhance the understanding of the physical phenomena under investigation. Additionally, time-averaged Schlieren images enabled the visualization of the flow field morphology and understanding of the underlying physics. Cylindrical holes both aligned with the flow and with a 30 deg compound angle and 7-7-7 fan-shaped holes aligned with the flow were considered. The results indicated performance improvements in both centerline and laterally averaged adiabatic effectiveness induced by the action of the oblique shock forming upstream of the hole. The same shock is responsible for modifying the shape of the adiabatic effectiveness contour, determining a higher peak along the centerline, which linearly decreases outward instead of showing the more typical bell shape. The fan-shaped holes outperform the other holes at any blowing ratio (BR) over 0.8 for supersonic main flow operations. Performance of cylindrical holes in the supersonic flow follows the trend typically encountered in subsonic flows, while also bringing the peak in adiabatic effectiveness upstream and closer to the hole exit.

A mathematical model was developed to investigate the water vapor spontaneous condensation under supersonic flow conditions. A numerical simulation was performed for the water vapor condensable supersonic flows through Laval nozzles under different flow friction conditions. The comparison between numerical and experimental results shows that the model is accurate enough to investigate the supersonic spontaneous condensation flow of water vapor inside Laval nozzles. The influences of flow friction drag on supersonic spontaneous condensation flow of water vapor inside Laval nozzles were investigated. It was found that the flow friction has a direct effect on the spontaneous condensation process and therefore it is important for an accurate friction prediction in designing this kind of Laval nozzles.

The lethality effect of high power laser on target is simulated with CFD method under different conditions of supersonic air flow on the surface of the target. Materials used in the experiments are 2cm aluminum plate. With the Mach number changing from 1 to 5, the lethality effects of the high power laser can be obtained from the simulations under these conditions of supersonic air flow. The flow-structure-laser coupling impact on the failure time of the target is discussed based on the simulation. Results show that with the increase of mach number, the effect on the aluminum plate is increase first and then decrease by the pressure. Because that it is obvious that the maximum area of pressure is away from the center of deformation region when the mach number is bigger than 5 . At the same time, when mach number is increase, the aerodynamic heating play more important role than the convective heat transfer on the temperature field of aluminum plate. there are two impacts from the supersonic flow. Firstly , the flow can produce the pressure on the surface of the aluminum plate. Secondly, the flow can produce aerodynamic heat on the aluminum plate.

Shock losses and Pitot tube measurements in non-ideal supersonic and subsonic flows of Organic Vapors

High temperature, high supersonic Mach number flows can be encountered in different innovative propulsion systems configurations such as Rotating Detonation Combustors (RDCs) and scramjets, requiring the use of film cooling for the thermal management. For this reason, a clear understanding of the performance of film cooling in supersonic flow is paramount. This experimental study investigated the performance of film cooling in a supersonic flow at Mach 1.65 comparing it with subsonic flow operations at Mach 0.3. Adiabatic effectiveness measurements were performed using Pressure Sensitive Paint. This technique also provided pressure field maps, useful to enhance the understanding of the physical phenomena under investigation. Additionally, time-averaged schlieren images enabled the visualization of the flow field morphology and understanding of the underlying physics. Cylindrical holes both aligned with the flow and with a 30 degrees compound angle and 7-7-7 fan-shaped holes aligned with the flow were considered. The results indicated performance improvements in both centerline and laterally averaged adiabatic effectiveness induced by the action of the oblique shock forming upstream of the hole. The same shock is responsible for modifying the shape of the adiabatic effectiveness contour, determining a higher peak along the centerline which linearly decreases outward instead of showing the more typical bell shape. The fan-shaped holes outperform the other holes at any blowing ratio (BR) over 0.8 for supersonic main flow operations. Performance of cylindrical holes in supersonic flow follow the trend typically encountered in subsonic flows while also bringing the peak in adiabatic effectiveness upstream and closer to the hole exit.

The effect of surface roughness on breakaway separation

NuScale Power, Inc. is commercializing a 45 Megawatt electric light water nuclear reactor NuScale Power Module (NPM). Each NPM includes a containment vessel, a reactor vessel, a nuclear reactor core, an integral steam generator, and an integral pressurizer. The NuScale Power Module is cooled by natural circulation. The primary coolant in the Reactor Pressure Vessel is heated in the nuclear core, it rises through a central riser, it spills over and encounters the helical coil steam generator, it is cooled as steam is generated inside the steam generator, and it is again heated in the nuclear core. The Steam Generator also must be designed to provide adequate heat transfer, to allow adequate primary reactor coolant flow, and to provide adequate steam flow to produce the required power output. This paper presents the CFD results that describe the transport phenomena on the heat transfer and fluid flow dynamics in helical coil steam generator tubes. The ultimate goal of the CFD modeling is to predict the steam outlet conditions associated with the chosen helical coil tube geometries, solving the primary and secondary flow region together coupled with the helical coil tube. However, current studies are focused on the primary side with the heat flux boundary condition assigned on the outer surface of the helical coil steam generator. In this study, the ANSYS CFX v. 12.1 [1] was used to solve the three-dimensional mass, momentum and energy equations. The helical coil steam generator has complex geometry and modeling entire geometry requires the enormous memory that is beyond our hardware capability and is not practical. Therefore, geometry was limited to 1 degree of the wedge and 5% of the total length in the middle. Only external flow, single phase flow around the helical coils, is simulated using the standard k-ε model and shear stress transport model. From the results of the numerical simulation, the pressure drop and temperature profiles were determined. It is important to understand thermal hydraulic phenomena for the design and performance prediction of the reactor internal.

Experimental investigation of a directionally enhanced DHX concept for high temperature Direct Reactor Auxiliary Cooling Systems