scholarly journals Экспериментальное исследование нового способа уменьшения волнового сопротивления профиля при трансзвуковых скоростях

2020 ◽  
Vol 46 (12) ◽  
pp. 34
Author(s):  
М.А. Брутян ◽  
А.В. Волков ◽  
А.В. Потапчик
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Wave Drag ◽  
New Method ◽  
Optical Studies ◽  
Supercritical Airfoil ◽  
Aerodynamic Loads ◽  
Compression Waves ◽  
Supersonic Zone

A new method is proposed for wave drag reducing of a supercritical airfoil at transonic speeds, which is associated with the organization of a microwave section in a local supersonic zone on the upper surface. Simultaneously with optical studies, aerodynamic loads acting on a model were obtained. It is experimentally established that the proposed approach leads to the formation of a system of weak compression waves, to consecutive deceleration of the supersonic flow, to decrease of the Mach number in front of the shock wave, and, as a consequence, to weakening of its intensity and to reducing of the airfoil wave drag

Higher-Order Solutions for Unsteady Hypersonic and Supersonic Flow

1974 ◽  
Vol 25 (1) ◽  
pp. 59-68 ◽  
Author(s):  
W H Hui ◽  
J Hamilton
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Power Series ◽  
Unsteady Flow ◽  
The Body ◽  
Wedge Flow ◽  
First Order ◽  
Order Solution ◽  
Method Of Solution

SummaryThe problem of unsteady hypersonic and supersonic flow with attached shock wave past wedge-like bodies is studied, using as a basis the assumption that the unsteady flow is a small perturbation from a steady uniform wedge flow. It is formulated in the most general case and applicable for any motion or deformation of the body. A method of solution to the perturbation equations is given by expanding the flow quantities in power series in M−2, M being the Mach number of the steady wedge flow. It is shown how solutions of successive orders in the series may be calculated. In particular, the second-order solution is given and shown to give improvements uniformly over the first-order solution.

Simple Formulae for Supersonic Flow past a Cone

1975 ◽  
Vol 26 (1) ◽  
pp. 11-19 ◽  
Author(s):  
W H Hui
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Surface Pressure ◽  
Simple Formula ◽  
Pressure Coefficient ◽  
Circular Cone ◽  
Number Range ◽  
Thin Shock Layer ◽  
Cone Vertex

SummaryThe problem of the supersonic flow with attached shock wave past a circular cone at zero angles of attack is treated, using the thin-shock-layer expansion. The solution is calculated to the fourth approximation. A simple formula is then derived for the surface pressure coefficient by the application of the parameter-straining technique and it is shown to be very accurate for the whole Mach number range for which the shock remains attached to the cone vertex.

XXVII.—The Rotational Field behind a Bow Shock Wave in Axially Symmetric Flow using Relaxation Methods

1952 ◽  
Vol 63 (4) ◽  
pp. 371-380
Author(s):  
A. R. Mitchell ◽  
Francis McCall
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Bow Shock ◽  
Axially Symmetric ◽  
Relaxation Methods ◽  
Bow Shock Wave ◽  
Constant Vorticity ◽  
Complete Field ◽  
Rotational Field

SynopsisThe relaxation technique of R. V. Southwell is developed to evaluate mixed subsonic-supersonic flow regions with axial symmetry, changes of entropy being taken into account. In the problem of a parallel supersonic flow of Mach number I·8 impinging on a blunt-nosed axially symmetric obstacle, the new technique is used to determine the complete field downstream of the bow shock wave formed. Lines of constant vorticity and Mach number are shown in the field, and where possible a comparison is made with the corresponding 2-dimensional problem.

Iterative Solutions of the Equations Jor Plane Oblique Shock Waves

1959 ◽  
Vol 63 (587) ◽  
pp. 669-672 ◽  
Author(s):  
A. R. Collar
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Free Stream ◽  
Oblique Shock ◽  
Oblique Shock Wave ◽  
Two Dimensional ◽  
Iterative Solutions ◽  
Flow Through ◽  
Image Position

If a plane oblique shock wave, inclined to the free stream at the angle ε, is produced in two-dimensional supersonic flow of Mach number M by (for example) a wedge which deflects the flow through an angle δ, the equation connecting these quantities may be writtenIn this form, δ is given explicitly when M, ε are fixed. Similarly, we may obtain M explicitly when ε, δ are fixed; equation (1) may be written (see, for example, Liepmann and Puckett, Equation 4.27)

A study of the flow structure for Mach reflection in steady supersonic flow

2010 ◽  
Vol 656 ◽  
pp. 29-50 ◽  
Author(s):  
B. GAO ◽  
Z. N. WU
Keyword(s):  
Shock Wave ◽  
Supersonic Flow ◽  
Mach Reflection ◽  
Reflected Shock Wave ◽  
Inflexion Point ◽  
Compression Waves ◽  
Expansion Waves ◽  
Sonic Line ◽  
Reflected Shock ◽  
Starting Point

In this paper we study the waves generated over the slipline and their interactions with other waves for Mach reflection in steady two-dimensional supersonic flow. We find that a series of expansion and compression waves exist over the slip line, even in the region immediately behind the leading part of the reflected shock wave, previously regarded as a uniform flow. These waves make the leading part of the slipline, previously regarded as straight, deviate nonlinearly towards the reflecting surface. When the transmitted expansion waves from the upper corner first intersect the slipline, an inflexion point is produced. Downstream of this inflexion point, compression waves are produced over the slipline. By considering the interaction between the various expansion or compression waves, we obtain a Mach stem height, the shape and position of the slipline and reflected shock wave, compared well to computational fluid dynamics (CFD) results. We also briefly consider the case with a subsonic portion behind the reflected shock wave. The global flow pattern is obtained through CFD and the starting point of the sonic line is identified through a simple analysis. The sonic line appears to coincide with the first Mach wave from the upper corner expansion fan after transmitted from the reflected shock wave.

scholarly journals Возникновение усилий на клине при взаимодействии скользящего разряда со сверхзвуковым потоком

2021 ◽  
Vol 47 (9) ◽  
pp. 33
Author(s):  
В.М. Бочарников ◽  
В.В. Голуб ◽  
А.С. Савельев
Keyword(s):  
Shock Wave ◽  
Reynolds Number ◽  
Mach Number ◽  
Supersonic Flow ◽  
Wind Tunnel ◽  
Static Pressure ◽  
Discharge Energy ◽  
Pulse Heat ◽  
The Impact ◽  
Sliding Discharge

It was investigated generation of forces on the wedge during the interaction of a sliding discharge in the supersonic flow. Sliding discharge was generated in a supersonic atmospheric-vacuum wind tunnel with a flow Mach number M = 2, Reynolds number of about 106, and static pressure in the flow p = 0.15 bar. It was found that the impact of the discharge is not limited by the appearance of a shock wave, and also significantly changes the properties of the flow around the wedge due to pulse heat generation. As the discharge energy increases, the force from its impact on the wedge also increases. Force produced by the sliding discharge in the supersonic flow is several times greater than in the quiescent air regardless of its pressure: 0.15 bar or 1 bar.

The Displacement of the Nose Shock Wave from a Pitot Tube under a Streamwise Velocity Gradient

1967 ◽  
Vol 71 (682) ◽  
pp. 725-726 ◽  
Author(s):  
D. Sloan
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Supersonic Flow ◽  
Velocity Gradient ◽  
Streamwise Velocity ◽  
Uniform Flow ◽  
Pitot Tube ◽  
Pressure Reading

Summary:—The pressure reading of a pitot tube in a supersonic flow corresponds to the position of the nose shock wave. Results of measurements of the shock stand-off distance are presented for three different streamwise Mach number gradients and for a uniform flow. These results are compared with existing theoretical values.

Experimental Research of Gasdynamic Liquid Drops Breakup in the Supersonic Flow with an Oblique Shock Wave

2020 ◽  
Author(s):  
K. Yu. Arefyev ◽  
O. V. Guskov ◽  
A. N. Prokhorov ◽  
A. S. Saveliev ◽  
E. E. Son ◽  
...  
Keyword(s):  
Shock Wave ◽  
Supersonic Flow ◽  
Experimental Research ◽  
Oblique Shock ◽  
Oblique Shock Wave ◽  
Liquid Drops
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