# TsAGI Science Journal, Issue 2, Year 2015

**NUMERICAL INVESTIGATION OF A STEADY VISCOUS INCOMPRESSIBLE FLUID FLOW OVER A FAMILY OF ELLIPTIC CYLINDERS**

*G. L. Korolev*

*The numerical solution to the problem of the symmetric steady flow over an elliptic cylinder is obtained. The major axis of the cylinder is set perpendicularly to the uniform incident flow of a viscous incompressible fluid for the Reynolds numbers up to *Re *= 900. The dependences of the separation zone sizes and the drag coefficient on the Reynolds number and a relative body thickness are studied.*

**KEY WORDS: ***Navier—Stokes equations, elliptic cylinder, separation, numerical method, incompressible fluid, steady solution*

**AXISYMMETRIC BODY AT AN ANGLE OF ATTACK IN HYPERSONIC FLOW: HEAT PROBLEM**

*V. A. Bashkin, I. V. Egorov, & I. V. Ezhov*

*Hypersonic flow (*M_{∞ }= 7.9) over a Martian probe model of the Martian Sample Return Orbiter type with an isothermal surface (in which the temperature factor is T_{w}_{0} *= 0.375) at an angle of attack (0 ≤ α ≤ *40*°) is calculated based on the unsteady three-dimensional Navier—Stokes equations. The temperature field structure and the thermal characteristics in the flow symmetry plane and their evolution with the angle of attack are investigated.*

**KEY WORDS: ***axisymmetric body, angle of attack, hypersonic flow, Navier—Stokes equations, numerical simulation, flow symmetry plane, aerodynamic heating*

**CHARACTERISTICS INVESTIGATION OF 2D AND 3D INTAKES OF SUPERSONIC PASSENGER AIRCRAFT**

*V. A. Vinogradov,** V. A. Stepanov, & Ya. A. Melnikov*

*Two intake configurations, three-dimensional (3D) and planar [two-dimensional (2D)], for a cruising (*M = *1.6) supersonic passenger aircraft (SPA) are considered. The intake includes a supersonic inlet and large-curvature subsonic diffuser (SD). Preliminary incoming airflow deceleration and boundary-layer control (BLC) in the 3D inlet case are carried out with the help of a swept wedge placed in front of the inlet due to a transverse pressure gradient being formed. In the 2D inlet case, boundary-layer bleed is achieved by installing the inlet on the wedge at some height from the surface. Further deceleration is realized in a system of shocks with terminal normal one. The results of the flowing inlet with the SD and boundary-layer bleed in the throat in throttling modes at *M* _{∞ }*=

0.8–1.8 were obtained with the help of numerical simulations based on the solution of the 3D Navier—Stokes equations. Flow stability was confirmed when the pressure in the SD exit section was raised 1–4 times compared with the external free-stream pressure. It was found that the total pressure recovery coefficient σ

*grows up to*σ =

*M*0.89–0.97 when using BLC. The drag estimations of the considered inlets—including the additive, wave drag components, and also the drag of the bleeding wedge— show that the summary drag of the 3D inlet is 1.5–3.5 times less than that of the 2D inlet in the

*=*

_{∞ }0.8–1.6 range when the efficiency of the flow deceleration in the inlet is the same.

**KEY WORDS: ***supersonic passenger aircraft, 3D intake, flow rate, coefficient of pressure, drag coefficient*

**PARAMETRIC EXPERIMENTAL INVESTIGATIONS ON THE CROSS CORRELATION OF THE PRESSURE FLUCTUATIONS GENERATED BY A CASCADE OF “FORWARD-FACING STEP/BACKWARD-FACING STEP” COMBINATIONS**

*A. Yu. Golubev** & B. M. Efimtsov*

*The results of laboratory parametric experimental investigations of the pressure fluctuation fields generated by the flow over a cascade of “forward-facing step/backward-facing step” combinations are presented. This cascade of steps models the pressure fluctuation fields generated by the flow over aircraft fuselage windows. The module |*φ|* of the normalized cross spectrum of the pressure fluctuations for the points in the cascade of forward-facing step/backward-facing step combinations significantly exceeds the normalized cross-spectrum module at the same observation points in the case of an undisturbed boundary layer. Of all the configurations considered, both at the beginning and end of the cascade, the maximum cross correlation is revealed in the pressure fluctuation fields generated by configuration 4 (i.e., the backward-facing step/backward-facing step). The normalized cross spectra of the pressure fluctuations generated by the configurations in the cascade increase when moving from the beginning to the end of the cascade. The dimensionless step height in the cascade does not considerably affect the cross-spectrum module |*φ|* or its argument arg *φ* and dimensionless phase velocity U _{ph}/U. The absence of intermediate combinations of the forward-facing step/backward-facing step in the cascade has no significant effect on the cross correlation of the pressure fluctuation fields generated by the combinations situated downstream in the flow.*

**KEY WORDS: ***pressure fluctuations, forward-facing step/backward-facing step combination, cross correlation*

**MEASUREMENTS OF DEFORMATION OF THE PASSENGER PLANE WING IN FLIGHT BY THE VIDEOGRAMMETRY METHOD**

*V. P. Kulesh*

*The results of the deformation measurement of a passenger airplane wing in flight using the videogrammetry method are presented. Recording of the images of the left wing was performed manually through the window from a passenger seat above the wing using a common digital camera (Kodak Easy Share V1253, Rochester, New York, United States). The wing structure elements easily detected in the images were used as references, which were generally the corners of the mechanization elements. The references were mainly grouped in six sections. The measurement uncertainty was estimated by calculation; for linear displacements the uncertainty was from 6 mm in the neighboring section to 10 mm at the wing tip and for angular displacements the uncertainty was from 0.2 to 0.8°, respectively. The measurements were carried out for four regimes, namely, cruising flight, flight with deflected interceptors, flight with extended flaps, and while taxiing along the ground with a quenched speed. It was obtained that the maximal values of the bending elastic deformation at the wing tip varied in the range of 600–1200 mm in the vertical plane and in the range of —100 ± 150 mm in the horizontal plane. The total elastic deformation of the wing torsion achieved a value of 2–2.5°.*

**KEY WORDS: ***aerodynamic wing characteristics, bending deformation, torsion, geometric parameters, noncontact measurements, videogrammetry, flight tests*

**METHOD OF DETERMINATION OF RATIONAL DESIGN PARAMETERS USING THE FINITE-ELEMENT METHOD**

*E. V. Kasumov*

*The article discusses the technique of carrying out design calculations of wings made of composite materials. Reviewed are the calculation results of some rational parameters of thin-walled structures elements under the influence of aerodynamic forces.*

**KEY WORDS: ***numerical experiment, designing, strength, stress—strain analysis*

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