Low Subsonic Flow Research Articles

Influence of the Blunt Trailing-Edge Thickness on the Aerodynamic Characteristics of the Very Thick Airfoil

In this paper, the NWT600 airfoil with a thickness ratio of 60% is taken as the research object. The aerodynamic performance of the airfoil is analyzed by experiments and numerical simulations. The results simulated by various turbulence models used in the 2D steady-state RANS method are compared, including the Spalart–Allmaras model, k-ω SST model, k-ε realizable model, and Reynolds stress (linear pressure-strain) model. The influence of blunt trailing-edge thickness on aerodynamic characteristics is studied by adding thickness symmetrically. The results show that even under the low subsonic flow with a Mach number of 0.149, the airflow is prone to severe separation. The aerodynamic performance of the airfoil is very different from that of the conventional thin airfoil. Although the 2D steady-state RANS models overestimate the pressure on the surface of the airfoil in most cases, it is qualitatively acceptable to predict the pressure distribution of the very thick airfoil. Numerical results simulated by the Reynolds stress model are in the best agreement with the experimental data. It is also found that symmetrically thickening the trailing edge effectively improves the maximum lift coefficient and reduces the drag coefficient at a small angle of attack.

Fineness Ratio Effects on the Flow Past an Axisymmetric Body at High Incidence

The flow past an axisymmetric body at a sufficiently high angle of attack becomes asymmetric and unsteady. Several authors identified three different flow regions for bodies of large fineness ratio at low subsonic flow and high incidence: a steady region in the forebody and two unsteady regions in the rear body. Unsteady Reynolds Averaged Navier–Stokes (URANS) codes with eddy viscosity turbulence models or Reynolds stress turbulence models fail to capture the unsteady flow region. These methods are overly dissipative and resolve only frequencies far lower than turbulent fluctuations. Scale-Adaptive-Simulation (SAS) provides an alternative method to afford the problem of these massively separated flows at high Reynolds numbers without addressing the problem to Large Eddy Simulation (LES). This paper applies SAS to study the effect of slenderness on the flow. The numerical solutions show that the flow becomes more unstable as the fineness ratio increases, and the three flow regions are clearly recognizable. For low fineness ratios, only one of the two unsteady regions is visible. The good agreement between the sectional forces and pressure coefficients with their corresponding experimental data for an ogive-cylinder configuration allows an analysis of the flow structure with a fair degree of confidence.

Исследование развития локализованных возмущений в пограничном слое плоской пластины в условиях повышенной степени турбулентности набегающего потока за уступом поверхности

In a wind tunnel on a flat plate in a separated flow behind a rectangular step, the emergence and development of localized disturbances generated by low-frequency impulse deviations of the local surface section under conditions of low and moderate degrees of the incoming flow turbulence is studied. The results were obtained by hot-wire anemometry at low subsonic flow velocity. It was found that impulse deviations of the wall generate disturbances, which are socalled Streaky structures and wave packets of oscillations. The separation of the laminar boundary layer accelerates the growth of wave packets with subsequent turbulization of the near-wall flow. The specific features of the behavior of localized disturbances under conditions of moderate degree of free-stream turbulence are revealed.

Flow induced oscillation of two rigid rectangular plates in a side-by-side configuration

The present work is an experimental study of two oscillating rigid plates placed in side-by-side configuration, hinged at their leading edges, subjected to low subsonic flow. This problem is investigated using smoke-wire flow visualization, hot-wire anemometry, and time resolved particle image velocimetry. It is found that beyond a critical Reynolds number, the plates set into oscillatory motion. This critical Reynolds number depends on the gap between the plates. It is also seen that this value of Reynolds number, at lower values of gap to thickness ratio (<7) is significantly higher than that of the single plate configuration value. The frequency and amplitude of the oscillating plates at various gaps and Reynolds numbers have been studied and compared with the characteristics of an oscillating single plate. It is also found that depending on the gap and acceleration of the free-stream, there exist two modes of oscillation - (i) in-phase and (ii) out-of-phase. For gap to thickness ratio less than 10, only in-phase oscillations take place for all values of free-stream velocity considered in the present work, whereas, when this ratio is greater than 10, the mode of oscillation depends on the initial conditions up to a certain free-stream velocity, beyond which the plates switch to in-phase mode. Smoke wire flow visualization technique along with time resolved particle image velocimetry reveal that the vorticity distributions around the plates are responsible for the initiation of the two modes of oscillations.

An implicit implementation of the characteristic boundary condition in a fully coupled pressure-based flow solver

The article deals with an implicit formulation of the pressure far field boundary condition, also known as the characteristic boundary condition, in a pressure-based coupled solver. This boundary condition applies to compressible flows over the entire Mach regime, and is derived by invoking the Riemann invariants to implicitly express the flow variables at inlets and outlets in terms of their values inside the domain. A set of inviscid an turbulent flow cases that include pressure far field boundary conditions are tested, namely: low subsonic compressible flow past a NACA0012 airfoil at 10° angle of attack; transonic flow over circular bump; supersonic flow over a series of slender bumps; and DLR-F6 Wing-Body-Nacelle-Pylon Aircraft. Predictions using the prescribed coupled solver are in good agreement with similar results obtained with density-based methods and/or experimental data.

Interaction of an Optical Discharge with a Shock Wave

The results of an experimental study of the impact of the focused pulsed-periodic radiation from a CO2 laser on a gas-dynamic structure in a supersonic jet are presented. The radiation of the CO2 laser is propagated across the stream and focused by a lens on the axis of the supersonic jet. To register the flow structure, a shadow scheme with a slit and a flat knife located along the flow is used. The image is fixed by a speed camera with an exposure time of 1.5 μs and a frame rate of 1000 1/s. In the flow, the plasma initiated by the pulsedperiodic laser is visualized in order to identify and determine the period of plasma development, as well as the motion of the initial front of the shock wave. It is shown that at the transverse input of laser radiation into the stream the periodic structure of the thermal trace is created with the formation of an unsteady shock wave from the energy release zone. At small repetition rates of laser radiation pulses, the thermal spot interacts with the flow in the pulsed mode. It is shown that elliptic nonstationary shock waves are formed only at low subsonic flow velocities and in a stationary atmosphere. The process of nonstationary ignition by an optical discharge of a methane–air mixture during a subsonic outflow into a motionless atmosphere is shown experimentally. The results of optical visualization indicate burning in the trace behind the optical discharge region.

Non-intrusive acoustic measurement of flow velocity and temperature in a high subsonic Mach number jet

In previous studies, sonic anemometry and thermometry have generally been used to measure low subsonic Mach flow conditions. Recently, a novel configuration was proposed and used to measure unheated jet velocities up to Mach 0.83 non-intrusively. The objective of this investigation is to test the novel configuration in higher temperature conditions and explore the effects of fluid temperature on mean velocity and temperature measurement accuracy. The current work presents non-intrusive acoustic measurements of single-stream jet conditions up to Mach 0.7 and total temperatures from 299 K to 700 K. Comparison of acoustically measured velocity and static temperature with probe data indicate root mean square (RMS) velocity errors of 2.6 m s−1 (1.1% of the maximum jet centerline velocity), 4.0 m s−1 (1.2%), and 8.5 m s−1 (2.4%), respectively, for 299, 589, and 700 K total temperature flows up to Mach 0.7. RMS static temperature errors of 7.5 K (2.5% of total temperature), 8.1 K (1.3%), and 23.3 K (3.3%) were observed for the same respective total temperature conditions. To the authors’ knowledge, this is the first time a non-intrusive acoustic technique has been used to simultaneously measure mean fluid velocity and static temperatures in high subsonic Mach numbers up to 0.7. Overall, the findings of this work support the use of acoustics for non-intrusive flow monitoring. The ability to measure mean flow conditions at high subsonic Mach numbers and temperatures makes this technique a viable candidate for gas turbine applications, in particular.

A correlation-based algebraic transition model

A correlation-based algebraic transition model that relies on local flow information is proposed. The model is qualified as an algebraic model, or a zero-equation model since it includes an intermittency function in place of an intermittency equation that is found in one- or two-equation models. The basic idea behind the model is that, instead of deriving new equations for intermittency transport, existing transport terms of the Spalart–Allmaras (S-A) turbulence model can be used. To this end, the production term of the S-A model is multiplied with the proposed intermittency function γBC; thereby the turbulence production is damped until it satisfies some turbulence onset requirements. The proposed formulation also depends on local information that uses empirical correlations to detect the transition onset using less equations and less calibration constants than other higher order models. The model is first validated against some widely-used zero and variable pressure gradient flat plate test cases with quite successful results. Second, the model is employed for some low Reynolds number airfoil cases with very promising results. Third, the model is applied for a turbine cascade case with success. Finally, two different three-dimensional wing flow cases were calculated under transonic and low subsonic flow conditions. To this end, the DLR-F5 wing subject to a transonic Mach number of 0.82 and the low-speed NREL wind turbine flow case are simulated and good agreement with experiments are observed. The results indicate that the proposed model may become an alternative for other models as it uses less computational resources with equivalent or higher accuracy characteristics that is quite advantageous for the computational fluid dynamics design in industry.

An Efficient Semi-implicit Solver for Direct Numerical Simulation of Compressible Flows at All Speeds

We develop a semi-implicit algorithm for time-accurate simulation of compressible shock-free flows, with special reference to wall-bounded flows. The method is based on partial linearization of the convective fluxes in such a way to suppress, or at least mitigate the acoustic time step limitation. Together with replacement of the total energy equation with the entropy transport equation, this approach avoids the inversion of block-banded matrices involved in classical methods, which is replaced by much less demanding inversion of standard banded matrices. The method is extended to deal with implicit integration of viscous terms and to multiple space dimensions through approximate factorization, and used as a building block of a semi-implicit Runge–Kutta scheme which guarantees third-order of accuracy in time (Nikitin in Int J Numer Methods Fluids 51:221–233, 2006). Numerical experiments are carried out for isotropic turbulence, plane channel flow, and flow in a square duct. All available data support higher computational efficiency than existing methods, and saving of computer resources ranging from 85% under low-subsonic flow conditions (down to $$M_0 \sim 0.1$$ ), to about 50% in supersonic flow.

Fluidic thrust vectoring using transverse jet injection in a converging nozzle with aft-deck

This paper summarizes the results of an experimental study carried out to implement fluidic thrust vectoring in yaw direction using transverse jet injection in a converging nozzle with an elliptic exit and triangular shaped aft-deck. The study was restricted to low subsonic flow regime typical for Unmanned Aerial Vehicle (UAV) applications. The momentum of the core nozzle flow was kept constant while varying the momentum of the secondary-jet. The surface pressure distribution, exhaust jet velocity field and thrust were measured. Secondary-jet injection was found to create a virtual exit plane at an angle to the stream wise direction of the nozzle flow resulting in the turning of the core-jet away from the injection slot and turning the direction of thrust (termed vectored thrust). The angle of the vectoring was found to increase with increase in momentum of the secondary-jet but only up to a certain maximum. The vectoring was found to increase the static pressure at the nozzle inlet, which is a very crucial observation in this study as it is directly related to engine performance. The secondary-jet injection was found to reduce the width of the core-jet and the magnitude of this reduction was found to increase in stream wise direction and also increase with increase in the momentum of the secondary-jet. Core flow behavior was different in the region with aft-deck compared to region without aft-deck. Jet spreading was expedited by the secondary-jet and aft-deck. Multiple shear layers were present in the flow with the secondary-jet switched on, which were otherwise absent. Secondary-jet injection was found to result in the formation of vortices with axes oriented along the direction of the secondary-jet at the injection slot and these vortices convect downstream with the core flow. These vortices were found to excite the shear layer, resulting in better jet mixing and hence augment the mass entrainment by the core flow.

On the computation of low-subsonic turbulent pipe flow noise with a hybrid LES/LPCE method

Aeroacoustic computation of a fully-developed turbulent pipe flow at Reτ = 175 and M = 0.1 is conducted by LES/LPCE hybrid method. The generation and propagation of acoustic waves are computed by solving the linearized perturbed compressible equations (LPCE), with acoustic source DP(x,t)/Dt attained by the incompressible large eddy simulation (LES). The computed acoustic power spectral density is closely compared with the wall shear-stress dipole source of a turbulent channel flow at Re τ = 175. A constant decaying rate of the acoustic power spectrum, f -8/5 is found to be related to the turbulent bursts of the correlated longitudinal structures such as hairpin vortex and their merged structures (or hairpin packets). The power spectra of the streamwise velocity fluctuations across the turbulent boundary layer indicate that the most intensive noise at ω + z R < 1.5).

CFD Analysis on Different Shapes of Winglet at Low Subsonic Flow

This paper focuses on aerodynamic characteristic ofaircraft wing with and without winglet at low subsonic flow.NACA 653-218 profile was applied on four models including elliptical, semicircular and straight winglet. CFD analysis was performed to compare the lift to drag ratio in these models. Spallart Allmarus turbulence model and 3D unstructured tetrahedral mesh was used to compute the flow around the winglet prototype. It was found that elliptical winglet gave maximum lift to drag ratio.

An approach for the numerical prediction of the compressible supercavitating flows over ultrahigh-speed underwater object

Purpose Ultrahigh-speed projectile running in water with the velocity close to the speed of sound usually causes large supercavity. The computation of such transonic cavitating flows is usually difficult, thus high-speed model reflecting the compressibility of both the liquid and the vapor phases should be introduced to model such flow. The purpose of this paper is to achieve a model within an in-house developed solver to simulate the ultrahigh-speed subsonic supercavitating flows. Design/methodology/approach An improved TAIT equation adjusted by local temperature is adopted as the equation of state (EOS) for the liquid phase, and the Peng-Robinson EOS is used for the vapor phase. An all-speed variable coupling algorithm is used to unify the computations and regulate the convergence at arbitrary Mach number. The ultrahigh-speed (Ma=0.7) supercavitating flows around circular disk are investigated in contrast with the case of low subsonic (Ma=0.007) flow. Findings The characteristic physical variables are reasonably predicted, and the cavity profiles are compared to be close to the experimental empirical formula. An important conclusion in the compressible cavitating flow theory is verified by the numerical result that, at any specific cavitation number the cavity’s size and the drag coefficient both increase along with the rise of Mach number. On the contrary, it is found as well that the cavity’s slenderness ratio decreases when Mach number goes up. It indicates that the compressibility has different influences on the length and the radius of the supercavity. Originality/value A high-speed model reflecting the compressibility of both the liquid and the vapor phases was suggested to model the ultrahigh-speed supercavitating flows around underwater projectiles.

Effects of short splitter plates on vortex shedding and sound generation in flow past two side-by-side square cylinders

Effects of short splitter plates on vortex shedding and sound generation in a low subsonic flow past two side-by-side square cylinders were examined experimentally at Reynolds numbers $${{{Re}}} = 1.0 - 3.3 \times 10^4$$ . The experiment was mainly conducted with the center-to-center distance between the two cylinders of 3.6d (d is the side length of a square cylinder) where vortex shedding from the two cylinders was synchronized with anti-phase relation, generating a quadrupole-like sound source that radiated in-phase sound in the far field. The results showed that the attachment of short splitter plates whose length (c) was ≤0.5d could reduce the sound pressure level of Aeolian tone significantly. Even with the shortest splitter plates of $$c{/}d=0.1$$ , SPL was reduced by 6 dB at Mach number $$M_{\infty }=0.15$$ . This was in contrast to the case of a single square cylinder, for which the attachment of a short splitter plate <0.2d led to no noticeable noise reduction. It was also shown that even when short splitter plates with a spanwise length as long as or less than the correlation length of shed vortices were attached on the two cylinders in a staggered array, the anti-phase vortex shedding and the corresponding in-phase sound radiation were still dominant.

Experimental investigation of linear and nonlinear aeroelastic behavior of a cropped delta wing with store in low subsonic flow

The present paper aims to investigate flutter phenomenon for cropped delta wing with an external store using numerical and experimental methods in a subsonic and incompressible flight regime. Wing structure is modeled based on von Karman plate theory. The unsteady vortex lattice method is applied to wing aerodynamic model and a slender-body theory is used for store aerodynamic model. The experimental tests have been conducted in an incompressible subsonic wind tunnel. The comparison indicates a satisfactory agreement between the experimental results and the theoretical analyses. Moreover, the effects of different parameters such as wing thickness, wing aspect ratio, spanwise store position, store mass, the aerodynamics of the store, underwing clearance and store center of gravity on both flutter speed and instability boundary of the wing are studied both analytically and experimentally. Also, the system behaviors and the oscillation amplitude are examined experimentally for post-flutter where time history simulation and phase portrait are obtained for various velocities. The FFT analysis of time history measured data shows that beside the dominant harmonic frequency, both superharmonic and subharmonic frequencies are clearly identifiable. The results obtained from experimental bifurcation diagram indicate the presence of hysteresis loop regions corresponding to the stable LCOs.

Dynamic characteristics of MarcoPolo-R Entry Capsule in low subsonic flow

The selection of the most suitable configuration between several candidate aeroshapes for the MarcoPolo-R ERC is performed by means of subsonic dynamic stability tests. These tests were conducted in the vertical free jet facility VMK at the German Aerospace Center (DLR) in Cologne, employing the free flight technique. Several different model configurations with respect to the shape, position of the centre of gravity, shoulder radius, and mass were tested. The motion of the models during free flight was recorded using an optical motion tracking system. The measurement of the position and the attitude of the models, in combination with flow parameters and an exact flow profile characterization, allows for a shape discrimination based on aerodynamic stability characteristics of the models.

Intermittency of the near pressure field induced by a compressible coaxial jet

A novel method of analysis based on the combined use of the wavelet transform and linear stochastic estimation (LSE) is applied to pressure and velocity signals obtained in a coaxial jet at low subsonic flow conditions. The aim of the analysis is to identify and clarify the role of intermittency in the generation of noise. The wavelet analysis extracts intermittent events from the near field pressure signals and the LSE technique identifies velocity fluctuations which are linearly associated with the selected pressure fluctuations. A frequency–wavenumber filtering procedure is also applied to the near field pressure in order to separate the hydrodynamic and the acoustic components. The results achieved helps to identify and localize fluid dynamic events responsible for the generation of near field hydrodynamic and acoustic intermittent pressure fluctuations.

Active flutter suppression of a multiple-actuated-wing wind tunnel model

In this study, a multi-input/multi-output (MIMO) time-delay feedback controller is designed to actively suppress the flutter instability of a multiple-actuated-wing (MAW) wind tunnel model in the low subsonic flow regime. The unsteady aerodynamic forces of the MAW model are computed based on the doublet-lattice method (DLM). As the first attempt, the conventional linear quadratic-Gaussian (LQG) controller is designed to actively suppress the flutter of the MAW model. However, because of the time delay in the control loop, the wind tunnel tests illustrate that the LQG-controlled MAW model has no guaranteed stability margins. To compensate the time delay, hence, a time-delay filter, approximated via the first-order Pade approximation, is added to the LQG controller. Based on the time-delay feedback controller, a new digital control system is constructed by using a fixed-point and embedded digital signal processor (DSP) of high performance. Then, a number of wind tunnel tests are implemented based on the digital control system. The experimental results show that the present time-delay feedback controller can expand the flutter boundary of the MAW model and suppress the flutter instability of the open-loop aeroelastic system effectively.

Laminar Boundary Layer Separation Affected by a Local Temperature Perturbation

Backward-facing step flow with laminar boundary layer separation forced by a local stationary temperature disturbance was experimentally investigated. The results of the study were obtained in quiet wind-tunnel conditions at low subsonic flow velocity. To generate the temperature perturbation, the test model was equipped with a Peltier element placed flush with the wall before the line of separation. At first, the hot-wire technique was adapted to examine the flow with temperature gradients; then, the velocity characteristics were determined making clear the effects of heat transfer on boundary layer separation

Energy harvesting from limit cycle oscillation of a cantilever plate in low subsonic flow by ionic polymer metal composite

In this work, energy harvesting from limit cycle oscillation of low aspect ratio rectangular cantilever plates in low subsonic flow by ionic polymer metal composite is considered. The plate is modeled in according to classic plate theory with Von-Karman strain–displacement relations for modeling large deflections due to midplane stretching. Aerodynamics flow is modeled based on modified vortex lattice aerodynamics model. Linear and nonlinear aeroelastic characteristics of the considered model are accurately examined and the effects of ionic polymer metal composite energy harvesting properties on flutter margin and limit cycle oscillation amplitudes are investigated. The obtained results show that output power of [Formula: see text] is generated when the flow velocity reaches to [Formula: see text] , and position of ionic polymer metal composite on the plate has great effect on the amount of power produced. The extraction of energy from plate vibration in conjunction with large strain characteristics of ionic polymer metal composite materials produces static deflection of plate. This static deflection produces stiffness hardening of the entire system, accordingly reduces the amplitude of limit cycle oscillation. Obtained results show that ionic polymer metal composite actuator has more influence on limit cycle oscillation of plate, while its effect on flutter instability is negligible.