scholarly journals Comparison of oblique shock wave angle in analytical and numerical solution

2019 ◽  
Vol 31 ◽  
pp. 137-142 ◽  
Author(s):  
Assen Marinov
Keyword(s):  
Shock Wave ◽  
Shock Waves ◽  
Wave Drag ◽  
Oblique Shock Wave ◽  
Friction Drag ◽  
Total Drag ◽  
Induced Drag ◽  
Drag And Lift ◽  
Wave Angle

The drag of the subsonic aircraft is largely formed by the skin friction drag and lift-induced drag. At transonic flight occurs shock wave. Determination of shock wave angle is important part of design of every aircraft, which working in supersonic airflow regimes. Formation of shock waves cause formation the wave drag. The wave drag could account about 35% from total drag of aircraft. Shock wave angle is directly linked with the intensity of itself. This work compares shock wave angle calculations using analytical and numerical solving methods.

Numerical Study in Wing Tip Vortex for a Modified Commercial Boeing Aircraft

2008 ◽  
Author(s):  
Ricardo Hernandez-Rivera ◽  
Abel Hernandez-Guerrero ◽  
Cuauhtemoc Rubio-Arana ◽  
Raul Lesso-Arroyo
Keyword(s):  
Numerical Study ◽  
Tip Vortex ◽  
Total Drag ◽  
The Core ◽  
Induced Drag ◽  
Wing Tip Vortex ◽  
Drag And Lift ◽  
Wing Tips ◽  
Wing Tip ◽  
Constant Mach Number

Recent studies have shown that the use of winglets in aircrafts wing tips have been able to reduce fuel consumption by reducing the lift-induced drag caused by wing tip vortex. This paper presents a 3-D numerical study to analyze the drag and lift forces, and the behavior of the vortexes generated in the wing tips from a modified commercial Boeing aircraft 767-300/ER. This type of aircraft does not contain winglets to control the wing tip vortex, therefore, the aerodynamic effects were analyzed adding two models of winglets to the wing tip. The first one is the vortex diffuser winglet and the second one is the tip fence winglet. The analyses were made for steady state and compressible flow, for a constant Mach number. The results show that the vortex diffuser winglet gives the best results, reducing the core velocity of the wing tip vortex up to 19%, the total drag force of the aircraft up to 3.6% and it leads to a lift increase of up to 2.4% with respect to the original aircraft without winglets.

Approximate formula of weak oblique shock wave angle

AIAA Journal ◽  
1992 ◽  
Vol 30 (3) ◽  
pp. 837-839 ◽  
Author(s):  
Hua-Shu Dou ◽  
Hsueh-Ying Teng
Keyword(s):  
Shock Wave ◽  
Approximate Formula ◽  
Oblique Shock ◽  
Oblique Shock Wave ◽  
Wave Angle

Oblique shock waves in a dusty-gas flow over a wedge

1986 ◽  
Vol 408 (1834) ◽  
pp. 61-78 ◽  
Keyword(s):  
Shock Wave ◽  
Gas Flow ◽  
Deflection Angle ◽  
Equations Of Motion ◽  
Gas Analysis ◽  
Oblique Shock ◽  
Oblique Shock Wave ◽  
Dusty Gas ◽  
Wave Angle ◽  
Flow Deflection

Supersonic flows of a dusty gas past a wedge are studied theoretically. An oblique shock wave emanates from the apex of the wedge at the same angle as in the case of a pure gas, but bends back because of the presence of the particles. It is shown from an equilibrium-gas analysis that the extent of decrease in the shock-wave angle is larger for smaller velocity of the uniform stream. When the flow-deflection angle is small enough, the oblique shock wave developing fully at large distances from the apex has a fully dispersed transition structure. On the other hand, it is partly dispersed when the flow-deflection angle is large. Details of the development of the oblique shock wave as the distance from the apex increases are clarified by solving the equations of motion numerically. The particles colliding with the wedge are assumed to stick to or reflect elastically from its surface. It is shown that the reflected particles affect the flow significantly in the neighbourhood of the wedge.

Reconsideration of oblique shock wave reflections in steady flows. Part 2. Numerical investigation

1995 ◽  
Vol 301 ◽  
pp. 37-50 ◽  
Author(s):  
J. Vuillon ◽  
D. Zeitoun ◽  
G. Ben-Dor
Keyword(s):  
Shock Wave ◽  
Shock Waves ◽  
Numerical Investigation ◽  
State Of The Art ◽  
Oblique Shock Wave ◽  
Bottom Surface ◽  
Steady Flows ◽  
Wave Reflections ◽  
The One ◽  
Reflecting Surfaces

The reflection of shock waves over straight reflecting surfaces in steady flows was investigated numerically with the aid of the LCPFCT algorithm. The findings completely supported the experimental results which were reported in Part 1 of this paper (Chpoun et al. 1995). In addition, the dependence of the resulting shock wave configuration on the distance between the trailing edge of the reflecting wedge and the bottom surface, inside the dual-solution domain, was studied. As a result of this study, as well as the one reported in Part 1, the state of the art of shock wave reflections in steady flows was reconsidered.

Accurate decomposition method and identification technology of drag

2020 ◽  
Vol 34 (14n16) ◽  
pp. 2040114
Author(s):  
Jia-Lei Yu ◽  
Jing Jin ◽  
Chun Shao ◽  
Miao Zhang ◽  
Tie-Jun Liu
Keyword(s):  
Shock Wave ◽  
Integral Method ◽  
Physical Mechanism ◽  
Wave Drag ◽  
Irreversible Processes ◽  
Viscous Drag ◽  
Numerical Dissipation ◽  
Far Field ◽  
Induced Drag ◽  
Field Integral

Aiming at accurate decomposition and identification of drag, the drag prediction technology based on the mid/far-field integral method is developed. The method decomposes the far-field drag into entropy drag and induced drag according to its physical mechanism, and introduces an appropriate entropy correction to eliminate the numerical dissipation by analyzing the influence of the trailing integral section position on the entropy drag calculation. Based on the analysis of thermodynamic reversible processes and irreversible processes, the drag is refined into viscous drag, shock wave drag, induced drag and pseudo-drag. The mid-field integral method is used to calculate the separate contribution of viscous drag, shock wave drag and induced drag by calculating the limited integral domain. Numerical results show that the developed method is feasible in accurately reflecting the physical mechanism and predicting the drag ratio. Thus, it provides a reliable tool for drag reduction of large passenger aircraft.

The non-linear refraction of shock waves by upstream disturbances in steady supersonic flow

1970 ◽  
Vol 43 (1) ◽  
pp. 1-33 ◽  
Author(s):  
Sheldon Weinbaum ◽  
Arnold Goldburg
Keyword(s):  
Shock Wave ◽  
Shock Waves ◽  
Interaction Parameter ◽  
Wave Interaction ◽  
Oblique Shock ◽  
Rational Approach ◽  
Oblique Shock Wave ◽  
Local Ratio ◽  
Non Linear ◽  
Shock Refraction

The general problem studied is the propagation of an oblique shock wave through a two-dimensional, steady, non-uniform oncoming flow. A higher-order theory is developed to treat the refraction of the incident oblique shock wave by irrotational or rotational disturbances of arbitrary amplitude provided the flow is supersonic behind the shock. A unique feature of the analysis is the formulation of the flow equations on the downstream side of the shock wave. It is shown that the cumulative effect of the downstream wave interactions on the propagation of the shock wave can be accounted for exactly by a single parameter Φ, the local ratio of the pressure gradients along the Mach wave characteristic directions at the rear of the shock front. The general shock refraction problem is then reduced to a single non-linear differential equation for the local shock turning angle θ as a function of upstream conditions and an unknown wave interaction parameter Φ. To lowest order in the expansion variable θΦ, this equation is equivalent to Whitham's (1958) approximate characteristic rule for the propagation of shock waves in non-uniform flow. While some further insight into the accuracy of Whitham's rule does emerge, the theory is not a selfcontained rational approach, since some knowledge of the wave interaction parameter Φ must be assumed. Analytical and numerical solutions to the basic shock refraction relation are presented for a broad range of flows in which the principal interaction occurs with disturbances generated upstream of the shock. These solutions include the passage of a weak oblique shock wave through: a supersonic shear layer, a converging or diverging flow, a pure pressure disturbance, Prandtl–-Meyer expansions of the same and opposite family, an isentropic non-simple wave region, and a constant pressure rotational flow. The comparison between analytic and numerical results is very satisfactory.

scholarly journals Comparation of Effects Generated by Shock Waves of Selected Explosives

2014 ◽  
Vol 8 (2) ◽  
pp. 24-32
Author(s):  
Kamil Boc ◽  
Štefan Jangl ◽  
Dagmar Vidriková ◽  
Martin Vysocký
Keyword(s):  
Shock Wave ◽  
Shock Waves ◽  
Wave Effects ◽  
Explosive Device

Abstract Article deals with assessment of exactly acquired values of shock wave effects generated in the course of explosion of an explosive device and with comparation of the current options for determination of magnitude of their effect by the means of mathematic models using experimentally measured values.

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