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1.
The mechanism of the origin of shock oscillations on NACA0012 aerofoils is investigated using a moving grid thin layer Navier
Stokes code. The method used to understand the mechanism is to initiate the shock oscillations on an aerofoil by moving the
aerofoil from a regime of steady transonic flow into a regime of periodic flow by a change in airflow incidence. The results
indicate that the shock induced bubble plays a leading role in the origin of shock oscillations and the trailing edge has
an affect on its amplitude.
Received 1 April 1997 / Accepted 1 December 1997 相似文献
2.
Time-resolved stereo PIV measurements of shock–boundary layer interaction on a supercritical airfoil
Time-resolved stereo particle-image velocimetry (TR-SPIV) and unsteady pressure measurements are used to analyze the unsteady
flow over a supercritical DRA-2303 airfoil in transonic flow. The dynamic shock wave–boundary layer interaction is one of
the most essential features of this unsteady flow causing a distinct oscillation of the flow field. Results from wind-tunnel
experiments with a variation of the freestream Mach number at Reynolds numbers ranging from 2.55 to 2.79 × 106 are analyzed regarding the origin and nature of the unsteady shock–boundary layer interaction. Therefore, the TR-SPIV results
are analyzed for three buffet flows. One flow exhibits a sinusoidal streamwise oscillation of the shock wave only due to an
acoustic feedback loop formed by the shock wave and the trailing-edge noise. The other two buffet flows have been intentionally
influenced by an artificial acoustic source installed downstream of the test section to investigate the behavior of the interaction
to upstream-propagating disturbances generated by a defined source of noise. The results show that such upstream-propagating
disturbances could be identified to be responsible for the upstream displacement of the shock wave and that the feedback loop
is formed by a pulsating separation of the boundary layer dependent on the shock position and the sound pressure level at
the shock position. Thereby, the pulsation of the separation could be determined to be a reaction to the shock motion and
not vice versa. 相似文献
3.
The results of an experimental study of the transonic flow behind the trailing corner edges of oversized cone-cylinder bodies of different dimensions are presented. Emphasis is placed on the induced transonic flow restructuring and the local aerodynamic forces accompanying the process, both steady and unsteady. The restructuring that occurs with increase in the freestream Mach number is studied. 相似文献
4.
The paper presents the development and application of a three-sensor wedge probe to measure unsteady aerodynamics in a transonic
turbine. CFD has been used to perform a detailed uncertainty analysis related to probe-induced perturbations, in particular
the separation zones appearing on the wedge apex. The effects of the Reynolds and Mach numbers are studied using both experimental
data together with CFD simulations. The angular range of the probe and linearity of the calibration maps are enhanced with
a novel zonal calibration technique, used for the first time in compressible flows. The data reduction methodology is explained
and demonstrated with measurements performed in a single-stage high-pressure turbine mounted in the compression tube facility
of the von Karman Institute. The turbine was operated at subsonic and transonic pressure ratios (2.4 and 5.1) for a Reynolds
number of 106, representative of modern engine conditions. Complete maps of the unsteady flow angle and rotor outlet Mach number are documented.
These data allow the study of secondary flows and rotor trailing edge shocks. 相似文献
5.
6.
G. Jourdan L. Biamino C. Mariani C. Blanchot E. Daniel J. Massoni L. Houas R. Tosello D. Praguine 《Shock Waves》2010,20(4):285-296
The mitigation of a planar shock wave caused by a cloud of calibrated water droplets was studied both experimentally and numerically.
Experiments were carried out, with different shock wave Mach numbers ranging from 1.1 to 1.8, in a vertical shock tube coupled
with a droplet generator which produced a well-characterized cloud of droplets of 120, 250 and 500 μm in diameter. By exploiting
such an experimental set-up, we successfully measured the attenuation of a normal shock wave when passing through the water
droplet cloud. This series of experiments allowed to identify the main parameters of this investigation and a clear dependence
between the attenuation of the shock wave and terms governing the regimes of droplet breakup has been found. On the other
hand, to support this experimental approach, 1D unsteady calculations were performed in similar configurations. Although the
mathematical model based on an Eulerian/Eulerian approach was actually incomplete, the first comparisons between the experiments
and the simulations were rather interesting and pointed out the need to improve the physical model, by taking into account
the fragmentation and the vaporization of the droplets submitted to the shock wave as well as the size distribution of the
water spray. 相似文献
7.
Robert D. Lotz Brian E. Thompson Christopher A. Konings Farhad Davoudzadeh 《国际流体数值方法杂志》1997,24(4):355-373
Numerical uncertainties are quantified for calculations of transonic flow around a divergent trailing edge (DTE) supercritical aerofoil. The Reynolds-averaged Navier–Stokes equations are solved using a linearized block implicit solution procedure and mixing-length turbulence model. This procedure has reproduced measurements around supercritical aerofoils with blunt trailing edges that have shock, boundary layer and separated regions. The present effort quantifies numerical uncertainty in these calculations using grid convergence indices which are calculated from aerodynamic coefficients, shock location, dimensions of the recirculating region in the wake of the blunt trailing edge and distributions of surface pressure coefficients. The grid convergence index is almost uniform around the aerofoil, except in the shock region and at the point where turbulence transition was fixed. The grid convergence index indicates good convergence for lift but only fair convergence for moment and drag and also confirms that drag calculations are more sensitive to numerical error. © 1997 by John Wiley and Sons, Ltd. 相似文献
8.
In transonic flow conditions, the shock wave/turbulent boundary layer interaction and flow separations on wing upper surface induce flow instabilities, ‘buffet’, and then the buffeting (structure vibrations). This phenomenon can greatly influence the aerodynamic performance. These flow excitations are self‐sustained and lead to a surface effort due to pressure fluctuations. They can produce enough energy to excite the structure. The objective of the present work is to predict this unsteady phenomenon correctly by using unsteady Navier–Stokes‐averaged equations with a time‐dependent turbulence model based on the suitable (k–ε) turbulent eddy viscosity model. The model used is based on the turbulent viscosity concept where the turbulent viscosity coefficient (Cμ) is related to local deformation and rotation rates. To validate this model, flow over a flat plate at Mach number of 0.6 is first computed, then the flow around a NACA0012 airfoil. The comparison with the analytical and experimental results shows a good agreement. The ONERA OAT15A transonic airfoil was chosen to describe buffeting phenomena. Numerical simulations are done by using a Navier–Stokes SUPG (streamline upwind Petrov–Galerkin) finite‐element solver. Computational results show the ability of the present model to predict physical phenomena of the flow oscillations. The unsteady shock wave/boundary layer interaction is described. Copyright © 2005 John Wiley & Sons, Ltd. 相似文献
9.
The starting of an axisymmetric convergent-divergent nozzle, with the result that supersonic flow is formed within almost the entire channel, is modeled, as applied to the hypersonic aerodynamic setup of the Institute of Mechanics of Moscow State University. A successful starting is realized when the nozzle is thrown in a uniform supersonic air flow at a fairly high Mach number. The steady flow structure is studied. It is numerically shown that in the convergent section of the channel there arises an oblique shock wave whose interaction with the nozzle axis leads to the formation of a reflected shock and a curvilinear Mach disk with a region of unsteady subsonic flow in the vicinity of the throat. The mathematical model is based on the two-dimensional Euler equations for axisymmetric gas flows. 相似文献
10.
Transition between regular and Mach reflection of shock waves: new numerical and experimental results 总被引:1,自引:0,他引:1
M.S. Ivanov D. Vandromme V.M. Fomin A.N. Kudryavtsev A. Hadjadj D.V. Khotyanovsky 《Shock Waves》2001,11(3):199-207
New numerical and experimental results on the transition between regular and Mach reflections of steady shock waves are presented.
The influence of flow three-dimensionality on transition between steady regular and Mach reflection has been studied in detail
both numerically and experimentally. Characteristic features of 3D shock wave configuration, such as peripheral Mach reflection,
non-monotonous Mach stem variation in transverse direction, the existence of combined Mach-regular-peripheral Mach shock wave
configuration, have been found in the numerical simulations. The application of laser sheet imaging technique in streamwise
direction allowed us to confirm all the details of shock wave configuration in the experiments. Close agreement of the numerical
and experimental data on Mach stem heights is shown.
Received 23 November 2000 / Accepted 25 April 2001 相似文献
11.
Jae-Soo Kim 《国际流体数值方法杂志》1993,17(7):567-587
The unsteady flow over an oscillatory NACA0012 aerofoil has been simulated by the calculation with Euler equations. The equations are discretized by an implicit Euler in time, and a second-order space-accurate TVD scheme based on flux vector splitting with van Leer's limiter. Modified eigenvalues are proposed to overcome the slope discontinuities of split eigenvalues at Mach = 0·0 and ± 1·0, and to generate a bow shock in front of the aerofoil. A moving grid system around the aerofoil is generated by Sorenson's boundary fitted co-ordinates for each time step. The calculations have been done for two angles of attack θ = 5·0° sin (ωt) and θ = 3·0° + 3·0° sin (ωt) for the free-stream Mach numbers 2·0 and 3·0. The results show that pressure and Mach cells flow along characteristic lines. To examine unsteady effects, the responses of wall pressure and normal force coefficients are analysed by a Fourier series expansion. 相似文献
12.
Charles Koeck 《国际流体数值方法杂志》1985,5(5):483-500
The unsteady Euler equations are numerically solved using the finite volume one-step scheme recently developed by Ron-Ho Ni. The multiple-grid procedure of Ni is also implemented. The flows are assumed to be homo-enthalpic; the energy equation is eliminated and the static pressure is determined by the steady Bernoulli equation; a local time-step technique is used. Inflow and outflow boundaries are treated with the compatibility relations method of ONERA. The efficiency of the multiple-grid scheme is demonstrated by a two-dimensional calculation (transonic flow past the NACA 12 aerofoil) and also by a three-dimensional one (transonic lifting flow past the M6 wing). The third application presented shows the ability of the method to compute the vortical flow around a delta wing with leading-edge separation. No condition is applied at the leading-edge; the vortex sheets are captured in the same sense as shock waves. Results indicate that the Euler equations method is well suited for the prediction of flows with shock waves and contact discontinuities, the multiple-grid procedure allowing a substantial reduction of the computational time. 相似文献
13.
Shock wave interaction with a sphere is one of the benchmark tests in shock dynamics. However, unlike wind tunnel experiments, unsteady drag force on a sphere installed in a shock tube have not been measured quantitatively. This paper presents an experimental and numerical study of the unsteady drag force acting on a 80 mm diameter sphere which was vertically suspended in a 300 mm x 300 mm vertical shock tube and loaded with a planar shock wave of M
s
= 1.22 in air. The drag force history on the sphere was measured by an accelerometer installed in it. Accelerometer output signals were subjected to deconvolution data processing, producing a drag history comparable to that obtained by solving numerically the Navier-Stokes equations. A good agreement was obtained between the measured and computed drag force histories. In order to interpret the interaction of shock wave over the sphere, high speed video recordings and double exposure holographic interferometric observations were also conducted. It was found that the maximum drag force appeared not at the time instant when the shock arrived at the equator of the sphere, but at some earlier time before the transition of the reflected shock wave from regular to Mach reflection took place. A negative value of the drag force was observed, even though for a very short duration of time, when the Mach stem of the transmitted shock wave relfected and focused at the rear stagnation point of the sphere.Received: 31 March 2003, Accepted: 7 July 2003, Published online: 2 September 2003 相似文献
14.
This paper describes a numerical simulation of bow shock formation ahead of a sphere at steady supersonic flow in the Mach
number range of 1.025–1.20. Turbulent viscous flow results are presented using the Spalart–Allmaras turbulence model. The
purpose of this study is to determine the shock standoff distance for a spherical projectile at slightly supersonic free flight
speeds. Results are compared to experimental data, including double exposure holographic interferograms obtained from a 40 mm
polycarbonate sphere launched by a light gas gun. The shock standoff distance was determined from the interferograms. The
present numerical simulations were found to agree with previously published data, and reached down to M = 1.025—a range where almost no previously published data exists. The computed flow structure and shock wave locations agree
well with recently obtained free-flight interferograms. 相似文献
15.
V. P. Zamuraev A. P. Kalinina 《Journal of Applied Mechanics and Technical Physics》2005,46(5):664-669
The possibility of controlling the aerodynamic characteristics of airfoils with the help of local pulsed-periodic energy addition into the flow near the airfoil contour at transonic flight regimes is considered. By means of the numerical solution of two-dimensional unsteady equations of gas dynamics, changes in the flow structure and wave drag of a symmetric airfoil due to changes in localization and shape of energy-addition zones are examined. It is shown that the considered method of controlling airfoil characteristics in transonic flow regimes is rather promising. For a zero angle of attack, the greatest decrease in wave drag is obtained with energy addition at the trailing edge of the airfoil.__________Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 46, No. 5, pp. 60–67, September–October, 2005. 相似文献
16.
In this paper, we propose a novel method for evaluating the frequency response of shock accelerometers using Davies bar and
interferometry. The method adopts elastic wave pulses propagating in a thin circular bar for the generation of high accelerations.
The accelerometer to be examined is attached to one end of the bar and experiences high accelerations of the order of 103∼105 m/s2. A laser interferometer system is newly designed for the absolute measurement of the bar end motion. It can measure the motion
of a diffuse surface specimen at a speed of 10−3 ∼100 m/s. Uncertainty of the velocity measurement is estimated to be±6×10−4 m/s, proving a high potential for use in the primary calibration of shock accelerometers. Frequency characteristics of the
accelerometer are determined by comparing the accelerometer's output with velocity data of the interferometry in the frequency
domain. Two piezoelectric-type accelerometers are tested in the experiment, and their frequency characteristics are obtained
over a wide frequency range up to several ten kilohertz. It is also shown that the results obtained using strain gages are
consistent with those by this new method.
Paper was presented at the 1992 SEM Spring Conference on Experimental Mechanics held in Las Vegas, NV on June 8–11. 相似文献
17.
This paper presents the results of an experimental study of the unsteady nature of a hypersonic separated turbulent flow.
The nomimal test conditions were a freestream Mach number of 7.8 and a unit Reynolds number of 3.5×107/m. The separated flow was generated using finite span forward facing steps. An array of flush mounted high spatial resolution
and fast response platinum film resistance thermometers was used to make multi-channel measurements of the fluctuating surface
heat trtansfer within the separated flow. Conditional sampling analysis of the signals shows that the root of separation shock
wave consists of a series of compression wave extending over a streamwise length about one half of the incoming boundary layer
thickness. The compression waves converge into a single leading shock beyond the boundary layer. The shock structure is unsteady
and undergoes large-scale motion in the streamwise direction. The length scale of the motion is about 22 percent of the upstream
influence length of the separation shock wave. There exists a wide band of frequency of oscillations of the shock system.
Most of the frequencies are in the range of 1–3 kHz. The heat transfer fluctuates intermittently between the undisturbed level
and the disturbed level within the range of motion of the separation shock wave. This intermittent phenomenon is considered
as the consequence of the large-scale shock system oscillations. Downstream of the range of shock wave motion there is a separated
region where the flow experiences continuous compression and no intermittency phenomenon is observed.
The project supported by National Natural Science Foundation of China 相似文献
18.
Separated Flow and Buffeting Control 总被引:2,自引:0,他引:2
D. Caruana A. Mignosi C. Robitaillié M. Corrège 《Flow, Turbulence and Combustion》2003,71(1-4):221-245
In transonic flow conditions, the shock wave/turbulent boundary layer interaction and the flow separations on the upper wing surfaces of civil aircraft induce flow instabilities, ‘buffet’ and then structural vibrations, ‘buffeting’. Buffeting can greatly affect aerodynamic behavior. The buffeting phenomenon appears when the aircraft's Machnumber or angle of attack increases. This phenomenon limits the aircraft's flight envelope. The objectives of this study are to cancel out or decrease the aerodynamic instabilities (unsteady separation, movement of the shock position) due to this type of flow by using control systems. The following actuators can be used: ‘Vortex Generators’ situated upstream of the shock position, a ‘Bump’ located at the shock position, and a new moving part designed by ONERA, situated on the trailing edge of the wing, the ‘Trailing Edge Deflector’ or TED. It looks like an adjustable ‘Divergent Trailing Edge’. It is an active actuator and can take different deflections or be driven by dynamic movements up to 250 Hz. Tests were performed in transonic 2D flow with models well equipped with unsteady pressure transducers. For high lift coefficients, a selected static position of the ‘Trailing Edge Deflector’ increases the wing's aerodynamic performances and delays the onset of buffet. Furthermore, in 2D flow buffet conditions, the ‘Trailing Edge Deflector’, driven by a closed-loop active control using the measurements of the unsteady wall static pressures, can greatly reduce buffet. The aerodynamic performances are not improved to the same extent by the bump actuator. From our experience, there is no effect on buffet or separated flow because of the incorrect positioning of the bump. All that can be observed is a local improvement on the intensity of the shock wave when the bump is very precisely situated at the shock position. Vortex generators have a great impact on the separated flow. The separated flow instabilities are greatly reduced and buffet is totally controlled even for strong instabilities. The aerodynamic performances of the airfoil are also greatly improved. 相似文献
19.
Nonlinearities arise in aerodynamic flows as a function of various parameters, such as angle of attack, Mach number and Reynolds number. These nonlinearities can cause the change from steady to unsteady flow or give rise to static hysteresis. Understanding these nonlinearities is important for safety validation and performance enhancement of modern aircraft. A continuation method has been developed to study nonlinear steady state solutions with respect to changes in parameters for two‐dimensional compressible turbulent flows at high Reynolds numbers. This is the first time that such flows have been analysed with this approach. Continuation methods allow the stable and unstable solutions to be traced as flow parameters are changed. Continuation has been carried out on two‐dimensional aerofoils for several parameters: angle of attack, Mach number, Reynolds number, aerofoil thickness and turbulent inflow as well as levels of dissipation applied to the models. A range of results are presented. Copyright © 2012 John Wiley & Sons, Ltd. 相似文献