اشترك بالحزمة الذهبية واحصل على وصول غير محدود شمرا أكاديميا
تسجيل مستخدم جديدA new similarity law for transonic-supersonic flow
89
0
0.0
(
0
)
تأليف
Luoqin Liu
اسأل ChatGPT حول البحث
ﻻ يوجد ملخص باللغة العربية
How to determine accurately and efficiently the aerodynamic forces of the aircraft in high-speed flow is one of great challenges in modern aerodynamics. In this Letter we propose a new similarity law for steady transonic-supersonic flow over thin bodies. The new similarity law is based on the local Mach number frozen principle. It depends on both the specific heat ratio and the free-stream Mach number. The new similarity law enables one to determine the lift and drag coefficients of the aircraft from that of a reference state which is more reachable. The validity of the new similarity law has been confirmed by the excellent agreement with numerical simulations of both two-dimensional airfoil flows and three-dimensional wing flows.
قيم البحث
اقرأ أيضاً
Within the transonic regime, the aeroelastic problems exhibit many unique characteristics compared with subsonic and supersonic cases. Although a lot of research has been carried out in this field, the underlying mechanisms of these complex phenomena
are not clearly understood yet, resulting in a challenge in the design and use of modern aircraft. This review summarizes the recent investigations on nonclassical transonic aeroelastic problems, including transonic buzz, reduction of transonic buffet onset, transonic buffeting response and frequency lock-in phenomenon in transonic buffet flow. After introducing the research methods in unsteady aerodynamics and aeroelastic problems, the dynamical characteristics as well as the physical mechanisms of these phenomena are discussed from the perspective of the fluid mode. In the framework of the ROM (reduced order model) -based model, the dominant fluid mode (or the eigenvalue) and its coupling process with the structural model can be clearly captured. The flow nonlinearity was believed to be the cause of the complexity of transonic aeroelasticity. In fact, this review indicates that the complexity lies in the decrease of the flow stability in the transonic regime. In this condition, the fluid mode becomes a principal part of the coupling process, which results in the instability of the fluid mode itself or the structural mode, and thus, it is the root cause of different transonic aeroelastic phenomena.
The flow structure obtained when Localized Arc Filament Plasma Actuators (LAFPA) are employed to control the flow issuing from a perfectly expanded Mach 1.3 nozzle is elucidated by visualizing coherent structures obtained from Implicit Large-Eddy Sim
ulations. The computations reproduce recent experimental observations at the Ohio State University to influence the acoustic and mixing properties of the jet. Eight actuators were placed on a collar around the periphery of the nozzle exit and selectively excited to generate various modes, including first and second mixed (m = +/- 1 and m = +/- 2) and axisymmetric (m = 0). In this fluid dynamics video http://ecommons.library.cornell.edu/bitstream/1813/13723/2/Alljoinedtotalwithmodetextlong2-Datta%20MPEG-1.m1v, http://ecommons.library.cornell.edu/bitstream/1813/13723/3/Alljoinedtotalwithmodetextlong2-Datta%20MPEG-2.m2v}, unsteady and phase-averaged quantities are displayed to aid understanding of the vortex dynamics associated with the m = +/- 1 and m = 0 modes excited at the preferred column-mode frequency (Strouhal number 0.3). The unsteady flow in both contains a broad spectrum of coherent features. For m = +/- 1, the phase-averaged flow reveals the generation of successive distorted elliptic vortex rings with axes in the flapping plane, but alternating on either side of the jet axis. This generates a chain of structures where each interacts with its predecessor on one side and its successor on the other. Through self and mutual interaction, the leading segment of each loop is pinched and passes through the previous ring before rapidly breaking up, and the mean jet flow takes on an elliptic shape. The m = 0 mode exhibits relatively stable roll-up events, with vortex ribs in the braid regions connecting successive large coherent structures.
The present work reports on the flow physics of turbulent supersonic flow over backward facing step (BFS) at Mach 2 using LES methodology where the dynamic Smagorinsky model is used for SGS modeling, while POD is invoked to identify the coherent stru
ctures present in the flow. The mean data obtained through the computations is in good agreement with the experimental measurements, while the iso-surfaces of Q-criterion at different time instants show the complex flow structures. The presence of counter rotating vortex pair in the shear layer along with the complex shock wave/boundary layer interaction leading to the separation of boundary layer is also evident from the contours of both Q and the modulus of vorticity. Further, the POD analysis reveals the presence of coherent structures, where the first and second modes confirm the vortical structures near the step as well as along the shear layer in the downstream region; while the second, third and fourth modes confirm the presence of vortices along the shear layer due to Kelvin-Helmholtz (K-H) instability. Moreover, POD as well as frequency analysis is extended at different planes to extract the detailed flow features.
We numerically examine the mechanisms that describe the shock-boundary layer interactions in transonic flow past an oscillating wing section. At moderate and high angles of incidence but low amplitudes of oscillation, shock induced flow separation or
shock-stall is observed accompanied by shock reversal. Even though the power input to the airfoil by the viscous forces is three orders of magnitude lower than that due to the pressure forces on the airfoil, the boundary layer manipulates the shock location and shock motion and redistributes the power input to the airfoil by the pressure forces. The shock motion is reversed relative to that in an inviscid flow as the boundary layer cannot sustain an adverse pressure gradient posed by the shock, causing the shock to move upstream leading to an early separation. The shock motion shows a phase difference with reference to the airfoil motion and is a function of the frequency of the oscillation. At low angles of incidence, and low amplitudes of oscillation, the boundary layer changes the profile presented to the external flow, leads to a slower expansion of the flow resulting in an early shock, and a diffused shock-foot caused by the boundary layer.
We are concerned with the stability of multidimensional (M-D) transonic shocks in steady supersonic flow past multidimensional wedges. One of our motivations is that the global stability issue for the M-D case is much more sensitive than that for the
2-D case, which requires more careful rigorous mathematical analysis. In this paper, we develop a nonlinear approach and employ it to establish the stability of weak shock solutions containing a transonic shock-front for potential flow with respect to the M-D perturbation of the wedge boundary in appropriate function spaces. To achieve this, we first formulate the stability problem as a free boundary problem for nonlinear elliptic equations. Then we introduce the partial hodograph transformation to reduce the free boundary problem into a fixed boundary value problem near a background solution with fully nonlinear boundary conditions for second-order nonlinear elliptic equations in an unbounded domain. To solve this reduced problem, we linearize the nonlinear problem on the background shock solution and then, after solving this linearized elliptic problem, develop a nonlinear iteration scheme that is proved to be contractive.
سجل دخول لتتمكن من نشر تعليقات
التعليقات
جاري جلب التعليقات
سجل دخول لتتمكن من متابعة معايير البحث التي قمت باختيارها
