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Results of a numerical study of supersonic flows formed in channels with a square cross section are reported. Configurations consisting of a constricting entrance section formed by four compression wedges aligned at a right angle to each other and a subsequent channel with a constant cross section are considered. The initial shock waves formed on the nose compression wedges pairwise intersect each other in dihedral corners of the configuration along the swept lines, and a complex system of reflected and subsequent interacting shock waves of the general three-dimensional position is formed further downstream. The data are obtained in a supersonic range of freestream Mach numbers M = 2–4 for compression wedge angles of 3° and 8° in flows with both regular and irregular interaction of the initial shock waves in dihedral corners of the configuration. [ABSTRACT FROM AUTHOR]

The paper presents the results of a numerical study of gas dynamics and integral parameters of the flow at the channel entrance located behind a conical or plane shock wave. The freestream Mach number range is M = 2–4 and the range of the angles of the compression surfaces of the wedge and cone is δ= 10–90°. Data on the flow structure at the channel entrance, mass-averaged Mach numbers, total pressure loss coefficient, and flow rate coefficients are obtained. A comparative analysis of these parameters is performed, and the advantages and drawbacks of the channel entrance positions in various types of the flow are noted. [ABSTRACT FROM AUTHOR]

Numerical investigations of the asymmetry of a fluid flow in the channel of an air intake which is due to the interaction of a compression shock, generated by the shell ring of the air-intake channel, with the boundary layer on the longitudinally ribbed lower surface of this channel have been performed on the basis of the Reynolds-averaged Navier–Stokes equations and the SST k–ω model of turbulence. Data on the structure of the fluid flow in such a channel with a narrowing inlet part and a next part of constant cross section, the limiting flow lines on the longitudinally ribbed lower surface of the channel, and the profiles of the velocity of the flow and the total pressure in the boundary layer in the channel have been obtained for the case where the Mach number of the fluid flow incoming to the channel is M = 4. [ABSTRACT FROM AUTHOR]

The paper presents the results of a numerical simulation of conditions for starting a supersonic air intake using the bypass of part of the captured air through longitudinal slots. The air-intake duct consists of a narrowing entrance section and a constant-cross-section throat, both elements being constructed using flat surfaces, and the air intake contains longitudinal slots stretching along the lower flat surface. Various variants of the sectional bleeding of air are investigated, and changes in the flow structure in the intake duct depending on the air bleeding variant are demonstrated. Numerical modeling of the three-dimensional flow was performed for free-stream Mach number M∞ = 4 on the basis of Reynolds-averaged Navier-Stokes equations and the k−ω SST model of turbulence. [ABSTRACT FROM AUTHOR]

Results of numerical simulations of a supersonic flow in a channel with a rectangular cross section having a longitudinally ribbed bottom surface are reported. The channel contains a tapered inlet section, where the cowl lip generates a shock wave incident onto the bottom surface, and a posterior constant-area section. The present numerical simulations are performed for a viscous three-dimensional flow with the use of the Reynolds-averaged Navier-Stokes equations and k-ω SST turbulence model. The computations are performed for the free-stream Mach number M = 4. [ABSTRACT FROM AUTHOR]

Results of a numerical study of the flow structure in the rectangular duct of a model air inlet with the interaction of the shock wave generated by the inlet cowl with the turbulent boundary layer on the inner duct surfaces are reported. A configuration involving a convergent entrance section and a constant-cross-section throat was analyzed. The numerical simulation of the three-dimensional flow with free-stream Mach number M = 4 was performed on the basis of the Reynolds-averaged Navier-Stokes equations and the k-ω SST model of turbulence. As a result of the calculations, the limit streamlines and the distributions of static pressure on the duct surfaces, as well as the distributions of total pressure and the fields of Mach number in duct cross sections were obtained. The structure of the flow is analyzed and the possibility of occurrence of longitudinal vortex structures in the corner region of the intake-duct throat is shown. [ABSTRACT FROM AUTHOR]

The results of the numerical modeling of the supersonic flow in an axisymmetric duct in which a pseudo-shock arises are presented. The duct includes the frontal inlet with a funnel-shaped part of initial compression of the supersonic flow and with a cylindrical throat part as well as the subsequent (cylindrical or diverging) diffuser where the flow slows down to a subsonic velocity. The flow conditions at the freestream Mach number M = 6 have been considered. Numerical computations of the flow have been done using a Navier-Stokes equations code and the k-ω SST turbulence model. As a result of computations, such flow parameters have been determined as the location of the pseudo-shock beginning, the length of the pseudo-shock supersonic part, the pressure distribution on the duct wall, the total pressure losses as well as the characteristics of flow turbulence. In particular, the variation of the turbulence intensity and turbulent viscosity along the pseudo-shock length have been examined and, based on these characteristics, the possibility of determining the location of a cross section, in which the pseudo-shock can be treated as completed, have been considered. [ABSTRACT FROM AUTHOR]

Results of calculations of the sonic boom produced by a supersonic passenger aircraft in a cruising regime of flight at the Mach number M = 2.03 are presented. Consideration is given to the influence of the lateral dihedral of the wings and the angle of their setting, and also of different locations of the aircraft engine nacelles on the wing. An analysis of parametric calculations has shown that the intensities of sonic boom generated by a configuration with a dihedral rear wing and by a configuration with set wings remain constant, in practice, and correspond to the intensity level created by the optimum configuration. Comparative assessments of sonic boom for tandem configurations with different locations of the engine nacelles on the wing surface have shown that the intensity of sonic boom generated by the configuration with an engine nacelle on the windward side can be reduced by ~14% compared to the configuration without engine nacelles. In the case of the configuration with engine nacelles on the leeward size of the wing, the profile of the sonic-boom wave degenerates into an N-wave, in which the intensity of the bow shock is significantly reduced. [ABSTRACT FROM AUTHOR]

The results of the numerical modeling of a flow with a pseudo-shock in an axisymmetric duct are presented. The duct included a frontal inlet with the initial funnel-shaped compression part and the cylindrical throat part as well as the subsequent expanding diffuser. To create a flow with a pseudo-shock, the duct was throttled with the use of the outlet converging insert. Numerical computations of the axisymmetric flow have been conducted on the basis of the solution of the Reynolds-averaged Navier−Stokes equations and with the use of the k-ω SST turbulence model. As a result of computations, such parameters of the flow were determined as the location of the beginning of the pseudo-shock, the length of its supersonic part, the velocity profiles in different cross sections of the pseudo-shock, the pressure distribution on the duct wall, the total pressure recovery factor, and others. The behavior of these parameters at the freestream Mach number М = 6 was analyzed versus the diffuser opening angle and different degree of the inlet duct throttling. [ABSTRACT FROM AUTHOR]

Aerodynamic design of an innovative supersonic three-dimensional air-intake including a starting device has been carried out. The intake concept is based on the use of an external compression ramp in the form of a caret wavefider. In a design flow regime, this ramp generates a plane oblique shock wave lying on its swept leading-edges and a subsequent two-dimensional isentropic compression wave. The intake starting is provided with a device special flaps of which in the throat are opened slightly in order to enlarge it and to form here both longitudinal and cross slots for air bleed and boundary-layer diversion. The same device can also serve for regulation of intake operating regimes depending on the flight velocity. The experimental model intake with a design Mach number MD) - 2 has been developed and tested at flow Mach numbers 1·5, 1·75, and 2·0 in the T-313 supersonic wind tunnel based at ITAM. Flow patterns about the intake have been observed, and flow parameters characterising the intake performance have been measured. The experimental results obtained have confirmed the workability of the starting/regulating device; the data on the efficiency of the model intake have demonstrated its good characteristics in terms of the total pressure recovery. [ABSTRACT FROM AUTHOR]

The results of a numerical investigation of supersonic off-design flow past waveriders at the freestream Mach numbers M∞ = 4 and 8 are presented. Flow regimes with M∞ both greater and smaller than the design value Md are analyzed. Configurations based on the flows behind plane shocks and described by power-law functions are considered. The results are obtained by the finite-volume solution of the Euler equations using higher-order TVD Runge-Kutta schemes. [ABSTRACT FROM AUTHOR]

The results of numerical and experimental studies of a model configuration of a hypersonic vehicle realizing the principle of compression convergence of the entire jet captured by an air-intake from the oncoming stream are presented. The external flow past a convergent air-intake integrated with a transversely-concave nose compression surface is investigated over the freestream Mach number range from 2 to 6. The salient features of the flow pattern near air-intakes with sidewalls of different length are established and the influence of a limitation on the lateral flow near the external compression wedge on the flow rate characteristics of the air-intake is established. The numerical calculations are performed within the framework of the inviscid gas model using higher-order TVD schemes; the calculated and experimental results are compared. [ABSTRACT FROM AUTHOR]