Exhaust Flow Field Research Articles

Performance Evaluation of Different Rotating Detonation Combustor Designs Using Simulations and Experiments

Abstract Rotating Detonation Engines (RDEs) present a potential avenue for enhancing the efficiency of gas turbine combustors through the utilization of a detonation-driven process. However, integrating an RDE with a downstream turbine poses a significant challenge due to the shock-laden and highly unsteady nature of the RDE exit flow, coupled with a high degree of flow periodicity. In contrast, gas turbines are designed to operate with relatively small velocity and temperature fluctuations at the turbine inlet. The objective of the present study is to develop an understanding of how geometric profiling can improve RDE performance and mitigate exhaust flow field unsteadiness. Three RDE designs are analyzed: an annular combustor with a constant cross-sectional area, an annular combustor with a converging nozzle near the exit, and an annular combustor with a rapid to gradual (RTG) area convergence. Three-dimensional unsteady reacting simulations are conducted for each configuration using the same fuel-oxidizer composition and mass flow rate condition. All simulations are validated against experimental measurements, and RDE performance is assessed based on total pressure gain/loss, unsteadiness at the exit of the RDE, and detonation effectiveness. Results show significant performance improvement for both the convergent nozzle and the RTG profile compared to the constant cross-sectional area configuration, and the RTG design performed better than the convergent nozzle design. Therefore, strategically constricting the flow in an RDE can be used to optimize the performance of an RDE and should be considered in future designs.

Numerical study on the effect of buoyancy-driven pollution source on vortex ventilation performance

High-temperature buoyancy-driven pollutants that generally exist in industrial buildings and kitchens often be eliminated by local exhaust ventilation systems. The high-temperature pollutant plume driven by buoyancy significantly affects the local exhaust process, resulting in the change in the exhaust flow field and reduction of the pollutant capture efficiency. In this study, the effects of thermal buoyancy-driven pollution sources on the vortex flow field and pollutant removal efficiency of a vortex ventilation system are evaluated through experiments and numerical simulations. The results show that with an increase in the buoyancy flux of the pollution source, the minimum negative pressure and maximum tangential velocity in the vortex flow first increase and then decrease, that is, the buoyancy plume first enhances and then weakens the vortex flow. A critical buoyancy flux exists, and the value of the critical buoyancy flux increases with an increase in the vortex ventilation intensity. Furthermore, the capture of high-temperature pollutants by vortex flow can be categorized as “plume-dominated” and “vortex-dominated”. These two types exhibit significant differences in pollutant transport paths and pollutant capture times during pollutant removal. In addition, on the basis of the pollutant capture efficiency, the vortex airflow performance with respect to capturing high-temperature buoyancy-driven pollutants can be classified as low-efficiency, transition, high-efficiency, and invalid zones. In the process of capturing high-temperature pollutants, vortex ventilation systems should be avoided in the low-efficiency and invalid zones.

Research on characteristics of thermoelectric generator with integrated muffler and catalytic converter

Automobile exhaust contains a large amount of heat loss, which has a high potential for energy saving. In order to recover the heat in the exhaust gas and improve the compatibility of thermoelectric generator (TEG), muffler and catalytic converter, an integrated automobile exhaust thermoelectric generator (IATEG) integrating muffler, and catalytic converter is proposed. Research on its performance by establishing a three-dimensional model, On the basis of finite element simulation, the model’s surface temperature field, muffler transmission loss, exhaust back pressure and conversion efficiency are compared and studied when the catalytic converter is placed in the front part (A), middle part (B) and rear part (C) of the model. The results show that when the exhaust temperature is 350 °C and the inlet velocity is 30 m/s, the average surface temperature of model B is the highest, which is 291.95 °C. The exhaust back pressure of the three models is almost the same and meets the requirements of QC/T 631-2009 (15 kPa). Compared with the original muffler of the automobile, the transmission loss of the three models is greater than that of the original muffler. From the perspective of improving the conversion efficiency of NOx, the exhaust flow field distribution of the catalytic converter of Model B is more uniform and has a smaller impact on the conversion efficiency of NOx. Therefore, it is recommended that IATEG choose the structure of Model B. This article demonstrates the possibility of IATEG integrating muffler and catalytic converter in terms of performance.

Afterburning effect on thermal environment of four-engine liquid rockets at different altitudes

The liquid rocket propellant burns inadequately in the combustion chamber; therefore, the exhaust gas reacts with oxygen and the heat that was generated, which in turn affects the thermal environment. By using finite-rate chemical kinetics and discrete ordinates method, the calculation model for the four-engine rocket exhaust flowfield is established, and numerically investigates the afterburning effect on the thermal environment of the engine exhaust plume under different altitudes. The results show that the afterburning reactions have little influence on the Mach number exhaust flow, and the difference between the peak temperatures of reacting and non-reacting flows is less than 5%. However, a significant increase in temperature of the mixing layer can be found, and the afterburning effect on the thermal environment reduces with increases in the flight altitude. At the same altitude, the temperature increment of the reacting flow in the far-field region is higher than that of the near-field region. Moreover, as the flight altitude increases from 15 to 35 km, the increase of the rocket base heat flux caused by the afterburning can vary from 9.4% to 2.7%. The results provide a theoretical basis for the designing of the thermal protection systems of multi-engine rockets.

Experimental Study of the Jet Engine Exhaust Flow Field of Aircraft and Blast Fences

A combined blast fence is introduced in this paper to improve the solid blast fences and louvered ones. Experiments of the jet engine exhaust flow (hereinafter jet flow for short) field and tests of three kinds of blast fences in two positions were carried out. The results show that the pressure and temperature at the centre of the jet flow decrease gradually as the flow moves farther away from the nozzle. The pressure falls fast with the maximum rate of 41.7%. The dynamic pressure 150 m away from the nozzle could reach 58.8 Pa, with a corresponding wind velocity of 10 m/s. The temperature affected range of 40°C is 113.5×20 m. The combined blast fence not only reduces the pressure of the flow in front of it but also solves the problems that the turbulence is too strong behind the solid blast fences and the pressure is too high behind the louvered blast fences. And the pressure behind combined blast fence is less than 10 Pa. The height of the fence is related to the distance from the jet nozzle. The nearer the fence is to the nozzle, the higher it is. When it is farther from the nozzle, its height can be lowered.

The Numerical Analysis of Flow Field on Warship Deck

This study aims to simulate the exhaust flow field of ship by the method of computational fluid dynamics (CFD) concerning with the interference by exhaust temperature, shape of stack and rolling angles etc.. In this research wind tunnel test for a corvette has been performed to attain associated experimental data, which were used as a reference basis. During simulation process several configurations of stacks have been selected, and combining with various rolling angles, exhaust temperatures and velocities, we have gene- rated numerous cases from which the diffusion paths and temperature distribution of the exhaust flow field can be clearly observed and analyzed. In terms of numerical simulation, the packaged program computational fluid dynamics software has been adopted. The simulation results also possess the same trend as the experimental data, which have initially confirmed the methods developed here can be used for the arrangement of stack and superstructure at the stage of initial and conceptual design of ships. Keywords—computational fluid dynamics, diffusion distribution, wind tunnel test, smoke diffusion, atmospheric pressure. I. Int ro d uct io n

F-35A and F-35B aircraft ground run-up acoustic emissions

A multi-organizational effort led by the Air Force Research Laboratory conducted acoustic emissions measurements on the F-35A and F 35B aircraft at Edwards Air Force Base, California, in September 2013. These measurements followed American National Standards Institute/Acoustical Society of America S12.75-2012 to collect noise data for community noise models and noise exposures to aircraft personnel. In total, over 200 unique locations were measured with over 300 high fidelity microphones. Multiple microphone arrays were deployed in three orientations: circular arcs, linear offsets from the jet-axis centerline, and linear offsets from the jet shear layer. The microphone arrays ranged from distances 10 ft outside the shear layer to 4000 ft from the aircraft with angular positions ranging from 0° (aircraft nose) to 160° (edge of the exhaust flow field). A description of the ground run-up acoustic measurements, data processing, and the resultant data set is provided.

Kitchen and Toilet of the Residence Exhaust Flow Field Analysis

Adopting the k-ε standard model, the CFD simulation software to simulate the indoor kitchen and toilet different row of indoor air volume air distribution in natural ventilated circumstance, by comparison results show that different row of indoor air volume changes in the rate of secondary pollution rate, and then to provide the change rule of indoor air quality protection reference basis.

Propagation characteristics of shock waves driven by gaseous detonation waves

We experimentally investigated propagation characteristics of the shock wave driven by a gaseous detonation wave emerging from the open end of a cylindrical detonation tube. In the present study, we visualized the shock wave and exhaust flowfields using a shadowgraph optical system and we obtained peak overpressure in the tube axial direction and the continuous shape transformation of shock waves around the tube open end. We also obtained overpressure histories of the shock wave using piezo-pressure transducers within 201 m from the open end of the tube. We normalized and classified these results by four regions using non-dimensional pressure and distance which are independent of variety of mixture and tube diameter. In the vicinity of the open end of the tube, the shock wave is nearly planar and does not significantly attenuate, and the peak overpressure maintains approximately C–J pressure. Subsequently, the shock wave attenuates rapidly, transforming from quasi-spherical to spherical. Farther from the tube open end, the shock wave propagates with approximately sound characteristic so that the peak overpressure decreases proportional to 1/r. Eventually, the shock wave begins to attenuate more rapidly than ideal sound attenuation, which may be due to the viscous effect.

Comparison Between Numerically Simulated and Experimentally Measured Flowfield Quantities Behind a Pulsejet

Pulsed combustion is receiving renewed interest as a potential route to higher performance in air breathing propulsion and ground based power generation systems. Pulsejets offer a simple experimental device with which to study unsteady combustion phenomena and validate simulations. Previous computational fluid dynamics (CFD) simulations focused primarily on pulsejet combustion and exhaust processes. This paper describes a new inlet sub-model which simulates the fluidic and mechanical operation of a valved pulsejet head. The governing equations for this sub-model are described. Sub-model validation is provided through comparisons of simulated and experimentally measured reed valve motion, and time averaged inlet mass flow rate. The updated pulsejet simulation, with the inlet sub-model implemented, is validated through comparison with experimentally measured combustion chamber pressure, inlet mass flow rate, operational frequency, and thrust. Additionally, the simulated pulsejet exhaust flowfield, which is dominated by a starting vortex ring, is compared with particle imaging velocimetry (PIV) measurements on the bases of velocity, vorticity, and vortex location. The results show good agreement between simulated and experimental data. The inlet sub-model is shown to be critical for the successful modeling of pulsejet operation. This sub-model correctly predicts both the inlet mass flow rate and its phase relationship with the combustion chamber pressure. As a result, the predicted pulsejet thrust agrees very well with experimental data.

Optimal rigid-body motions

Optimal rigid-body angular motions are investigated in the absence of direct control over one of the angular velocity components. A numerical survey of first-order necessary conditions for optimally reveals that, over a range of boundary conditions, there are, in general, several distinct extremal solutions. A classification in terms of subfamilies of extremal solutions is presented. Domains of existence of the extremal subfamilies are established. Locus of Darboux points are obtained, and global optimality of extremal solutions is observed in relation to Darboux points. Local optimality for the candidate minimizers is verified by investigating the second-order necessary conditions. TUDIES of optimal flight generally employ dynamic models based on point-mass equations of motion. Indeed, for many purposes these models must be further simplified, leading to energy models and the like. Recently there has been considerable interest in highly agile or supermaneuver able aircraft. Maneuvers of interest in these applications include rapid fuselage pointing and suggest the need to include rigidbody rotational dynamics in optimal flight studies. The usual flight mechanics wisdom is that the rigid-body rotations are fast compared with translational motions (short-period visa-vis phugoid frequencies). This suggests a singular perturbation scheme to divide and conquer these problems. However, some of the maneuvers discovered in simulator studies are delicately choreographed so that considerable care is needed in choice of coordinates and in time-scale selection. The key point is that careful analysis is needed to sort out various effects and to gain insight into these supermaneuvers. The present study is motivated by the problem of angular velocity control for a supermaneuvera ble aircraft. In one scenario of interest, primarily the low-dynamic-pressure regime at high angles of attack, the usual aerodynamic control surfaces are ineffective for generating control moments. In such flight regimes, it is proposed that thrust-vectored propulsive controls be uspd as the primary source of control moments, which may be generated by deflecting vanes in the exhaust flowfield of a jet-powered aircraft. Owing to the location of engines, we anticipate that the rolling moment (in the body-fixed axes) that can be generated is going to be small in comparison to the pitching and the yawing moments. As an extreme case, we consider the case when the rolling moment is identically zero. Hence any rolling motion is controlled indirectly through the pitch and yaw motions. Thus we are motivated to study optimal rigid-body rotational maneuvers in the absence of rolling moments.

Subscale-model and full-scale engine mixed-flow exhaust system performance comparison

A full-scale engine test of the NASA/General Electric Company (GE) Energy Efficient Engine (E) was conducted to demonstrate the E engine concept and evaluate its performance. The test program, performed at the GE outdoor engine test facilities in Peebles, Ohio, included a detailed evaluation of the total pressure and temperature profiles at the exit of the mixed-flow exhaust system to determine its mixing effectiveness. Subscale model tests of the same mixed-flow exhaust system had been previously conducted at FluiDyne Engineering Corporation in Minneapolis as part of the GE E mixer aerodynamic technology development program. The scalemodel and full-scale engine nozzle exit survey data and the calculated mixing effectiveness are compared and discussed. The results indicate the effectiveness of the full-scale engine mixing to be 5% higher than that of the scale model as a result of a geometric difference and higher turbulence levels in the engine exhaust flowfield.

Interaction of Gun Exhaust Flowfields

I between the air column forced out ahead of a projectile and the subsequent propellant gas exhaust are observed experimentally. By comparing data taken at fully ambient conditions with that acquired in firings from an evacuated gun tube, the interactions are demonstrated to significantly influence the near muzzle flow, a region through which the projectile passes and within which muzzle devices are located.

Integrated Scramjet Nozzle/Afterbody Performance Analysis

A two-dimensiona l flowfield computer program developed for preliminary performance analysis of integrated scramjets is presented. Scramjet flowfields are computed by means of the shock-capture technique, with real gas thermodynamic properties for frozen and equilibrium compositions. The program computes internal and external flowfields with multiple shock interactions. Interaction of the exhaust (underexpanded or overexpanded) with the external stream and vehicle afterbody also is considered. Forces and moments are computed due to stream thrust and external surface pressure distributions. Typical program results for parametric nozzle/afterbody design and off-nominal performance analysis are presented. N recent years, increased attention has been given to the application of supersonic combustion ramjets (scramjets) to airbreathing launch vehicles, advanced civil aviation, and hypersonic research aircraft. One of the scramjet technology areas requiring particular attention is the interaction of the airbreathing propulsion system with the vehicle aerodynamics and the resulting effects on the overall vehicle performance. Particularly significant and as yet not well defined is the interaction of the vehicle afterbody and the exhaust of the integrated scramjet.M The integrated scramjet propulsion system utilizes the vehicle afterbody as an extension of the nozzle, producing increased effective exhaust velocities and thrust at the expense of incurred trim penalties. Therefore, a requirement exists for the means to define accurately scramjet nozzle/afterbody exhaust flowfields. A critical review of this problem indicated that detailed analysis should include consideration of real gas three-dimensional flowfields with chemical-kinetic effects, multiple shocks, heat transfer, boundary layer growth, and interaction with the inviscid flow. In view of the complexity of the problem, simplified practical methods of analysis must be developed which retain the most essential features of the flow phenomena, fulfill preliminary design requirements, and produce realistic results. At the present time, the computer programs of Refs. 5 and 6 appear to be the only published programs specifically developed to meet this requirement. These programs consider inviscid and adiabatic quasi-three-dimensional flowfields with real gas equilibrium hydrogen-air chemistry. These programs are based on the reference plane technique, with a table lookup thermodynamic properties routine. This paper represents a simpler scramjet flowfield analysis computer program suitable for preliminary design analyses. Since the majority of the current scramjet configurations are of the modular quasi-two-dimensional type, it is felt that for preliminary design .purposes the flowfield can be described sufficiently well by means of a two-dimensional program. The two-dimensional scramjet flowfield program7 considers inviscid and adiabatic supersonic flow in ducts and coflowing dissimilar streams with real or ideal gas properties. The methodology used in this program is the shock capture technique (SCT) described in Refs. 8 and 9. Thermodynamic properties are evaluated by means of a special-purpose sub

Prediction of rocket exhaust flowfields

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Comments on exhaust flow field and surface impingement

Comments on exhaust flow field and surface impingement