I. I. Mazhul

Numerical modeling of the conditions for realization of flow regimes in supersonic axisymmetric conical inlets of internal compression

The results of the numerical investigation of flow regimes in the axisymmetric inlets of internal compression at a supersonic flow around them are presented in the work. The main attention is paid to the determination of the ranges of the duct geometric convergence, in which a supersonic inflow in the inlet realizes. The investigation has been carried out at high supersonic freestream velocities corresponding to the Mach numbers М = 2−8 by the example of the frontal conical (funnel-shaped) inlets with the angles of the internal cone wall inclination δw = 7.5−15° under the variation of the relative area of the throat cross section. The flow structure alteration was studied and the critical relative areas of the inlet throat were determined, at which either there is no starting of the inlet in the process of flow steadying at the initial subsonic flow in it or a flow breakdown occurs in the process of flow steadying at an initial supersonic inflow. Numerical computations of the axisymmetric flow were done on the basis of the solution of the Navier–Stokes equations and the k-ω SST turbulence model.

Numerical investigation of supersonic flow breakdown at the inlet duct throttling

The work presents the results of investigating the process of supersonic flow deceleration in a duct of the two-dimensional inlet throttled by variation of the outlet cross-sectional area. An inlet with three external compression shock waves designed for the freestream Mach number Md = 7 was considered as an example for the investigation. A one-dimensional analysis of the conditions for realization of the supersonic flow deceleration regimes in the inlet duct with two throats — in the inlet entrance and at the inlet duct outlet, has been carried out. The parametric numerical computations of two-dimensional inviscid or turbulent flows in the inlet were performed with the use of the Euler and Navier—Stokes codes of the program package FLUENT. The critical conditions for the nonuniform flow in the outlet throat bringing to choking the inlet duct were determined.

Gasdynamic design of a two-dimensional supersonic inlet with the increased flow rate factor

A two-dimensional inlet of external compression with the increased flow rate factor at high supersonic velocities is constructed by the method of gasdynamic design. Its feature is that a flow with the initial oblique shock wave and the subsequent centered isentropic compression wave is formed over the external compression ramp of the inlet. These waves interact with one another so that a resulting stronger oblique shock wave and a velocity discontinuity arise in front of the entrance to the inlet internal duct. An example of an inlet configuration with the design flow regime corresponding to the Mach number Md = 7 is considered. The characteristics of this inlet were obtained in the range of the free-stream Mach numbers M = 4–7 with the use of a Navier—Stokes code for turbulent flow. They are compared with characteristics of an equivalent conventional shocked inlet. As computations have shown, the inlet with the isentropic compression wave has much higher values of flow rate factor φ at Mach numbers M < Md. So, for example, at M = 4 the value φ ≈ 0.72 for it is by 33 % higher in comparison with φ ≈ 0.54 for the equivalent shocked inlet.

Numerical investigation of flow over two sweepback wedges at M = 4 and 6

Results of numerical simulation are discussed: simulation was carried out for a configuration of two wedges with sweepback leading edges placed on a pre-compression ramp in a way that skewed surfaces of the wedges deflect the compressed flows in the opposite directions. It was demonstrated that this configuration produces a flow with irregular interaction in the plane of symmetry for shock waves produced by sweepback wedges. The shock waves formed by the skew wedges induce 3D boundary layer separations along sweepback leading edges of the wedges. Flows in the separation zones are directed toward the plane of symmetry of this configuration; they interact and produce in the central part a “swollen” zone of separation flow with a typical S-shaped profile of velocity. Simulation data was obtained for the free stream flow with Mach number M = 4 and 6 and based upon Navier—Stokes equations and k-ωSST turbulence model using FLUENT computation code. Inviscid flow described by Euler equations was considered as well.

Numerical simulation of the side pylon effect on the aerodynamic characteristics of models at their wind tunnel tests

The possible influence of fastening the models on a side pylon at their tests in wind tunnels on their aerodynamics at supersonic flow speeds has been considered. The physical problem of the pylon and the model interference has been investigated, and the estimates of the pylon influence on integral aerodynamic characteristics have been obtained. The numerical computations of the flow have been done using the averaged Navier-Stokes equations and the SST k-ω turbulence model in the range of freestream Mach numbers M = 2.5-5. As the investigation object the “classical” body of revolution of large aspect ratio is considered, which has a cruciform forward fins and six-blade tail stabilizers.

Supersonic flow over paired compression wedges of different sweep installed on the pre-compression surface

The results of the numerical modeling of the flow over configurations consisting of two adjacent wedges with swept leading edges located on a flat surface of the pre-compression ramp are presented. The flow around compression wedges with different sweeps of leading edges is considered: zero sweep (χ = 0), positive sweep (χ > 0), and negative sweep (χ < 0). Swept wedges deflect the flows compressed by them either to the sides opposite to the configuration symmetry plane (χ > 0) or, on contrary, one towards another (χ < 0). The computations have been done with solution of the averaged Navier—Stokes equations and with the use of the SST k-ω turbulence model at the freestream Mach number M = 6. The difference in flow structure is analyzed, which includes, in particular, quasi-conical three-dimensional separations of the turbulent boundary layer on the pre-compression surface, which are induced by shock waves generated by swept wedges. The characteristics of the side flow spread on compression wedges of different sweep are presented.

On the question of starting conditions for frontal axisymmetric inlets tested in hot-shot wind tunnels

The work presents the results of an analysis of starting conditions for some frontal axisymmetric inlets of internal compression tested at freestream Mach numbers М = 3−8.4 in the hot-shot wind tunnels based at Khristianovich Institute of Theoretical and Applied Mechanics (ITAM). The results of these inlets test are compared with the data of numerical computations of inviscid, laminar, and turbulent flows carried out by the pseudo-unsteady method. There were determined the inlet throat areas limiting either with regard to the inlet starting or with regard to providing the maximally possible degree of geometric compression of the inlet-captured supersonic airstream at its deceleration in the already started inlet. Reshaping of computed flow patterns in the inlets depending on the variation of the minimal cross section of the inlet internal duct is analyzed.

Numerical simulation of the influence of the control surfaces deflection on the aerodynamics of a slender axisymmetric configuration

A possible influence of the deflection of control surfaces on the aerodynamics of an axisymmetric slender configuration at supersonic flow speeds is considered. A classical configuration consisting from the fuselage in the form of a body of revolution and having cross frontal fins and six-blade trailing stabilizers is considered as the investigation object. The physical flow pattern at the deflection of horizontal fin consoles is investigated and the estimates are obtained for the influence of this deflection on both the characteristics of elements (the body and stabilizers) as well as on the integral aerodynamic characteristics of the entire configuration. Numerical computations of the flow have been done at the freestream Mach number М = 3 in the range of attack angles α = 0−10° and the angles of the control surfaces deflection δcs = ±5° on the basis of the averaged Navier−Stokes equations and the SST k-ω turbulence model.

Experimental characteristics of a supersonic three-dimensional air inlet with adjustable throat

In the present paper, we discuss results of an experimental study of performance characteristics of a 3D inlet with a flow-metering device at free-stream Mach numbers M = 1.75 and 2. The inlet was designed using gas-dynamic design methods. The initial external compression in the inlet is achieved using a V-shaped body called a waverider. The inlet is provided with a special device for its starting, also permitting regulation of the internal channel cross-sectional area in the throat region with the help of paired rotary panels, throat doors. The flow-rate and total-pressure loss characteristics in the throat of the model inlet were determined as functions of the degree of opening of the throat doors.

Flow turbulization in a pseudo-shock forming in an axisymmetric duct with a frontal inlet

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.