Analysis of Unstarted Supersonic Flutter in Cascade by Semiactuator Disk Theory.

In this paper, some preliminary experiments and the calculations were performed to clarify the flow field, in which the rods surrounded by a porous cavity were normally inserted into the main supersonic flow. As a result, it is found that the starting shock wave severely interacts with the rods, the bow shock wave, its reflections, and the porous wall, which are numerically well predicted under some conditions. In case of the single rod, the main supersonic flow is sucked in front of the rod and blown downstream of the rod. Furthermore, the suction mass flux through the porous holes depends on the number and the locations of the rods. On the other hand, the three rods aligned parallel to the flow make favorable suction area downstream of the rods and also generate as less total pressure loss as that for a single rod. In addition, the spanwised three rods cause the strongest bow shock wave making the large mass flux through the porous wall and also large total pressure loss downstream of the rods. The calculation also suggests that interaction between the bow shock wave and boundary layer along the upper wall results in the boundary layer separation.

A numerical study was carried out for the combustion induced by oblique shock wave over a two-dimensional wedge of finite length. A compressible fluid dynamics code was used with a detailed hydrogen-oxygen combustion mechanism for the numerical study. For an off-attaching condition of oblique detonation waves (ODW) predicted by Rankine-Hugoniot theory with equilibrium chemistry, periodically oscillating combustion was induced by oblique shock wave for certain conditions. Investigation of the numerical results presents that the mechanism of the periodically oscillating combustion is an interaction between shock and reaction waves coupled with chemical kinetic effects. The oscillatory motion at the wedge surface was similar to the periodic motion of one dimensional detonation wave. For the investigation of the criterion of a periodically oscillating combustion, a series of numerical simulations were carried out for wedges of different lengths to change the ratio between fluid dynamic time scale and chemical time scale. At the offattaching condition of ODW predicted by RankineHugoniot theory with equilibrium chemistry, present results reveals three different regimes of combustion. (1) If a fluid dynamic residence time is shorter than a chemical induction time behind an oblique shock wave, combustion may not occur or incomplete shock-induced combustion was observed. (2) If a fluid dynamic residence time is longer than a chemical induction time behind an oblique shock wave, but the former is shorter than the later behind a bow shock wave at same Mach number, oblique shock wave still attaches at the wedge nose, even at the off-attaching condition of oblique detonation waves, and oscillatory combustion was observed. (3) If a fluid dynamic residence time is longer than a chemical induction time behind a bow shock wave, the oblique shock wave is no longer exist and detached bow shock-induced combustion or a detached overdriven detonation wave was observed. Introduction Oblique detonation wave (ODW) stabilized over a body was given interests for a last decade, since is it considered as a promising combustion mechanism for novel hypersonic propulsion systems such as a ram accelerator or an oblique detonation wave engine. Classical theory of ODW with Rankine-Hugoniot theory assuming immediate heat addition behind an oblique shock wave was summarized by Pratt[1], and Shepherd[2] summarized recent theoretical and experimental results of ODW for propulsion application. 0 5 10 15 20 25 30 35 40 45 0 15 30 45 60 75 90

Reentry capsule configurations significantly differ from each other due to entry conditions and mission requirements. This paper describes numerical simulations of the viscous flow over the Beagle and the OREX (Orbital Reentry EXperiments) configurations for freestream Mach numbers in the range of 1.2 - 5.0. The flow fields over the reentry modules are obtained by solving time-dependent axisymmetric compressible laminar Navier-Stokes equations. The fluid mechanics equations are discretized in spatial coordinates employing a finite volume method, which reduces the governing equations to semi-discretized ordinary differential equations. Temporal integration is carried out using a two-stage Runge-Kutta time-stepping scheme. A local time-stepping is employed to get the steady state solution. The numerical simulation is performed on a single-block structured grid. The flow field features around the reentry capsules such as bow shock wave, sonic line, expansion fan and recirculation flow in the base-shell region are well captured by the present numerical computation. The effects of the geometrical parameters of the module, such as aspect ratio, frontal segment bluntness, fore body cone angle and shoulder rounding radius on the wake throat width, distance from the wake throat to the base of the model, flow departure angle and aerodynamic drag are analyzed using the numerically simulated flow field.

Transmission and reflection of acoustic and entropy waves through a stator–rotor stage

The flowfield over a sharp-tipped double cone axisymmetric configuration is carried out by solving time dependent axisymmetric laminar compressible Navier-Stokes equations for freestream Mach number range of 2.0 - 6.0. The fluid dynamics equations are discretized in spatial coordinates using integral formulation in conjunction with finite volume method which reduces the governing equations to semi-discretized ordinary differential equations. Temporal integration is performed employing multistage Runge-Kutta time-stepping scheme. A local time step is used to achieve steady-state solution. The numerical computation is carried out on a mono-block with structured grids. The flowfield features over the sharp-tipped double cone configuration such as conical shock wave, separation region, separation and reattachment shock wave and bow shock wave in the cone region, expansion fan, and recirculation flow in the base region are well captured at Mach number 3 and 6 which are identical to the Edney VI type shock interaction. The velocity vector, Mach contours plots and variations of surface pressure coefficient along the sharp-tipped double cone configuration are analyzed at various Mach numbers. The fore-body aerodynamic drag is calculated employing computed pressure distribution. The paper presents the influence of the freestream Mach number on the flows with shock interactions over the sharp-tipped double cone geometry.

A preliminary inviscid version of the NASA LeRC/Mississippi State University 3-D aeroelastic computational fluid dynamics code, TURBO-AE, is evaluated using operating conditions from three test cases from an AlliedSignal Engines (AE) advanced technology integrally bladed fan rotor. Two of the cases corresponded to conditions where high op-line flutter vibrations were observed, and the third case was flutter-free during testing. TURBO-AE steady (no blade vibration) computations were performed for each test case and a reasonable match was found with the steady 3-D viscous code DAWES. Following completion of the steady computations, TURBO-AE flutter (blade vibration) analyses were completed by harmonically oscillating the airfoil with mode shapes computed by the structural finite-element code ANSYS. Aerodynamic work per vibration cycle and modal damping was computed by TURBO-AE at 0° and 180° Inter Blade Phase Angles (IBPA). For IBPA = 0°, TURBO-AE predicted aerodynamic damping to be low but positive. Further, for IBPA = 180°, the TURBO-AE predictions indicate a very high level of aerodynamic damping. Flutter was observed on the rotor only for IBPA = 32.7°, and TURBO-AE analyses are in process for this condition. Unsteady pressure perturbation contour plots are presented to aid in the visualization of the TURBOAE flow solution. Results presented in this paper indicate that, although TURBO-AE still requires significant development, it is making the difficult transition from a research code to a flutter and forced-response prediction tool for the U.S. aeropropulsion community.

A series of 1.27 m air gap discharge tests under negative switching impulses were carried out to characterize shock wave expansion behavior along leader channels. A high-speed Schlieren system was used to identify shock wave propagation characteristics near the high-voltage electrode. According to the Schlieren images, three distinct types of shock waves were recognized during the leader propagation and after the air gap breakdown: (1) spherical shock wave (SSW), (2) cylindrical shock wave (CSW), and (3) bow shock wave (BSW). To the best of our knowledge, it is the first time to recognize SSWs at the point where the space leader meets the main channel and BSWs along the leader branch after the main leader bridges the air gap. According to the proposed velocity calculation method, the propagation velocity of SSW and CSW was calculated as a function of the shock radius, as well as the shock wave head of BSW. Moreover, a transient magnetohydrodynamics model was developed to predict shock wave front propagation characteristics of both straight and bending channels after breakdown. The predicted shock wave propagation velocity and shock radius showed good agreement with Schlieren images.

New VLA maps at 2 cm of the 13-arcsec-scale linear radio source in the center of NGC 1068 are described. The northeast lobe shows a limb-brightened conical morphology, very sharp leading edges, and a magnetic field running parallel to these edges. The spectral index between 2 and 6 cm in these line-brightened regions is near 1.0. The northeast subpeak has a very steep radio spectrum between 18 and 2 cm which is attributed to inverse Compton losses of the relativistic electrons on the infrared photons. The spectral indices in the southwest lobe lie in the range 0.9-1.5 except in its northern parts, where a much larger index is found. The northeast lobe radio emission could arise in either the cocoon of old jet material which has passed through the internal shock in the ejecta and blown out to either side, or in interstellar material compressed by a bow shock wave driven into the galactic ISM. 45 references.

Grid fin is widely used in the field of aerospace because of its unique advantages in the aerodynamic and structural performance. Meanwhile, the aerodynamic and structural characteristics of the grid fin are quite different from those of the conventional rudder. Deep understanding of the aeroelastic characteristics and mechanism of the grid fin is of great reference value for the fine design of grid fin. In this paper, an efficient reduction-order-method-based aeroelastic analysis method is established, and the method is verified by the flutter case of AGARD 445.6 wing. Finally, based on this method, the flowfield characteristics and the variation of flutter characteristics with Mach number are studied in detail. The results show that in the subsonic region, the transonic flutter “dip” of the grid fin occurs earlier than that of the conventional wing due to the acceleration effect of the grid fin on the airflow. In the transonic region, when the flowfield is characterized by a bow shock wave, the critical flutter speed increases slightly with the increase of Mach number. However, in the supersonic region, when the flowfield is characterized by an oblique shock wave, the flutter critical velocity increases substantially with the increase of Mach number. It is found that the variation law of the flutter boundary with Mach number is very similar to that of the normal force and the pitching moment with Mach number, and there is a close relationship between them. The ratio of flutter speed of the grid fin to the freestream velocity will reach the minimum value near . At the same time, it is discovered that the flutter instability of the grid fin is mainly caused by the coupling of the bending mode and torsional mode of the grid fin, and the instability mode is the torsional mode. Meanwhile, it should be noted that the flutter coupling characteristics of the grid fin differ greatly from those of the conventional rudder surface due to the great difference in structural characteristics and unsteady aerodynamic characteristics.

In this research, the effects of some liquid side injection from nozzle wall into exhaust gas of combustion chamber are studied. The side injection against main flow is as elliptical solid thing that change the symmetric of flow field on nozzle wall and causes some different pressure distribution on wall, and finally causes thrust vector deviation. Flows interaction causes some physical phenomena as bow shock wave in front of injection region. This paper explain the effects of this wave and variation velocity & pressure distribution at different cross sections of flow field and comparison results of air and other liquid fluid in thrust vector control system. The results are compared with experimental data and have well agreement with them. The results show that Freon is one of best injection liquid for this type of thrust vector control. Performance of Injection is optimum in relative position 35 to 40% nozzle divergence length.

Generation and development of disturbances in a hypersonic viscous shock layer on a flat plate is studied both experimentally and numerically. The study is performed at the Mach number M∞ = 21 and the Reynolds number ReL = 1.44 × 105 and is aimed at elucidating the physical mechanisms that govern the receptivity and instability of the shock layer at extremely high hypersonic velocities. The experiments are conducted in a hypersonic nitrogen-driven wind tunnel. An electron-beam fluorescence technique, a Pitot probe and a piezoceramic transducer are used to measure the mean density and Mach number contours, as well as density and pressure fluctuations, their spectra and spatial distributions in the shock layer. Direct numerical simulations are performed by solving the Navier–Stokes equations with a high-order shock-capturing scheme in a computational domain including the leading and trailing edges of the plate, so that the bow shock wave and the wake behind the plate are also simulated. It is demonstrated that computational and experimental data characterizing the mean flow field, intensity of density fluctuations and their spatial distributions in the shock layer are in close agreement. It is found that excitation of the shock layer by external acoustic waves leads to generation of entropy–vortex disturbances with two maxima of density fluctuations: directly behind the shock wave and on the external edge of the boundary layer. At the same time, the pressure fluctuations decay inward into the shock layer, away from the shock, which agrees with the linear theory of interaction of shock waves with small perturbations. Thus, the entropy–vortex disturbances are shown to dominate in the hypersonic shock layer at very high Mach numbers, in contrast with the boundary layers at moderate hypersonic velocities where acoustic modes are most important. A parametric numerical study of wave processes in the shock layer induced by external acoustic waves is performed with variations of frequency, amplitude and angle of propagation of external disturbances. The amplitude of generated disturbances is observed to grow and decay periodically along the streamwise coordinate, and the characteristics of these variations depend on the frequency and direction of incident acoustic waves. The hypersonic shock layer excited by periodic blowing and suction near the leading edge is also investigated; in the experiments, this type of excitation is obtained by using an oblique-cut whistle. It is shown that blowing/suction generates disturbances resembling those generated by external acoustic waves, with similar spatial distributions and phase velocities. This result paves the way for active control of instability development in the shock layer by means of destructive interference of two types of disturbances. Numerical simulations are performed to show that instability waves can be significantly amplified or almost entirely suppressed, depending on the relative phase of blowing/suction and acoustic disturbances. Wind-tunnel experiments completely confirm this numerical prediction. Thus, the feasibility of delaying instability development in the hypersonic shock layer has been demonstrated for the first time.

The hypersonic flow field around a blunt cone was simulated using a high-order finite difference method. Fast acoustic waves, slow acoustic waves, entropy waves, and vortical waves were introduced into the free-stream to determine the influence of a free-stream with disturbances on the hypersonic flow field and boundary layer. The effect of disturbance type on the evolution of perturbations in the hypersonic boundary layer was analyzed. Fast Fourier Transform was adopted to analyze the effect of the disturbance type on the evolution of different modes in the boundary layer. A roughness element was introduced into the flow field to reveal the impact of the roughness element on hypersonic boundary layer receptivity. The results showed that a free-stream with disturbances affected the hypersonic flow field and boundary layer; acoustic waves had the greatest influence. The impact of slow acoustic waves on the flow field was mainly concentrated in the region between the shock and the boundary layer, whereas the influence of fast acoustic waves was mainly concentrated in the boundary layer. Multi-mode perturbations formed in the boundary layer were caused by the free-stream with disturbances, wherein the fundamental mode was the dominant mode of the perturbations in the boundary layer caused by fast acoustic waves, entropy waves, and vortical waves. The dominant modes of the perturbations in the boundary layer caused by slow acoustic waves were both the fundamental mode and the second harmonic mode. The roughness element changed the propagation process of different modes of perturbations in the boundary layer. In the downstream region of the roughness element, perturbations in the boundary layer caused by the slow acoustic waves had the greatest influence. The second harmonic mode in the boundary layer was significantly suppressed, and the fundamental mode became the dominant mode. The effects of fast acoustic waves and entropy waves on the boundary layer receptivity were similar, except the amplitude of the perturbations in the boundary layer caused by the fast acoustic waves was larger.

Experimental investigations of the aerodynamic damping of compressor blades are usually performed by forcing the blades externally to a motion that is similar to a given mode shape and inter-blade phase angle (IBPA) while measuring the motion-induced unsteady pressure distribution. Evaluating this to an aerodynamic work entry from the fluid to the blade, at least a local contribution to the flutter (in-)stability can be determined. Test rigs are usually non-rotating linear or rotating annular cascade wind tunnels. In the latter case, besides measuring in and transmitting data out of the rotating system, the excitation of the blades themselves is still a challenge. In the present case a blisk rotor with realistic fan blade geometries and flow conditions was investigated aeroelastically. For the excitation of the 1st bending (1F) mode shape of the blading a sector of five blades was chosen. In this sector the natural vibration of the blading, represented by constant vibration amplitudes and a given IBPA should be simulated. Therefore the blades were equipped with Macro Fiber Composites (MFC). These foils of piezoelectric material expand and contract due to the applied high voltage. A control system was developed to adjust the amplitude and the IBPA of the blade vibration. For the transmission of the high voltage, a separate data transmission system on the base of liquid metals was chosen. The blade vibration was measured by strain gauges and additionally monitored by a specific rig system. The aeroelastic investigations were carried out in the compressor test facility M2VP of the DLR Institute of Propulsion Technology in Cologne. During the measurement, the MFCs were able to excite the blades to a certain extent. The paper will present the technique to excite the blades of a compressor blisk by means of MFCs as well as achieved vibration amplitudes and limitations under engine-like operating conditions.

We present the results of 3-D MHD simulations of collisions between an equilibrium current and a plasmoid. Three typical equilibrium configurations were analyzed. Our simulation results show that when a plasmoid approaches a current loop, an active region is created in front of the plasmoid bounded on the front side by a bow shock wave and on the back side by a reverse shock wave. The collision process modifies the current system and a strong electric field is also induced in the active region. An additional magnetic field generated by the induced current upsets the initial equilibrium condition. As a result, the whole loop is compressed and heated. We found that when the plasmoid approaches the loop, before reaching it the induced electric field amounts to its maximum value. The current loop is curved under the collision. The core of the plasmoid can not drive into the loop, it is sprung back by a magnetic counterpressure. This collision process between a plasmoid and a current loop may be responsible for the triggering of a solar flare observed by Yohkoh.

In high-velocity high-altitude flight, the thick boundary layer on the surface of a flying vehicle merges with the bow shock wave so that, even at a large distance from a leading edge, they may form the so-called viscous shock layer (VSL). Like the boundary layer, the laminar VSL is unstable, and flow disturbances arising and developing in this layer induce its transition to the turbulent flow regime. However, the mechanisms governing their emergence and development in the hypersonic VSL may differ substantially from those investigated in supersonic flows with lower Mach numbers [1, 2]. A better understanding of the receptivity and instability mechanisms is a necessary condition for the development of efficient methods for controlling the laminar-turbulent transition in a hypersonic flow around flying vehicles.