Total Temperature Ratio Research Articles

Sensitivity of air/mist non-equilibrium phase transition cooling to transient characteristics in a compressor of gas turbine

A comparative study of performance characteristics in a gas turbine compressor before and after water injection into the air is carried out to explore the relationships between the transient characteristic and mist injection. The Eulerian-Lagrangian particle tracking method is used for solving a process of phase transition cooling of water/air coupling, and the fast Fourier transform (FFT) is adopted to analyze the frequency attributions of fluctuating quantities with time-dependent including total pressure ratio, total temperature ratio, and efficiency. Results show, after mist injection into the compressor, droplet evaporation phase change leads to a higher characteristic fluctuation in the low-frequency region. The sensitivity of efficiency fluctuation is easily affected by the mist injection by comparing with the dry air condition. Within the mist conditions of 1–5 μm droplet sizes and 0.5–5% water/air ratios, the fluctuation of pressure in the compressor reaches a peak in small mist particle and low water/air ratio, resulting in a greater total pressure ratio. The temperature fluctuation gets the minimum in small droplet diameter and large water/air ratio, leading to a lower total temperature ratio. The fluctuation intensity of total pressure ratio, total temperature ratio and efficiency of compressor are increased to 3 times, reduced to 7/9 times and dropped to 3/5 times as water/air ratio increases from 0.5% to 5%, respectively; then, the fluctuation strength correspondingly increases 2 times, decreases 3/4 times and descends 6/7 times with the increase of droplet sizes from 1 μm to 9 μm. Thus, the fluctuation caused by low water/air ratio is higher than the small droplet diameter. Furthermore, efficiency fluctuation gets a minimum in small droplet size and appropriate water/air ratio.

Thermo-mechanical characteristics of high-speed and heavy-load modified gears with elasto-hydrodynamic contacts

Abstract The thermo-mechanical characteristics of high-speed and heavy-load modified gears in non-Newtonian transient micro thermo-elasto-hydrodynamic contacts are investigated. The synthetic stiffness model is developed from the stiffness of both the oil film and gear pair. Scuffing prediction and mesh efficiency are determined by the total contact temperature and frictional loss ratio respectively. By utilizing the verified numerical procedure, effects of geometric parameters on aforementioned thermo-mechanical characteristics are discussed. The results show that crown modification and the increase of tooth number, module, and pressure angle contribute to vibration and impact reduction of gear drive. Larger tooth number and module are beneficial to scuffing resistance and efficiency improvement. It indicates that gear system can achieve high-performance transmission by rationally matching geometric parameters.

Understanding of two-stage continuous microreaction technology for in-situ generation and consecutive conversion of diazomethane

Abstract Microreaction technology (MRT) offers a safe way to deliver useful, highly active but hazardous diazomethane on demand within a closed system and then realize in-situ conversion to the desired product. In this work, the two-stage MRT is employed for the integrated synthesis of diazomethane and its consecutive conversion with benzoic acid, and the key factors resulting in the low product yield are investigated in depth. Extensive optimization of reaction conditions (e.g., the total flow rate, residence time, temperature and reactants ratio), solvent selection as well as heat management are performed. Special attention has to be devoted to the significantly detrimental effect of water solvent on the product yield. These results provide some fundamental understanding of the interlinked reaction chemistry and engineering, which could guide the MRT development for continuous operation when using hazardous compounds.

Numerical study on supersonic combustion of hydrogen and its mixture with Ethylene and methane with strut injection

In this paper, supersonic combustion and flow field of hydrogen and its mixture with ethylene and methane from strut injections in a Mach 2 supersonic flow are studied numerically. The fuel mixture of hydrogen, methane and ethylene represents the major products of pyrolysis of hydrocarbon fuels with large molecules such as kerosene as it acts as coolant and flows through cooling channels and absorbs heat. Detached Eddy Simulation with a reduced kinetic mechanism and steady flamelet model are applied to simulate turbulent combustion. The calculated temperature profiles of hydrogen are compared to the experimental results of DLR supersonic combustor for validation of the present numerical method. The supersonic combustion flows associated with shock waves, turbulent vortices and flame structures are studied. With addition of methane and ethylene, the flame zone moves further downstream of the strut and the maximum flow temperature at chamber exit decreases by 200 K. With analysis of total temperature ratios, it is found that combustion efficiency for hydrogen combustion is 0.91 and it decreases to 0.78 for the fuel mixture. The calculation of ignition delay time and flame speed reveals that fuel mixture of hydrogen and hydrocarbons has considerably larger delay time and smaller flame speed, that contributes to the weakened flame zone and lower combustion efficiency.

Flow characteristic of a multistage radial turbine for supercritical compressed air energy storage system

Compressed air in supercritical compressed air energy storage system expand from supercritical to atmospheric conditions at lower inlet temperature (<500 K) to generate MW scale power. Therefore, a new multistage radial turbine is adopted and the flow characteristic is investigated by numerical simulation. Effects of ideal gas model and tip clearance on the performance and flow field of the multistage turbine are revealed. Results show that ideal gas model can reveal flow pattern under supercritical condition correctly while leading to obvious deviation of isentropic enthalpy drop, entropy, and inlet-to-exit total temperature ratio. Relative differences for mass flow and efficiency are less than 2%, while the relative differences for output power reaches to 9.36%. For shrouded rotor, mixing of working fluid near hub, blade suction surface, and shroud is the main influencing factor of the flow loss in the rotor. For unshrouded rotor, leakage vortex promote mixture of the fluid deriving from the hub, shroud, and suction surface, and causes much higher flow loss in the channel of rotor. The rotors, which have higher blade height variation rate, present higher efficiency reduction when the tip clearance height is increased, which is because the proportion of tip clearance in blade inlet height increases with the increase of average aspect ratio, resulting in the increase of leakage flow at the leading edge of rotor blade. The pressure fluctuation near the tip clearance and efficiency reduction is also increased. The present study provides a reference for further design and optimization of the multistage radial turbines in compressed air energy storage.

Numerical investigation of air flow in a supersonic wind tunnel

In the framework of TsAGI’s supersonic wind tunnel modernization program aimed at improving flow quality and extending the range of test regimes it was required to design and numerically validate a new test section and a set of shaped nozzles: two flat nozzles with flow Mach number at nozzle exit M=4 and M=5 and two axisymmetric nozzles with M=5 and M=6. Geometric configuration of the nozzles, the test section (an Eiffel chamber) and the diffuser was chosen according to the results of preliminary calculations of two-dimensional air flow in the wind tunnel circuit. The most important part of the work are three-dimensional flow simulation results obtained using ANSYS Fluent software. The following flow properties were investigated: Mach number, total and static pressure, total and static temperature and turbulent viscosity ratio distribution, heat flux density at wind tunnel walls (for high-temperature flow regimes). It is demonstrated that flow perturbations emerging from the junction of the nozzle with the test section and spreading down the test section behind the boundaries of characteristic rhomb’s reverse wedge are nearly impossible to eliminate. Therefore, in order to perform tests under most uniform flow conditions, the model’s center of rotation and optical window axis should be placed as close to the center of the characteristic rhomb as possible. The obtained results became part of scientific and technical basis of supersonic wind tunnel design process and were applied to a generalized class of similar wind tunnels.

Effects of Jet Temperature on Broadband Shock-Associated Noise

This paper describes an experimental study of the effects of jet heating on broadband shock-associated noise. The noise measurements are supplemented with schlieren flow visualizations that show that the appearance of a Mach disk in the shock cell structure is consistent with the saturation of the broadband shock-associated noise levels for both highly under- and overexpanded jets. The effects of jet heating are then described for the design Mach-number-1.5 nozzle operating at four Mach numbers: two overexpanded ( and 1.4), and two underexpanded ( and 1.9). Total temperature ratios in the range of 1.0 to 2.2 are considered in increments of 0.2. These values are chosen to supplement heated air measurements at higher total temperature ratios (between 1.8 to 3.2) and to examine the effects of low levels of heating. It is shown that the peak broadband shock-associated noise rapidly approaches a saturation value in all cases. In the underexpanded cases, the peak broadband shock-associated noise level decreases with increasing total temperature ratio and the reverse is true for the overexpanded cases. Although the changes are very small, a systematic trend can be observed. The spectral shape of fundamental broadband shock-associated noise is found to be very similar for all the cases considered. An empirical prediction model for the fundamental broadband shock-associated noise is proposed.

Ethylene Flame Dynamics and Inlet Unstart in a Model Scramjet

Ethylene flame dynamics in a model scramjet installed in an open test section of an arc-heated hypersonic wind tunnel are experimentally investigated with freestreams of Mach 4.5–9. An ethylene fuel jet is injected into the model scramjet that is autoignited by high-enthalpy flows compressed and decelerated by a train of shock waves and boundary layers developing in the scramjet isolator/combustor. The ethylene flame behavior is captured by a high-speed movie camera through optical access windows on the model and characterized by surface pressure/temperature traces recorded at five different locations along the flowpath through the model scramjet. Downstream pressure buildup due to combustion heat release, shock-induced compression, and mass loading via the fuel jet triggers the flame propagation toward upstream, stabilizes the flame at a location in the combustor, or pushes the flame further upstream to cause inlet unstart. In particular, flame dynamics while the scramjet undergoes the inlet unstart with Mach 4.5 freestreams are investigated intensively. In addition, stagnation temperature, overall fuel concentration, and oxygen concentration in freestream flows are varied to investigate their influences on the flame dynamics. The flame propagates faster and anchors further upstream with increased total temperature, oxygen enrichment, and overall equivalence ratio slightly above unity.

Analyzing the impact of the inlet temperature on the acoustic noise production form a supersonic jet using large eddy simulations

Non-ideal expanded supersonic jets emerging from a convergent-divergent nozzle produce three different types of noise, i.e., the shock-associated broadband noise, the screech noise, and the turbulent mixing noise. Interesting to note that the screech tone outside of the nozzle was exposed mostly in downscaled laboratory experiments, while under realistic conditions the exhaust jet of a gas turbine engine does not show this phenomena. Apart from a geometric scaling difference, usually a lower temperature is employed in experimental studies. It is believed that the screech tone occurs due to self-excitation of the shear-layer in a feedback-loop. Acoustic waves generated by vortical structures interacting with a shock are propagating upstream within the subsonic region of the shear-layer or outside of it. They eventually hit the nozzle's lip and excite instabilities at a certain frequency. The compressible Navier-Stokes equations are simulated numerically by using Large Eddy Simulation approach. The effect of supersonic jet operation temperature onto the associated noise is investigated. The jet-exit Mach-number is 1.56, while the total temperature ratios considered at the inlet plane of the nozzle are 1.27, 2.05, and 3.65. The differences in the near-field acoustics will be presented in each of the cases and the flow-acoustic interaction will be analyzed and quantified.

Crackle Noise in Heated Supersonic Jets

Crackle noise from heated supersonic jets is characterized by the presence of strong positive pressure impulses resulting in a strongly skewed far-field pressure signal. These strong positive pressure impulses are associated with N-shaped waveforms involving a shocklike compression and, thus, is very annoying to observers when it occurs. Unlike broadband shock-associated noise which dominates at upstream angles, crackle reaches a maximum at downstream angles associated with the peak jet noise directivity. Recent experiments (Martens et al., 2011, “The Effect of Chevrons on Crackle—Engine and Scale Model Results,” Proceedings of the ASME Turbo Expo, Paper No. GT2011-46417) have shown that the addition of chevrons to the nozzle lip can significantly reduce crackle, especially in full-scale high-power tests. Because of these observations, it was conjectured that crackle is associated with coherent large scale flow structures produced by the baseline nozzle and that the formation of these structures are interrupted by the presence of the chevrons, which leads to noise reduction. In particular, shocklets attached to large eddies are postulated as a possible aerodynamic mechanism for the formation of crackle. In this paper, we test this hypothesis through a high-fidelity large-eddy simulation (LES) of a hot supersonic jet of Mach number 1.56 and a total temperature ratio of 3.65. We use the LES solver CHARLES developed by Cascade Technologies, Inc., to capture the turbulent jet plume on fully-unstructured meshes.

Influence of Compressibility on Film-Cooling Performance

Film-cooling is an active cooling technique that is widely used in conventional gas turbine and rocket engines to manage heat transfer between hot reacting gases and cooler structural components. It is also a candidate technology for the thermal protection of scramjet engines, in which combustion occurs under highly compressible conditions. This paper extends semi-empirical modeling ideas for incompressible wall-jet slot film-cooling to account for both convective and thermal compressibility effects. The new model shows that the influence of compressibility is determined by the magnitude of the average convective Mach number M c , the total temperature ratio θ 0 = T 0s /T 0∞ , and the flow Mach number M HS of the highest-speed stream in the film-cooling flow. In general, increasing M c and decreasing θ 0 improves film-cooling performance. These predictions are validated via comparison with experimental data in three compressible flow regimes: weakly compressible (M c , M HS ≤ 0.3 and 0.6 ≤ θ 0 ≤ 1), moderately compressible (0.3 1.0, and θ 0 < 0.3). The model also resolves disagreements in the literature over the importance of compressibility in film-cooling problems by showing that compressibility effects can be significant, provided the convective and flow Mach numbers are high enough and the total temperature ratio is low enough.

Multiple Surrogate Modeling for Axial Compressor Blade Shape Optimization

A major issue in surrogate model-based design optimization is the modeling fidelity. An effective approach is to employ multiple surrogates based on the same training data to offer approximations from alternative modeling viewpoints. This approach is employed in a compressor blade shape optimization using the NASA rotor 37 as the case study. The surrogate models considered include polynomial response surface approximation, Kriging, and radial basis neural network. In addition, a weighted average model based on global error measures is constructed. Sequential quadratic programming is used to search the optimal point based on these alternative surrogates. Three design variables characterizing the blade regarding sweep, lean, and skew are selected along with the three-level full factorial approach for design of experiment. The optimization is guided by three objectives aimed at maximizing the adiabatic efficiency, as well as the total pressure and total temperature ratios. The optimized compressor blades yield lower losses by moving the separation line toward the downstream direction. The optima for total pressure and total temperature ratios are similar, but the optimum for adiabatic efficiency is located far from them. It is found that the most accurate surrogate did not always lead to the best design. This demonstrated that using multiple surrogates can improve the robustness of the optimization at a minimal computational cost.

축류압축기 블레이드의 표면조도가 성능에 미치는 영향

Deterioration of surface of turbomachinery blades occurs in course of time due to many factors and hence reduces the performance of the machine. In this paper, the effects of surface roughness of transonic axial compressor blade on performance are studied considering a reference blade and a shape distorted (optimized) blade. Optimal blade is designed considering sweep and lean. Baldwin-Lomax turbulence model is used for flow field analysis and Cebeci-Smith roughness model is formulated for roughness modeling. It is found that, as the surface roughness increases, adiabatic efficiency, total temperature ratio and total pressure ratio decrease while Mach number increases. Performance deterioration is more severe in case of distorted blade as compared to reference blade.

The effects of intercooling and regeneration on the thermo-ecological performance analysis of an irreversible-closed Brayton heat engine with variable-temperature thermal reservoirs

A thermo-ecological performance analysis of an irreversible intercooled and regenerated closed Brayton heat engine exchanging heat with variable-temperature thermal reservoirs is presented. The effects of intercooling and regeneration are given special emphasis and investigated in detail. A comparative performance analysis considering the objective functions of an ecological coefficient of performance, an ecological function proposed by Angulo-Brown and power output is also carried out. The results indicate that the optimal total isentropic temperature ratio and intercooling isentropic temperature ratio at the maximum ecological coefficient of performance conditions (ECOPmax) are always less than those of at the maximum ecological function () and the maximum power output conditions () leading to a design that requires less investment cost. It is also concluded that a design at ECOPmax conditions has the advantage of higher thermal efficiency and a lesser entropy generation rate, but at the cost of a slight power loss.

Mach Number Effects on Jet Noise Sources and Radiation to Shallow Angles

LL jets have noise generation due to turbulent mixing; suchmixingleadstotheproductionofReynoldsstresses,thesourceterm within the Lighthill acoustic analogy [1]. In addition to thissource,ajetthatpossessescoherent,large-scalestructuresthattravelwith velocities exceeding the local speed of sound will emit Machwave radiation, a highly efficient noise production mechanism. Theamplitudeofjetnoisescaleswiththejetareaandtheexitvelocitytoanangle-dependentpowern.Thereisdisagreement,however,inthevalueofn.Basedondimensionalanalysis,Lighthill[1]foundntobe8. In anextensive examination of jet noise scaling, Viswanathan [2]observed that n 9:8 for the measurement location (30 deg) andambient to total temperature ratio of unity examined in this work.Venkatesh et al. [3] presented an extensive set of subsonic jet noisemeasurements that show the region of maximum noise generationcoincides withthecollapseoftheuniform, irrotationaljetcore.Thistechnical note examines the impact of Mach number on themechanism of jet noise generation, the scaling of jet noise, and theregion of maximum noise generation. It does this throughexamination of the time and frequency domain properties of theshallow angle jet noise from Mach 0.9, and ideally expandedMach 1.3 and 2.0 jets.A recent set of experiments showed that a Mach 1.3 jet possesseslarge amplitude acoustic peaks within the time-domain data andthesepeakscouldbeusedtorelatethegenerationoffarfieldradiatedsound to the evolution of the mixing layer [4–7]. These acousticpeaks are examined in this technical note to characterize the noisegeneration mechanism and source region of the three jets. For allMachnumbersstudied,thelargeamplitudeacousticpeaksoriginatefrom an ellipsoid-shaped region that is centered past the end of thepotentialcoreandthenoisesourceregionnarrowsinthecross-streamwith increasing Mach numbers. Unlike at Mach 0.9 and 1.3, theMach2.0acousticpeakshavetemporalasymmetry,whichisrelatedtothechangeinnoisegenerationmechanismfromturbulencemixingto Mach wave emission.

기체유동에서 발생하는 복합초킹 현상에 관한 연구

Compound choking frequently occurs at a minimum area of the flow passage, where two or more streams which have different stagnation properties are merged. This phenomenon is especially important in that the flow choking may not be given by Mach number, M=l at the nozzle throat. In order to obtain a detailed understanding of the flow characteristics involved in the compound flow choking, the two-dimensional, compressible, Navier-Stokes equations are solved using a fully implicit finite volume method and the predicted solutions are compared with the results of the one-dimensional theoretical analysis. Stagnation pressure and temperature of each stream are changed to investigate the effects on the compound choking. The results show that stagnation pressures of each stream affect Mach number and static pressure distributions downstream of the exit of the convergent nozzle. However, the flow characteristics of the compound choking are not significantly dependent on the total temperature ratio.

Aeroacoustic Measurement of Transient Hot Nozzle Flows

An assessment of a low-cost transient testing technique for obtaining acoustic measurements from hot, highspeed nozzle e ows is presented. A shock tunnel was used to produce short-duration (10‐20 ms) supersonic air jets from three sizes of American Society of Mechanical Engineers conic nozzles and a mixer-ejector nozzle typical of a high-speed civil transport (HSCT) aircraft application. A comparison between shock-tunnel transient noise data and steady-state data is presented. Theassessment establishes the uncertainty bounds on sound pressure level measurements over the range of frequency bands, nozzle pressure and total temperature ratios, and nozzles scales for which the facility can be used as a substitute and/or complementary mode of investigation for steady-state hot-e ow test facilities. Using the transient facility, far-e eld narrowband spectra were obtained at directivity angles from 65 to 145 deg, and the data were extrapolated to full-scale e ight conditions. The constraint of short test duration is alleviated through the use of multiple runs to reduce the uncertainty associated with transient acoustic measurements. Sound pressure level vs frequency trends with nozzle pressure ratio and directivity angle are comparablebetweenthesteady-stateandtransientdatafortheconicandmixer-ejectornozzles.Conicnozzleresults demonstrated that thetransient noise data replicatethe steady-state data to within §2-3 dB, and themagnitude of effectiveperceived noiselevel valuesagree to within 1 ‐3 dB depending on test condition and nozzlesize. The mixerejector model demonstrated agreement with the steady-state noise data of around 2 ‐5 dB on sound pressure level vs full-scale frequency over the range of nozzle pressure ratios and total temperature ratios relevant to the HSCT mixer-ejectordevelopmentprogram.Trendsinnoisevariationwithazimuthalangle,asmeasuredcircumferentially about the inlet axis, as well as variation with mixer area ratio were also captured using transient testing.

Simultaneous coherent anti-Stokes Raman spectroscopy measurements inhydrogen-fueled supersonic combustion

Coherent anti-Stokes Raman spectroscopy (CARS) has been used to measure H2 and H2O concentrations and N2 temperature simultaneously in a H2-fueled, air fed supersonic combustor. Profiles of each parameter at selected axial locations were generated by translating the measurement point vertically across the combustor at varying conditions of total temperature and equivalence ratio. The baseline combustor experiment contributes to an understanding of mixing and combustion processes and a data base for refining 2-D CFD codes. The combustor, the dual Stokes CARS system and the techniques employed are described. Sample results and potential improvements are also discussed.

Laminar-Turbulent Transition in Free Shear Layers

A STUDY has been performed to derive a threshold criterion for the existence of turbulence in a shear layer, and to see if such a threshold is useful in predicting transition onset in such a flow. This approach has already given satisfactory results in wake flows. As in the latter case, the present work is based on the existence of a minimum turbulence Reynolds number, and derives numerical results by utilizing established magnitudes of turbulence properties in the turbulent shear layer. Transition Reynolds numbers are predicted as a function of speed ratio, total temperature ratio, and Mach number jumps across the layer. New experiments are described in which free shear-layer transition was measured and, within an observed unit Reynolds-number effect, was found to parallel the predictions quite well.

Altitude Engine Test of a Turbofan Exhaust Gas Mixer to Conserve Fuel

A comparison of the specific fuel consumption was made with and without an internal mixer installed in a low bypass ratio, confluent flow turbofan engine. Tests were conducted at several Mach numbers and altitudes for core to fan stream total temperature ratios of 2.0 and 2.5 and mixing lengths of L/D = 0.95 and 1.74. For these test conditions, the specific fuel consumption improvement varied from 2.5 to 4.0 percent.