Search

Several ten-stages-compressor of a high bypass ratio turbofan engine was used as the research object, the change law of tip clearance and its effects on aerodynamic performance and stable boundary of the compressor were investigated based on a high speed compressor component, core engine and complete engine test facility. The mechanism and the sensitive area of tip clearance effect on the overall performance and each stage performance of multistage compressors was revealed by detailed simulation results of transonic and subsonic stage performance. The results show that the clearance change law of compressors of different sizes is consistent, which is mainly related to centrifugal force and temperature filed. Flow-pressure ratio and flow-efficiency characteristics of the compressor at design rotational speed Indicate a feature of whole reduction while the clearance increases, and the clearance has more impact on the back subsonic stage. With the increase of the tip clearance, the efficiency of the front transonic stages of the compressor does not decrease for the small ratio of clearance to the big blade height, but the work capacity of the back subsonic stage decreases obviously, and the influence of the clearance mainly concentrates on the back subsonic stage, which makes the compressor efficiency and stall margin droop. [ABSTRACT FROM AUTHOR]

Tracer pathline statistics have rarely, if at all, been used in the investigation of compressible subsonic, transonic and supersonic flows past a circular cylinder. Here, we report pathline histogram, conditional histogram and zonal histogram data, focusing on the characteristics of particle instantaneous flow reversal and their dependencies on Mach number and tracer release position. The far-upstream Mach number and the diameter-based Reynolds number used for these three flow regimes are $ ({0.2, 10000}), ({0.9, 3900}) $ (0.2 , 10000) , (0.9 , 3900) and $ ({1.2, 10000}) $ (1.2 , 10000) , respectively. The subsonic surface-release histogram profiles display unexpected persistent oscillations at a frequency $ 50 $ 50 times of the periodic vortex shedding frequency, reflecting the effect of a very small magnitude zig-zagging type of particle motion in a region very close to the wall and slightly behind the mean flow separation location. No similar high-frequency oscillation is observed in the corresponding transonic and supersonic histograms likely because, at $ \textrm{Ma} \ge 0.9 $ Ma ≥ 0.9 , periodic vortex shedding ceases to exist and the near-wake becomes a quasi-laminar recirculation zone bounded by a pair of converging slip-layers with a neck opening to the far-wake. At the instant of $ 5 $ 5 subsonic vortex-shedding periods after being released from the cylinder surface, there are $ 1.67\% $ 1.67 % , $ 35.2\% $ 35.2 % and $ 39.7\% $ 39.7 % particles that experience flow reversal in the subsonic, transonic and supersonic flows, respectively. Surprisingly, in the present transonic flow, growth of the turbulent far-wake behind the pair of lambda-shocks is constrained over a considerable streamwise distance. [ABSTRACT FROM AUTHOR]

To investigate the flow and drag characteristics of the ramjet flow path of a hybrid over-under turbo based combine cycle （TBCC），TBCC engine under transonic flight conditions， a high-Mach aircraft model with a cruise Mach number of 4.0 based on hybrid over-under TBCC engine was designed， and its performance in Ma∞=0.7—1.6，H∞=11 km to simulate the aircraft-engine internal/external flow and its coupling characteristics in the flight environment. The results show that in the whole transonic state， the static pressure of the air pressurized at the inlet of the ramjet inlet reaches 85%—90% of the free flow stagnation pressure， and the air flow is close to the stagnation state， indicating that there is a strong throttling flow effect in the combined inlet， and the throat of the ramjet flow path is the main contributor to the throttling effect of the combined inlet port. The exhaust flow of the ramjet tail nozzle is simultaneously interfered by the aircraft bypass flow and the exhaust system of the turbine channel， and the turbine channel exhaust jet has various interference mechanisms such as expansion compression and exhaust injection to the airflow of the ramjet tail nozzle itself. The drag analysis shows that the differential-pressure drag coefficient is two orders of magnitude higher than the friction drag coefficient of the inner surface， is the main source of the drag of the ramjet engine under the transonic state. In the subsonic state， the inlet port drag accounts for 60%—80%，which is the main drag component of the ramjet engine， in the Ma∞ >1.0 supersonic state， the inlet port drag ratio decreases with the flight Mach number and decreasing gradually， and the drag of the tail nozzle increases rapidly， indicating that the drag contribution gradually shifts to the tail nozzle， and the two tend to approach. [ABSTRACT FROM AUTHOR]

Flow over a cavity is one of the most intriguing problems in aeronautics. Although geometry is simple, the physics of cavity requires uttermost attention. In this study, various novel passive flow control techniques such as reshaping the aft wall as stair-stepped configuration or combinations of spoilers with reshaped aft wall are applied to cavity and effects of these techniques are investigated numerically. Combined configurations that are proposed in the present study are considered as a novelty to the literature. Analyses are performed with Detached Eddy Simulation method three-dimensionally in transonic regime (0.85 Mach) for a Reynolds number of ~107 based on the cavity length, 0.508 m, using open-source software OpenFOAM. Results are compared with both experimental data and each other, fundamentally in terms of Overall Average Sound Pressure Level. Further examinations are also performed for features such as Mach number, turbulent intensity and turbulent coherent structures. It is seen that combined passive flow control methods have reduced Overall Average Sound Pressure Level by ~10 dB. Newly proposed passive flow control methods have also reduced Overall Average Sound Pressure Level by ~6 dB. A high correlation between coherent turbulent structures and generated noise is observed. [ABSTRACT FROM AUTHOR]

Purpose: In the present research, the effect of the flexible shells method in unsteady viscous flow around airfoil has been studied. In the presented algorithm, due to the interaction of the aerodynamic forces and the structural stiffness (fluid-structural interaction), a geometrical deformation as the bump is created in the area where the shock occurs. This bump causes instead of compressive waves, a series of expansion waves that produce less drag and also improve the aerodynamic performance to be formed. The purpose of this paper is to reduce wave drag throughout the flight range. By using this method, we can be more effective than recent methods throughout the flight because if there is a shock, a bump will form in that area, and if the shock does not occur, the shape of the airfoil will not change. Design/methodology/approach: In this simulation pressure-based procedure to solve the Navier-Stokes equation with collocated finite volume formulation has been developed. For this purpose, a high-resolution scheme for fluid and structure simulation in transonic flows with an arbitrary Lagrangian-Eulerian method is considered. To simulate Navier-Stokes equations large eddy simulation model for compressible flow is used. Findings: A new concept has been defined to reduce the transonic flow drag. To reduce drag force and increase the performance of airfoil in transonic flow, the shell can be considered flexible in the area of shock on the airfoil surface. This method refers to the use of smart materials in the aircraft wing shell. Originality/value: The value of the paper is to develop a new approach to improve the aerodynamic performance and reduce drag force and the efficiency of the method throughout the flight. It is noticeable that the new algorithm can detect the shock region automatically; this point was disregarded in the previous studies. It is hoped that this research will open a door to significantly enhance transonic airfoil performance. [ABSTRACT FROM AUTHOR]

The purpose of this study is to analyze the subsonic and transonic water entry of the axisymmetric body with compressibility effect considered. The water-entry processes of axisymmetric body are numerically investigated, with Tait state equation and energy equation used to describe the compressibility of the liquid. Reasonable agreement has been obtained between the numerical results and the experimental measurements of trajectory, velocity and acceleration. The cavity evolution and dynamic characteristics of subsonic cases are investigated for vertical and oblique water-entry. The results show that the maximum drag coefficient of the oblique water entry is lower, and it is affected by a large moment which causes pitch angle and trajectory inclination angle to change significantly. In the oblique transonic conditions, the effect of compressibility of the liquid on the dynamic flow characteristics is analyzed. On this basis, the influence of entry angle on the hydrodynamic characteristics is investigated. It is concluded that the water-entry angle has a more significant influence on the trajectory characteristics. • Investigate the flow structure and dynamic characteristics of subsonic water-entry cases for vertical and oblique water-entry. • Analyze the effect of compressibility of the liquid on the dynamic flow characteristics of transonic water-entry cases. • Improve the understanding of the influence of entry angle on the hydrodynamic characteristics with compressibility considered. [ABSTRACT FROM AUTHOR]

The transient shock dynamics and drag characteristics of a projectile flying through a pipe 3.55 times larger than its diameter at transonic speed are analyzed by means of time-of-flight and pipe wall pressure measurements as well as computational fluid dynamics (CFD). In addition, free-flight drag of the 4.5-mm-pellet-type projectile was also measured in a Mach number range between 0.5 and 1.5, providing a means for comparison against in-pipe data and CFD. The flow is categorized into five typical regimes the in-pipe projectile experiences. When projectile speed and hence compressibility effects are low, the presence of the pipe has little influence on the drag. Between Mach 0.5 and 0.8, there is a strong drag increase due to the presence of the pipe, however, up to a value of about two times the free-flight drag. This is exactly where the nose-to-base pressure ratio of the projectile becomes critical for locally sonic speed, allowing the drag to be estimated by equations describing choked flow through a converging-diverging nozzle. For even higher projectile Mach numbers, the drag coefficient decreases again, to a value slightly below the free-flight drag at Mach 1.5. This behavior is explained by a velocity-independent base pressure coefficient in the pipe, as opposed to base pressure decreasing with velocity in free flight. The drag calculated by CFD simulations agreed largely with the measurements within their experimental uncertainty, with some discrepancies remaining for free-flying projectiles at supersonic speed. Wall pressure measurements as well as measured speeds of both leading and trailing shocks caused by the projectile in the pipe also agreed well with CFD. [ABSTRACT FROM AUTHOR]

Abstract A new unsteady aerodynamic modeling methodology for calculating transonic flutter characteristics is presented. The main idea of the methodology is to obtain the unsteady flow response to small amplitude periodic deformations of a structure over a large range of oscillation frequencies through the interpolation of the most dominant fluid dynamic modes obtained from Dynamic Mode Decomposition (DMD) of a few reference unsteady simulations at different oscillation frequencies. These simulations can be carried out by solving the Euler or RANS equations. The methodology can then be used to obtain a frequency-domain generalized aerodynamic force matrix, and stability analysis can be performed using standard flutter analysis methods such as the p - k method. The proposed methodology provides a very good estimate of the flutter boundary for the 2D Isogai airfoil and 3D AGARD 445.6 wing models, but at a lower computational cost than the traditional higher-fidelity Fluid–Structure Interaction (FSI) simulations. [ABSTRACT FROM AUTHOR]

Abstract For a high-load low-reaction compressor transonic rotor based on an increase in outlet axial velocity, the stage reaction is decreased by a substantial meridian contraction of the endwall. The axisymmetric endwall contouring has a deep impact on the aerodynamic performance of the rotor. Therefore, nine combinations of endwall contours with three different shapes including a linear wall, a sinusoidal wall and a symmetrical wall of the sinusoidal wall about the linear wall were firstly studied through 3D steady Reynolds Average Navier–Stokes simulations. The results show that the cases with a linear hub have the highest peak pressure ratio and the largest peak efficiency, whereas the cases, featured with a sinusoidal hub, have the widest stall margin, the largest mass flow and the lowest pressure ratio. Endwall contouring not only influences the shock structure across the flow passage but also dominates the location and intensity of separation. In addition, the effect of endwall contouring on loading distribution is also presented and the case with a sinusoidal hub is characterized by aft-loading and the other two cases have a feature of fore-loading. [ABSTRACT FROM AUTHOR]

The aerodynamic performance of a cylinder Pitot probe for velocity measurements in compressible non-ideal gas flows, such as those encountered in Organic Rankine Cycle (ORC) turbines, is investigated by means of Computational Fluid Dynamics. Numerical simulations are performed at subsonic and transonic conditions, and freestream Reynolds numbers are in the cylinder critical regime. The working fluid is the organic vapour Novec™ 649. Air flow simulations at similar inlet conditions are reported for comparison. Steady and unsteady RANS solutions are computed with the Spalart–Allmaras turbulence model. The results are assessed against experimental measurements collected in a wind tunnel. URANS is in good agreement with experimental data for all considered conditions, and delivers reasonably accurate estimations of the cylinder back pressure. Using a dense gas leads to a lower minimum pressure coefficient compared to air, alongside a reduced maximum Mach number due to the non-ideal speed of sound behaviour. In the experimentally studied range of compressibility factors and Mach numbers, discrepancies observed with respect to air flow are mostly an effect of the different isentropic exponents. In the transonic regime, shock waves causing boundary layer separation are weakened in the dense gas, but back pressure is also decreased, contributing to rising form drag. • In subsonic regimes, weakly non-ideal conditions essentially do not affect the flow. • The back pressure prediction is improved by URANS over RANS for transonic conditions. • Strongly non-ideal thermodynamic conditions cause the gas to expand more. • In diluted conditions, differences between flows are due to the isentropic exponent. [ABSTRACT FROM AUTHOR]

• Simulations of transonic non-Newtonian, wave-assisted slurry atomization in novel injector. • Unraveled visualization of annular wave cycle. • Nozzle exit marks dramatic increase in azimuthal variation of slurry waves. • Instabilities are excited in self-amplification cycle. • Adaptive mesh refinement found to be insufficient for this modeling application. The onset of three-dimensional instabilities during "Wave-Augmented Varicose Explosions" (WAVE) liquid disintegration is characterized for an annular flow of a shear-thinning slurry that is fed into a central transonic steam flow. Droplet production inside the nozzle is enhanced by ligaments radially flicking up from the slurry wave into the steam flow with radial:axial velocity ratios exceeding 0.5. The wave also leaves residual ligaments in its wake, which facilitate further disintegration. After birth, a wave spends 80% of the wave cycle period building up to peak height at the nozzle exit. Two effervescent mechanisms are provided as 1) steam penetrates the rising wave and surface deformation allows steam fingers to force through, and 2) the wave collapses on itself, trapping steam. Baroclinic torque drives the development of Rayleigh-Taylor (RT) instabilities and reaches values on the order of 1 × 10 13 1/s2. Both RT and Kelvin-Helmholtz instabilities are self-amplified in a viscosity-shear-temperature instability cycle because the slurry is non-Newtonian. Velocities inside the nozzle (wave formation region) are generally azimuthally similar (two-dimensional), but those outside the nozzle (radial bursting region) are azimuthally uncorrelated (three-dimensional). Inter-variable correlations show significant decoupling of quantities beyond the nozzle exit, and local strain rate fluctuations were found to correlate particularly well with bulk system pulsation. Although adaptive mesh refinement (AMR) can provide computationally efficient resolution of gas–liquid interfaces, this technique produced different results than an equivalent non-dynamic mesh when modeling WAVE. Gradients were particularly affected by AMR, and turbulent kinetic energy showed differences greater than 150% outside the nozzle. [ABSTRACT FROM AUTHOR]

We present an optimization platform for turbomachinery with complex mesh configuration in a parallel computation environment. A continuous adjoint solver for 3-D viscous internal flow is coded under the same parallel framework as the flow solver. To meet the various permitted extents of reshaping on blade surface and to cut down the computational cost in grid perturbation, a localized two-level mesh deformation method is developed based on Gaussian radial basis function (RBF). This method works efficiently for both the O mesh surrounding the blade and the O–H mesh inside tip gap. In the optimization of the transonic NASA Rotor 67 for high adiabatic efficiency with a mass flow rate constraint, an adjoint sensitivity analysis is conducted. The relations between the design sensitive regions and physical phenomena in internal flow are discussed. Flow fields before and after the adjoint optimization are investigated, including shock system, tip leakage flow, and flow separation. [ABSTRACT FROM AUTHOR]

The objective of this paper is to address the simulation of supercavitating flows around a slender body projectile at the sub-, trans- and supersonic speeds with the water compressibility effects considered. Based on the slender body theory (SBT) and the matched asymptotic expansion method (MAEM), the integro-differential equations for the cases of M a <1 and M a >1 are derived and solved. The results show that there is a good agreement of the supercavity shape for both the cases of M a <1 and M a >1 between the numerical simulation and the experimental results. The influence of Mach number on the supercavity length L , maximum radius R , slenderness λ , the corresponding growth rate of the supercavity length L ˜ , the growth rate of the maximum supercavity radius R ˜ , and the position of the supercavity maximum section (PSMS) are discussed. It can be found that L ˜ is much bigger than R ˜ , due to the large λ . It also concludes that the front part of the supercavity is thicker than the rear part, and the PSMS moves forward further at supersonic speed. Finally, the reason for the supercavity shape transition around M a = 1 is illustrated and it may be the c a use of the extension of perturbation zones at transonic speed. [ABSTRACT FROM AUTHOR]

An adjoint-response surface method is developed to provide efficient surrogate model in a parametrized design space for aerodynamic optimization of turbomachinery blades. Our goal is to improve the adiabatic efficiency or equivalently reduce the entropy generation through blade row with a mass flow rate constraint. Firstly, an aerodynamic sensitivity analysis is conducted with a viscous adjoint method to find the suitable number of control points on the suction surface of the transonic NASA rotor 67. Then quadratic polynomial (QP) based response surfaces of 4, 6 and 8 parameters are examined to validate the advantages of the gradient-enhanced model. In the following 24-parameter aerodynamic design optimization case, a steepest descent optimization (SDO) based on adjoint gradient is conducted, then QP based response surface model is constructed using both the values of cost function and its adjoint gradients with respect to geometry control parameters. We present the geometric features, overall aerodynamic improvements and flow details of optimal designs given by SDO and gradient-enhanced response surface model (GERSM). The effects of blade reshaping on shock system, tip clearance flow and flow separation at hub are examined. Also, off-design performances are analyzed regarding both adiabatic efficiency and stall margin. [ABSTRACT FROM AUTHOR]

Development of casing treatment geometries which will enhance the stable operating envelope of the compressor stage without sacrificing the efficiency is a great challenge for the designers. Bend skewed casing treatment with relatively lower porosities in the presence of plenum chamber is proven to be capable of improving the stall margin and gain in the compressor stage efficiency compared to other conventional casing treatments. Experimental investigations presented in this paper show very encouraging results with substantial improvement in the stable margin of the compressor stage with increase in the stage efficiency. The three bend skewed casing treatments designed for the experimental evaluations have porosities of 21%, 33% and 45%. The steady and transient performances of these three casing treatment geometries are evaluated at two axial extensions of 20% and 40%. Flow survey at the rotor exit shows distinct behaviors of the aerodynamic parameters in the tip region at near stall flow condition. The unsteady velocity measurements are performed with hot wire probe in the tip region ahead of the rotor leading edge to capture the fluctuations in the inlet axial velocity and the results are presented at the stall flow condition. The presence of the casing treatment changes the stalling behavior of the compressor stage. Highest porosity casing treatment results in 40.62% improvement in the stall margin with 0.88% improvement in the compressor stage efficiency at 20% axial extension. [ABSTRACT FROM AUTHOR]

The observation of distinct velocity ‘plateaus’ below the upper limits of the sonic velocity have frustrated many in the scientific community studying high-velocity dislocation dynamics. Liebfried and Frank derived the well-established elastic models of dislocation motion, showing dislocation core energy approaches a relativistically infinite level at the limiting sonic velocity due to a singularity. Eshelby predicted the possibility of a single stable ‘transonic’ velocity with the Peierls-Nabarro model, however he could not provide a physical mechanism to explain the acceleration from the sub-sonic velocity limit. Weiner proposed a linear elastic model where the atomic masses within a dislocation move in a coordinated (or harmonic) manner due to coupled momentum transfer, and used this to predict the limiting velocity at 0 K. In modern times, detailed phonon dissipation models have been developed to predict the dislocation velocity relationships at higher temperatures. However, few studies have been performed to show how phonon interactions with dislocations influence the atomistic behaviour. This paper presents a conceptual model based on the coupled motion of multiple atoms in the dislocation core. When compared with molecular dynamics simulations, the model predicts key qualitative and quantitative metrics, such as the lower and upper limiting dislocation velocities. Detailed atomistic analysis confirms the in-phase coordination of atoms that become disrupted by interactions with dislocations at both limiting velocities. The model provides a physically realistic mechanism that is capable of explaining the observation of subsonic and transonic velocity plateaus. [ABSTRACT FROM AUTHOR]

Some advanced aircrafts, especially those designed with using moderate aspect ratio, thin airfoil, trim, serrated, folding wings etc, are prone to encounter transonic abruptwing-stall, and furthermore bring out uncommanded lateral motion. To study the problem, due to the transonic abrupt wing stall, the free-to-roll test system was developed in the FL-3 Wind Tunnel. The static and free-to-roll wind tunnel tests for some aircraft models were conducted, including different configurations and different Mach number (0.8 ~ 0.95). The analysis used both the static and the dynamic wind test data to predict the abrupt wing stall sensitive range for the different configurations, and assess the corresponding severity of the uncommanded Lateral Motion. The characteristics of the uncommanded lateral motion caused by abrupt wing stall are predicted effectively at A VIC Aerodynamic Research Institute. The present study verifies the reliability of the testing facilities, establishes evaluation methods, and achieves transonic free-toroll test technique. [ABSTRACT FROM AUTHOR]

Detached-Eddy Simulations of flows in weapon bays with a generic store at different positions in the cavity and with flexible fins are presented in this paper. Simulations were carried out to better understand the fluid–structure interactions of the unsteady, turbulent flow and the store. Mach and Reynolds numbers (based on the missile diameter) were 0.85 and 326.000 respectively. Spectral analysis showed few differences in the frequency content in the cavity between the store with rigid and flexible fins. However, a large effect of the store position was seen. When the store was placed inside the cavity, the noise reduction reached 7 dB close to the cavity ceiling. The closer the store to the carriage position, the more coherent and quieter was the cavity. To perform a more realistic simulation, a gap of 0.3% of the store diameter was introduced between the fin root and the body of the store. Store loads showed little differences between the rigid and flexible fins when the store was inside and outside the cavity. With the store at the shear layer, the flexible fins were seen to have a reduction in loads with large fluctuations in position about a mean. Fin-tip displacements of the store inside the cavity were of the range of 0.2% of the store diameter, and in the range of 1–2% of store diameter when at the shear layer. [ABSTRACT FROM AUTHOR]

Eight turbulence models frequently used in aerodynamics have been employed in the detailed numerical investigations for transonic flows in the Sajben diffuser, to assess the predictive capabilities of the turbulence models for shock wave/turbulent boundary layer interactions (SWTBLI) in internal flows. The eight turbulence models include: the Spalart-Allmaras model, the standard model, the RNG model, the realizable model, the standard model, the SST model, the model and the Reynolds stress model. The performance of the different turbulence models adopted has been systematically assessed by comparing the numerical results with the available experimental data. The comparisons show that the predictive performance becomes worse as the shock wave becomes stronger. The model and the SST model perform much better than other models, and the SST model predicts a little better than the model for pressure on walls and velocity profile, whereas the model predicts a little better than the SST model for separation location, reattachment location and separation length for strong shock case. [ABSTRACT FROM AUTHOR]

The production of k-ω within standard Wilcox k-ω turbulence model is dependent on the mean strain-rate of the flow and does not take the rotation rate into account, resulting in the prediction of a weak vortex which can’t be sustained and diffuses quickly. A suitable sensor is defined to distinguish between shear layers and vortex cores, to increase the production of the dissipation rate (ω)within regions of highly rotational flow. k-ω with Pω enhancer turbulence model is applied to RANS solver for structured grid. Subsonic and transonic flows around a 65 o swept leading edge delta wing are simulated numerically. The computational results are compared with detailed experimental data obtained from a series of wind tunnel tests in the NTF at NASA and DNW-TWG at DLR. The comparisons show that the k-ω with Pω enhancer turbulence model produces better results when it comes to capturing the vortex core, the critical angle of attack and the position of vortex breakdown at subsonic and transonic flows, thus is suitable for simulating the vortical flow around delta wing. [ABSTRACT FROM AUTHOR]

Utilizing recent advancements in computational schemes for compressible, multiphase flows, this work features a parametric study of a pure liquid jet in supersonic crossflow that involves simulating the atomization process for four values of momentum-flux ratio. These simulations are validated against experimental results measured with high-speed X-ray imaging, which confirm the accuracy of the numerical approach. Also, the effect of numerical resolution on some flow behavior is investigated, revealing convergence of the jet shape and surface instability wavelength. Analysis of the resulting sprays includes statistical descriptions of the liquid distribution, liquid structures created through breakup, interfacial instabilities, and dominant flow features. As the flowrate increases, the spray penetrates further, it becomes more disperse, and less liquid impacts the wall, but the droplet size distribution changes little. The wavelength of instabilities on the windward side of the jet diverges from measured trends in subsonic crossflows. In a visualization of the time-averaged flow, counter-rotating vortices are observed along the jet core and in the wake, affecting the process of primary atomization and early droplet trajectories. • Large Eddy Simulations of liquid jet in supersonic crossflow at high Weber number. • Quantitative validation against experimental results using equivalent path length. • Shear-driven instability wavelengths differ from subsonic crossflow analogy. • Analysis of spray statistics and mean flow features. [ABSTRACT FROM AUTHOR]

Operational Modal Analysis (OMA) is a promising candidate for flutter testing and Structural Health Monitoring (SHM) of aircraft wings that are passively excited by wind loads. However, no studies have been published where OMA is tested in transonic flows, which is the dominant condition for large civil aircraft and is characterized by complex and unique aerodynamic phenomena. We use data from the HIRENASD large-scale wind tunnel experiment to automatically extract modal parameters from an ambiently excited wing operated in the transonic regime using two OMA methods: Stochastic Subspace Identification (SSI) and Frequency Domain Decomposition (FDD). The system response is evaluated based on accelerometer measurements. The excitation is investigated from surface pressure measurements. The forcing function is shown to be non-white, non-stationary and contaminated by narrow-banded transonic disturbances. All these properties violate fundamental OMA assumptions about the forcing function. Despite this, all physical modes in the investigated frequency range were successfully identified, and in addition transonic pressure waves were identified as physical modes as well. The SSI method showed superior identification capabilities for the investigated case. The investigation shows that complex transonic flows can interfere with OMA. This can make existing approaches for modal tracking unsuitable for their application to aircraft wings operated in the transonic flight regime. Approaches to separate the true physical modes from the transonic disturbances are discussed. [ABSTRACT FROM AUTHOR]

For steady supersonic flow past a solid convex corner surrounded by quiescent gas, if the pressure of the upcoming supersonic flow is lower than the pressure of the quiescent gas, there may appear a strong shock to increase the pressure and then a transonic characteristic discontinuity to separate the supersonic flow behind the shock-front from the still gas. In this paper, we prove the global existence, uniqueness, and stability of such flow patterns under suitable conditions on the upstream supersonic flow and the pressure of the surrounding quiescent gas, for the two-dimensional steady complete compressible Euler system. Mathematically, a global weak solution to a characteristic free boundary problem of hyperbolic conservation laws is constructed and shown to be unique and stable under the framework of front tracking method. [ABSTRACT FROM AUTHOR]

Large area expansion can be achieved effectively by wide angle diffusers in wind tunnels, while the airflow separation may occur at the same time. In order to ensure the flow field quality, the separation must be suppressed. Utilizing damp-screens with reasonable parameters is a valid method in engineering. With the background of 0.6m continuum transonic wind tunnel, so as to test the engineering design method of damp-screens, we verify the preliminary result by the method through the way of numerical simulation. From the calculation, by taking the scheme 3-4, loss coefficients for two screens are 1.6 and 1.0 respectively, the RMS value at the outlet of the wide angle diffuser will be 14. 5% ; considering the influence by the environment and adjusting the two coefficients to 0. 8 and 1. 0, the RMS value will be 16. 2%. The research indicates that, the engineering design method, in addition to numerical simulations, can be effectively applied in the design of wide angle diffusers. Suppressing the airflow separation, reducing the pressure loss, and improving the exit velocity uniformity can be achieved simultaneously. [ABSTRACT FROM AUTHOR]

In this paper, we study the stability of contact discontinuities that separate a C supersonic flow from a static gas, governed by the three-dimensional steady non-isentropic compressible Euler equations. The linear stability problem of this transonic contact discontinuity is formulated as a one-phase free boundary problem for a hyperbolic system with the boundary being characteristics. By calculating the Kreiss-Lopatinskii determinant for this boundary value problem, we conclude that this transonic contact discontinuity is always stable, but only in a weak sense because the Kreiss-Lopatinskii condition fails exactly at the poles of the symbols associated with the linearized hyperbolic operators. Both of the planar and nonplanar contact discontinuities are studied. We establish the energy estimates of solutions to the linearized problem at a contact discontinuity, by constructing the Kreiss symmetrizers microlocally away from the poles of the symbols, and studying the equations directly at each pole. The nonplanar case is studied by using the calculus of para-differential operators. The failure of the uniform Kreiss-Lopatinskii condition leads to a loss of derivatives of solutions in estimates. [ABSTRACT FROM AUTHOR]

The wind tunnel test section noise (pressure fluctuation) is one of the important indexes to evaluate the wind tunnel flow field, excessive noise not only significantly reducing the precision and accuracy of the collected data, but also arouses the model as well as wind tunnel parts buffeting response and causes serious fatigue damage to the materials, which is especially outstanding in transonic wind tunnels. This article discussed the causes of noise produced in the test section at first, with the common problems of general wind tunnels and the personalities of transonic wind tunnels included, and then based on wind tunnel tests studied sound laws of different test section wallboard type and model stand, analyzed the contribution for overall noise level from each component. [ABSTRACT FROM AUTHOR]

Abstract: A review of recent and historical work in the field of transonic and supersonic ground effect aerodynamics has been conducted, focussing on applied research on wings and aircraft, present and future ground transportation, projectiles, rocket sleds and other related bodies which travel in close ground proximity in the compressible regime. Methods for ground testing are described and evaluated, noting that wind tunnel testing is best performed with a symmetry model in the absence of a moving ground; sled or rail testing is ultimately preferable, though considerably more expensive. Findings are reported on shock-related ground influence on aerodynamic forces and moments in and accelerating through the transonic regime – where force reversals and the early onset of local supersonic flow is prevalent – as well as more predictable behaviours in fully supersonic to hypersonic ground effect flows. [Copyright &y& Elsevier]

This paper presents an overview of the work performed recently at ONERA on the control of the buffet phenomenon. This aerodynamic instability induces strong wall pressure fluctuations and as such limits aircraft envelope; consequently, it is interesting to try to delay its onset, in order to enlarge aircraft flight envelop, but also to provide more flexibility during the design phase. Several types of flow control have been investigated, either passive (mechanical vortex generators) or active (fluidic VGs, fluidic trailing-edge device (TED)). It is shown than mechanical and fluidic VGs are able to delay buffet onset in the angle-of-attack domain by suppressing the separation downstream of the shock. The effect of the fluidic TED is different, the separation is not suppressed, but the rear wing loading is increased and consequently the buffet onset is not delayed to higher angles of attack, but only to higher lift coefficient. Then, a closed loop control methodology based on a quasi-static approach is defined and several architectures are tested for various parameters such as the input signal, the objective function or, the tuning of the feedback gain. All closed loop methods are implemented on a dSPACE device calculating in real time the fluidic actuators command from the unsteady pressure sensors data. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to present a 2D study regarding the numerical simulation of flow within a transonic highly-loaded rotating cascade from an axial compressor. In order to describe an intricate flow pattern of a complex geometry and given specific conditions of cascade's loading and operation, an appropriate accurate flow model is a must. For such purpose, the Navier-Stokes equations system was used as flow model; from the computational point of view, the mathematical support is completed by a turbulence model. A numerical comparison has been performed for different turbulence models (e.g. KE, KO, Reynolds Stress and Spallart-Allmaras models). The convergence history was monitored in order to focus on the numerical accuracy. The force vector has been reported in order to express the aerodynamics of flow within the rotating cascade at the running regime, in terms of Lift and Drag. The numerical results, expressed by plots of the most relevant flow parameters, have been compared. It comes out that the selecting of complex flow models and appropriate turbulence models, in conjunction with CFD techniques, allows to obtain the best computational accuracy of the numerical results. This paper aims to carry on a 2D study and a prospective 3D will be intended for the same architecture. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to present a 2D study regarding the numerical simulation of flow within a transonic highly-loaded rotating cascade from an axial compressor. In order to describe an intricate flow pattern of a complex geometry and given specific conditions of cascade's loading and operation, an appropriate accurate flow model is a must. For such purpose, the Navier-Stokes equations system was used as flow model; from the computational point of view, the mathematical support is completed by a turbulence model. A numerical comparison has been performed for different turbulence models (e.g. KE, KO, Reynolds Stress and Spallart-Allmaras models). The convergence history was monitored in order to focus on the numerical accuracy. The force vector has been reported in order to express the aerodynamics of flow within the rotating cascade at the running regime, in terms of Lift and Drag. The numerical results, expressed by plots of the most relevant flow parameters, have been compared. It comes out that the selecting of complex flow models and appropriate turbulence models, in conjunction with CFD techniques, allows to obtain the best computational accuracy of the numerical results. This paper aims to carry on a 2D study and a prospective 3D will be intended for the same architecture. [ABSTRACT FROM AUTHOR]

Target influence on aerodynamic characteristics in video model deformation experiments has been ignored for high cost of wind tunnel tests, which causes an increase of uncertainty of aerodynamic force data. With the use of computational fluid dynamics, target influence on aerodynamic characteristics of the airfoil RAEE2822 was investigated through three cases to find the influence rules of different parameters. In case I, influence of target thickness ranging from 3H to 9H and target chordwise location ranging from 0.1 to 0.9 on both side surfaces of the airfoil was investigated. In case II, influence of angle of attack ranging from -4° to 8 ° was investigated. In case III, influence of four combination types of targets were investigated with the target thickness ranging from 1H to 9H. Computational results show that the target influence is sensitive to chordwise location and the relation between variation of aerodynamic force and target thickness is nonlinear. Aerodynamic force variation results of different angles of attack also indicate that influence of target on low surface is much smaller than that of target on upper surface when model has a positive angle of attack. [ABSTRACT FROM AUTHOR]

In our previous work, we have established the existence of transonic characteristic discontinuities separating supersonic flows from a static gas in two-dimensional steady compressible Euler flows under a perturbation with small total variation of the incoming supersonic flow over a solid right wedge. It is a free boundary problem in Eulerian coordinates and, across the free boundary (characteristic discontinuity), the Euler equations are of elliptic–hyperbolic composite-mixed type. In this paper, we further prove that such a transonic characteristic discontinuity solution is unique and L1–stable with respect to the small perturbation of the incoming supersonic flow in Lagrangian coordinates. [ABSTRACT FROM AUTHOR]

The linear stability based on Orr-Sommerfeld equations and the eN method was employed to determine the onset and control of transition over VFW-VF-2 and NACA653-018 aerofoil sections. For the VFW-VF-2 aerofoil, the effects of active suction on shock-boundary layer interaction were investigated. An enhanced lift to drag ratio was obtained for incidence angles below 6 degrees using suction coefficient as tiny as 0.0006. A parametric study were conducted to show the effects of the suction slot location, suction inclination angle and the extent of sucked flow through slots for flows around the NACA653-018 aerofoil. It was observed that transition was delayed regardless of suction positions. However, maximum delay in transition occurred when the sucked region was located between a critical point and the transition point. For all the cases studied here, lift to drag ratio was increased using active surface suction. [ABSTRACT FROM AUTHOR]

A transonic flow with a non-equilibrium condensation past NACA0012 profile whose aspect ratio AR is 1.0 with the angle of attack was analyzed by numerical analysis using a TVD scheme, and investigated using an intermittent indraft type supersonic wind tunnel. Transonic flows of 0.70-0.90 in free stream Mach number with variations of Φ and α were tested. For the same M and α, the increase in stagnation relative humidity Φ caused a decrease in the drag coefficient of profile (that is total) which is composed of the components of form, viscous, wave and condensation; however, the lift coefficient up to Φ = 50% increased in the opposite direction. As an example, in the case of M = 0.83, Φ = 50%, α = 3○ and T = 298 K, the decreasing rate of the coefficient of profile drag and the increasing rate of the lift coefficient compared to the case of Φ = 0% caused by non-equilibrium condensation amounted to 65% and 52%, respectively. In addition, for the same Φ and α, as the free stream Mach number M increased, at first, the lift coefficient increased slightly, and then suddenly severely dropped, and finally remained nearly constant. The suddenly dropped free stream Mach number in C became larger with an increase of Φ. It turned out that the drag coefficients of form and viscous were almost independent of Φ. The contribution of wave drag to the coefficient of profile drag for M = 0.83, Φ = 30%, α = 3° and T = 298 K amounted to approximately 79%, and in the case of Φ = 60%, the contribution of the non-equilibrium condensation to the reduction in the coefficient of the profile drag compared to the case without condensation amounted to 75%. Especially, for the case of Φ = 0% and α = 0°, there was an oscillatory flow region around M = 0.87. [ABSTRACT FROM AUTHOR]

Abstract: High-speed flow research has been sponsored and performed at differing levels of effort since the late 1800s. For example, hypersonic research has experienced numerous cycles since the 1950s, with various periods of high research activity, followed by equally long periods of very low activity. This lack of continuity in high-speed flow research has led to a situation where researchers of one “generation” often do not know what the researchers of previous “generations” have done, mainly due to large losses of institutional knowledge in government, industry, and academic organizations. Therefore, a chronically weak area in research papers, reports, and reviews is the complete identification of critical background documents that form the building blocks and intellectual heritage for modern compressible flow research. A method for systematically determining these critical references is presented in the context of its application to high-speed flow using Citation-Assisted Background, which is based on the assumption that many critical documents tend to be highly cited within the literature, although not necessarily recently. While Citation-Assisted Background is a highly systematic approach for identifying critical references, it is not a substitute for the judgement of the researchers, but rather complements their expertise. In this critical review of high-speed compressible flow, important documents have been identified using Citation-Assisted Background, but other documents have been added by the authors to enhance the picture provided by the highly-cited documents. [Copyright &y& Elsevier]

Abstract: Little has been published on the details of design and fabrication of differential mobility analyzers (DMAs) classifying 1nm particles with resolving powers of 30 or more. These DMAs must operate at Reynolds numbers Re⪢2000 (supercritical), requiring laminarizers and diffusers that have tended to make them large and heavy. Here we discuss design and fabrication criteria as well as observed performance for several variants of a miniature supercritical DMA weighing 2.7kg. In spite of the narrow working sections associated to small electrode radii of R 1=4mm and R 2=6 or 7mm (gap=R 2−R 1=Δ=2 or 3mm), a large enough Re (hence resolution Я down to 1nm diameter particles) is achieved via a diffuser enabling up to transonic sheath gas speeds. Several axial distances L between the inlet and outlet aerosol slits have been tested to facilitate classification of particles with diameters above 20nm. Я is limited by concentricity errors, magnified by the small gaps used. Nonetheless, short models with L/Δ=2 achieve consistently Я>40. The long 2 and 3nm gap model reaches typically Я>25–30, and exceptionally up to Я=40. This performance is with 1nm particles, and would probably be better with larger particles. While prior long supercritical DMAs have used slightly conical inner electrodes to accelerate and stabilize the flow, a cylindrical version of our long DMAs shows no signs of turbulent transition. [Copyright &y& Elsevier]

Firstly, this paper presents a new expression which changes the isentropic flow equation into Cauchy-Riemann equation, and the exact solution of subsonic flow around an arbitrary two-dimensional body is obtained. Secondly, another new affine transformation is employed, with a coordinate transformation, to change the isentropic flow equation into the hyperbolic equation through the transformations. Therefore, an exact solution of transonic flow is obtained. And the two solutions make the variables of velocity, pressure and density contain continuity near the sonic line. [ABSTRACT FROM AUTHOR]

The application of proper orthogonal decomposition (POD), which is an efficient method of data analysis for two dimensional airfoil transonic flow field reconstruction, has been proved successfully in this paper. The reduced order model (ROM) for flows has been constructed by using the snap shots from the knowledge of computational data or experimental data of flows. Based on this ROM, it is possible to reconstruct the entire flow field from the gappy data accurately in the range of snap shots. Using this approach, another application which is the two dimensional inverse airfoil design based on POD/ROM has been demonstrated successfully. This methodology demonstrates a great simplification for the inverse airfoil design problem. In this paper, the cases of transonic flow field reconstruction of RAE 2822 airfoil and the inverse design of NACA 63212 and Korn airfoil have been simulated. These cases prove that the method of flow field reconstruction and inverse airfoil design based on POD is efficient and utilizable in the range of transonic. [ABSTRACT FROM AUTHOR]

Abstract: We present the existence of the subsonic solution to a two-dimensional Riemann problem governed by a self-similar nonlinear wave equation where the boundary of the subsonic region consists of a transonic shock and the sonic circle. Thus the governing equation becomes a free boundary problem on the transonic shock and degenerates on the sonic circle. By utilizing the barrier methods and iterative methods, we show the well-posedness of the transonic shock in the entire subsonic region and thus establish the global solution. This result does not rely on any smallness of Riemann data. [Copyright &y& Elsevier]

Abstract: To analyze the procedure of gasdynamically generating nanoparticles, the turbulent mixing of the heavy gas tetraethoxysilane (TEOS - ) in a hot accelerated transonic air co-flow is numerically investigated. The precursor gas TEOS is premixed with nitrogen and then ejected through circular holes in the base of an injector into the subsonic region of a nozzle flow at a co-flow Mach number of and a temperature of . A multi-species large-eddy simulation (LES) capturing the most energetic turbulent scales is used to simulate the mixing process. Close behind the blunt trailing edge of the injector the results show the mixing to be dominated by pronounced free shear layers. Downstream of the injector the flow field is accelerated to supersonic speed and dominated by trailing vortices, which transport the partially mixed precursor gas but do not enhance further mixing. The analysis of the simulation results indicates that no further mixing occurs approximately 16 trailing edge thicknesses downstream of the trailing edge. [Copyright &y& Elsevier]

Abstract: Based on aerodynamic identification technology in which unsteady CFD (computational fluid dynamics) is used, the unsteady aerodynamic reduced order models (ROM) are constructed. Coupled with structural equations, we get the analyzable models for transonic aeroelasticity in state-space. A Mach number flutter trend of a typical airfoil section with a control surface is analyzed and agrees well with that of CFD/CSD (computational structural dynamics) direct coupling method. Then we study the effect of the structural parameters (natural frequency and the flap static unbalance) of the control surface on the transonic flutter system. We find some classical structural design rules are unfavorable for transonic flutter problem. For transonic flutter problem, the flap rotating mode becomes the predominant mode of aeroelastic system. Classical technology of mass static balance or over-balance maybe reduces the stability of transonic aeroelasticity. [Copyright &y& Elsevier]

By means of a high-speed gas flow generated by a miniature supersonic nozzle, we proposed a unique biolistic method to accelerate microparticle formulation of drugs to sufficient momentum, to penetrate the outer layer of human skin or mucosal tissue for the treatment of a range of diseases. One of the main concerns for designing and evaluating this system is ensuring microparticles delivery into human skin with a controllable velocity range and spatial distribution. The initial experimental work suggested that the performance of the transdermal delivery strongly depends on aerodynamics of the supersonic nozzles employed. In this paper, computational fluid dynamics (CFD) is utilized to characterize existing prototype biolistic delivery systems, the device with a converging-diverging supersonic nozzle (CDSN) and the device based on the contoured-shock-tube (CST) design, with the aim at investigating the transient gas and particle dynamics in the supersonic nozzles. Whenever possible, predicted pressure and Macb number histories, 2-D flow structures, and particle velocity distributions are made to compare with the corresponding experimental measurements to validate the implemented numerical approach. The gas-particle interaction and performance of two biolistic devices are interrogated and distinguished. Subsequently, the particle impact conditions are presented and discussed. It is demonstrated that the CST can deliver microparticles with a narrow and more controllable velocity range and spatial distribution. [ABSTRACT FROM AUTHOR]

Abstract: This paper presents an efficient implicit solver for transonic shock-tube flow, generated for a needle-free epidermal delivery of powdered vaccines. This transient transonic flow exhibits complicated flow phenomena. We proposed a unique method of implicit formation on the basis of analysing a model equation, which can separately deal with convection, diffusion and source terms. By combining this idea with flux-vector-splitting, an implicit flux-vector-splitting solver of the Navier–Stokes equations, which can avoid approximate-factorization (AF) or block-bidiagonalization, is developed. Numerical experiments show that it has obvious superiority over the conventional flux-vector-splitting scheme in terms of convergence and computing cost, meanwhile maintaining a high accuracy and robustness. [Copyright &y& Elsevier]

Abstract: The application of an iterative design method allows the aerodynamic design of engine nacelles and wings being part of complex aircraft configurations. The method couples a flow solver for the solution of the Euler/Navier–Stokes equations on unstructured grids with a design algorithm which solves the transonic potential equation for small perturbations via an inverse formulation. The accuracy and robustness of the design method is investigated for redesigns of engine nacelles and wings. Several design cases demonstrate the efficient application of the method for the aerodynamic design of installed engine nacelles and wings under influence of complex aircraft configurations. One application shows that the method also can be used to minimize aerodynamic installation effects resulting from the engine integration process by sole nacelle or wing contour shaping. A general design methodology for engine nacelles is derived which allows to take into account all aerodynamic effects on complex configurations during the design process. [Copyright &y& Elsevier]

This article reports the measurements of time-resolved heat transfer rate and time-resolved static pressure that have been made on the over-tip casing of a transonic axial-flow turbine operating at flow conditions that are representative of those found in modern gas turbine engines. This data is discussed and analysed in the context of explaining the physical mechanisms that influence the casing heat flux. The physical size of the measurement domain was one nozzle guide vane-pitch and from -20% to +80% rotor axial chord. Additionally, measurements of the time-resolved adiabatic wall temperature are presented. The time-mean data from the same set of experiments is presented and discussed in Part I of this article. The nozzle guide vane exit flow conditions in these experiments were a Mach number of 0.93 and a Reynolds number of 2.7 × 106 based on nozzle guide vane mid-height axial chord. The data reveal large temporal variations in heat transfer characteristics to the casing wall that are associated with blade-tip passing events and in particular the blade over-tip leakage flow. The highest instantaneous heat flux to the casing wall occurs within the blade-tip gap, and this has been found to be caused by a combination of increasing flow temperature and heat transfer coefficient. The time-resolved static pressure measurements have enabled a detailed understanding of the tip-leakage aerodynamics to be established, and the physical mechanisms influencing the casing heat load have been determined. In particular, this has focused on the role of the unsteady blade lift distribution that is produced by upstream vane effects. This has been seen to modulate the tip-leakage flow and cause subsequent variations in casing heat flux. The novel experimental techniques employed in these experiments have allowed the measurement of the time-resolved adiabatic wall temperature on the casing wall. These data clearly show the falling flow temperatures as work is extracted from the gas by the turbine. Additionally, these temperature measurements have revealed that the absolute stagnation temperature within the tip-gap flow can be above the turbine inlet total temperature, and indicates the presence of a work process that leads to high adiabatic wall temperatures as a blade-tip passes a point on the casing wall. It is shown that this phenomena can be explained by consideration of the flow vectors within the tip-gap, and that these in turn are related to the local blade loading distribution. The paper also assesses the relative importance of different time-varying phenomena to the casing heat load distribution. This analysis has indicated that up to half of the casing heat load is associated with the over-tip leakage flow. Finally, the implications of the experimental findings are discussed in relation to future shroudless turbine design, and in particular the importance of accounting for the high heat fluxes found within the tip-gap. [Copyright &y& Elsevier]

The over-tip casing of the high-pressure turbine in a modern gas turbine engine is subjected to strong convective heat transfer that can lead to thermally induced failure (burnout) of this component. However, the complicated flow physics in this region is dominated by the close proximity of the moving turbine blades, which gives rise to significant temporal variations at the blade-passing frequency. The understanding of the physical processes that control the casing metal temperature is still limited and this fact has significant implications for the turbine design strategy. A series of experiments has been performed that seeks to address some of these important issues. This article reports the measurements of time-mean heat transfer and time-mean static pressure that have been made on the over-tip casing of a transonic axial-flow turbine operating at flow conditions that are representative of those found in modern gas turbine engines. Time-resolved measurements of these flow variables (that reveal the details of the blade-tip/casing interaction physics) are presented in a companion paper. The nozzle guide vane exit flow conditions in these experiments were a Mach number of 0.93 and a Reynolds number of 2.7 × 106 based on nozzle guide vane mid-height axial chord. The axial and circumferential distributions of heat transfer rate, adiabatic wall temperature, Nusselt number and static pressure are presented. The data reveal large axial variations in the wall heat flux and adiabatic wall temperature that are shown to be primarily associated with the reduction in flow stagnation temperature through the blade row. The heat flux falls by a factor of 6 (from 120 to 20 kW/m2). In contrast, the Nusselt number falls by just 36% between the rotor inlet plane and 80% rotor axial chord; additionally, this drop is near to linear from 20% to 80% rotor axial chord. The circumferential variations in heat transfer rate are small, implying that the nozzle guide vanes do not produce a strong variation in casing boundary layer properties in the region measured. The casing static pressure measurements follow trends that can be expected from the blade loading distribution, with maximum values immediately upstream of the rotor inlet plane, and then a decreasing trend with axial position as the flow is turned and accelerated in the relative frame of reference. The time-mean static pressure measurements on the casing wall also reveal distinct circumferential variations that are small in comparison to the large pressure gradient in the axial direction. [Copyright &y& Elsevier]