The modern aviation industry has many applications for honeycomb composite panels (HCP). The HCP-manufactured product line becomes more extended. This dictates a need for most effective and inexpensive methods for nondestructive inspection for HCP integrity that can ensure a high level of flight safety. One of promising methods for nondestructive inspection is the thermal technique. This paper presents and analyses the method feasibility of revealing the delamination of outer skin from an HCP filling. This task is performed by numerical simulation and the paper presents a pictorial test with a sample of HCP typical of aircraft engineering. Our results demonstrate how the thermal imaging tools can be used for the fact of skin delamination with regard for technical restrictions found in real experiments with a physical sample of HCP.

A method is proposed to improve the aerodynamic performance of small-scale aircraft systems. The method is based on fundamental gas-dynamic phenomena, such as local separation of the boundary layer accompanied by the formation of the so-called separation bubbles, separation of the turbulent boundary layer, and flow stall from the leading edge of the airfoil, which change the entire flow structure around the body. Publications where the relationship of these phenomena is established and those that describe a control method eliminating the adverse consequences of flow separation with the use of a wavy surface of the airfoil are reviewed. The method is simple in implementation and offers many prospects. The areas of its applicability are determined, and the criteria of optimizing the wavy surface of the airfoil for particular operation conditions are provided. Results of investigations are reported, which show that the use of a wavy surface of wings or blades of flying vehicles can improve their aerodynamic characteristics. Various structural elements of the boundary layer are noted, such as local separation bubbles or a global separation region, and the criteria of their emergence are given, which assist in verification of numerical experiments.

The paper describes how the acoustic impact on an impinging jet induces a maldistribution in the radial velocity field within the nearwall zone of spreading jet. The r.m.s. velocity pulsation level in the nearwall jet is lower than in the natural jet. A higher length of the laminar flow zone in a microjet was noted. The development of sinusoid-type instability facilitates a higher combustion efficiency for a propane-butane fuel mixture in the impinging microjet; this reduces the soot emission. The tests demonstrated that the zero acoustic impact makes the flame spectrum more yellow (including the flame from the nearwall zone). This testifies about deficiency of oxidizer (air) and soot presence in the combustion products. The combustion efficiency for diffusive flame in the impinging jet depends on the nozzle diameter and nozzle-target distance.

This article is devoted to an analytical study of the possibility of using a varioform sectional wing as a power and control surface of a promising future aircraft. The paper provides a theoretical assessment of the possibility of full control of the aircraft due to a variable wavy surface.

In the work details of experimental design of highly bearing laminarized airfoils GU-25-5(11)8 modified with smooth and wavy surfaces is considered. The results of measuring their aerodynamic characteristics at Reynolds numbers from 150 000 to 500 000 are given.

A new type of weight measurements in the aerodynamic experiment is presented to this work. It is based on recording the nonstationary characteristics of a wing oscillating in a low subsonic flow with the help of the movable tunnel balance. Experiments were carried out at T-324 ITAM wind tunnel at Reynolds number about 150,000 and frequency by 0.5-2 Hz.

Macro- and microjets facilities for generation of the round and plane subsonic jets are designed and fabricated. Automated measuring system (AMS - 2) for the spatial - temporal hot - wire anemometric visualization of jet flow field is designed and fabricated. Coordinate device and unit of the measurement, collecting, storage and processing of hot - wire anemometric information were integrated in the AMS. Coordinate device is intended for precision movement of the hot wire probe in jet flow field according to the computer program. At the same time accuracy of the hot - wire probe movement is 5 microns on all three coordinates (x, y, z). Unit of measurement, collecting, storage and processing of hot - wire anemometric information is intended for the hot - wire anemometric measurement of the jet flow field parameters (registration of the mean - U and fluctuation - u' characteristics of jet flow velocity), their accumulation and preservation in the computer memory, and also carries out their processing according to certain programms.

The present work is devoted to the study of a new type of bearing surface with the ability to optimize interaction with the medium, as well as controlling the movement in space.

System developed for automated carrying out of hot–wire anemometrical measurements in the aerophysical experiment is presented in the given work. Measurements of the mean (U) and fluctuation (u’) streamwise velocity components, their input in a computer and data processing it is carried out automatically with help of this system. Data processing of the experimental results are carried out by the special program written in MATLAB environment.

An investigation of aerodynamic performance for micro air vehicle (MAV) wing is performed. It is shown, what is a reason of significant wing performance difference between tests, which were carried out in different wind tunnels. New method for estimation of wind tunnel condition influence on experimental results is suggested.

Results of an experimental study of a subsonic flow past a straight-wing model installed at fixed angle of attack in hysteresis range of flow velocity are reported. It was shown that, at the same angle of attack and at one and the same flow velocity, flows of two types are possible on the upper surface of the model, leading-edge flow stall and re-attached flow with laminar-turbulent transition in the vicinity of the leading edge. It was found that the wing flow could be transformed from stalled to re-attached one with the help of a local influence applied at one point on the model inside the stall zone. Visualization data for both flows, hot-wire anemometer measurements of flow velocity, and measurements of the amplitude and frequency spectra of flow pulsations in both flows, are reported.

The possibility of controlling flow separation on an oblique airfoil using dielectric-barrier discharge has been experimentally studied. The experiments were performed at subsonic flow velocities in a broad range of the angle of attack. The results of measurements of the velocity and surface pressure fields and an analysis of the flow patterns show that the application of electric discharge allows the interval of the angles of attack for separation-free flow past the airfoil to be significantly increased. Various discharge regimes have been studied, including those with continuous activation by single voltage pulses with a frequency of 0.5–5 kHz and by pulse trains at a repetition rate of 1–100 Hz. The efficiency of the flow separation control has been studied as dependent on the electrical parameters, frequency characteristics, and position of the discharge relative to the flow separation line.

Flow past model wings is experimentally investigated in a subsonic wind tunnel at large angles of attack at which the laminar boundary layer separates near the leading edge of the wing (flow stall). The object of the study was the flow structure within the separation zone. The carbon-oil visualization of surface streamlines used in the experiments showed that in the separation zone there exist one or more pairs of large-scale vortices rotating in the wing plane. Certain general properties of the vortex structures in the separation zone are found to exist, whereas the flow patterns may differ depending on the model aspect ratio, the yaw angle, and other factors.