Comprehensive data analysis of atmospheric and cyclonic activity based on worldwide meteorological and oceanology data as well as the comparison with tilts and strains precise measurements by far distanced instruments have been performed. The earlier proposed model of earthquakes triggering due to atmosphere, ocean and lithosphere interaction was confirmed. The interaction develops as successively arising hurricanes (typhoons) activity in form of spatial-temporal swings of the lower pressure areas over the tectonic plates. The process started 1–2 months before the 2023 Mw 7.8 Turkey earthquake and after some cyclones reduction, it resumed. It was at this time that a major seismic shock occurred. This study considers the cyclones interaction in the Indian Ocean, North Atlantic Ocean and Mediterranean Sea during December 2022–January 2023. Excitations of Indo-Australian, African, Eurasian and Arabian tectonic plates progressed as NW-SE spatial and temporal swings over seismogenic area and were accompanied by tilt-baric and strain-baric disturbances detected by instruments installed in Central and East Europe and Far East regions. Tilt-baric effects of 1.2 mas/hPa and strain-baric events were observed for the most intensive cyclones 2–7 weeks before the earthquake.

The paper presents the computational experiment results aimed at reducing the intensity level of closing shock in the sonic shock wave profile taking into account the propulsion system jets interaction. The significantly reducing possibility the sonic boom effect on the environment is shown.

Our previous investigations evidently show that synchronous observation of global atmosphere and lithosphere disturbances by means of 130-1,600 km spatially distributed precise instruments is the effective method of detecting precursors of large seismic events and other dangerous natural phenomena. This study expands the spatial range of our search up to 7,000-8,000 km and allows regional and global disturbances to be shared.

The comparison of tiltmeter and strainmeter data in the periods of the strongest earthquakes with tropical cyclone activities in the World Ocean during January–April 2014 is made. Main features of the observed co-seismic tilt and strain processes are consistent with the results obtained for the strongest events during 1997–2004. The time-frequency data analysis and the comparison of the analysis results with the anomalous geomagnetic and ionospheric activity come to an agreement with the observed phenomena. The obtained results have allowed the triggering mechanism of seismicity to be proposed. The process begins as spatial and temporal swings of the regions of tropical cyclone origins and the basins of their activity. The powerful cyclone development accompanies a wide range of earthquake precursory phenomena, including abnormal behavior of ultra-wideband (0.002 mHz–3 Hz) Earth's oscillations, which can be recorded at far distances up to 1000–10,000 km. The daily dissipation energy of the most powerful tropical cyclone (hurricane, typhoon) is estimated to have same order of magnitude as the energy released by an event of Mw 7–9, as well, atmospheric depressions are big enough to trigger a forthcoming strong earthquake. The triggering mechanism could be caused by quasi-static and time-dependent surface loading that produces vertical tension and shear deformations. This loading affects the seafloor and coastline where they fall close to the adjacent tectonic plate boundaries.

Results of a physical experiment aimed at measuring the profile of relative excess pressure on a control surface in the near zone of the disturbed region of a schematized model of supersonic passenger aircraft (SPA) are reported. Tests in the test section of the T-313 wind tunnel, aimed at identification of the optimum mounting of SPA-model suspension and ensuring measurements of the full profile of the disturbed-pressure wave involving the leading, intermediate, and closing shock waves, were carried out. Comparisons of calculated data with experimental results are presented. Using the revealed optimum model suspension, measurement results in good agreement with the results of numerical calculations are obtained. The numerical solution to the problem about the flow around the geometric model was obtained and the necessary measurements in the experiment were carried out at freestream Mach number M ∞ = 2.04 and angle of attack α = 4°.

Abstract The paper presents the computational experiment results aimed at studying the influence of the aircraft propulsion system on the sonic boom parameters. As a result, at large distances from the disturbances source at jet temperature of 2400° K and velocity of 1700 m/s at the nozzle exit, the trailing shock wave intensity in the sonic boom profile was significantly reduced by 50% compared to the profile generated by the aircraft without considering the jet influence.

This paper presents the results of numerical calculations and experiments in the form of distributions of disturbed pressures generated by the tandem configuration of a schematized model of a supersonic passenger airplane on the control surface in the disturbed region. The numerical solution of the problem on the flow over a geometrical model was carried out and the necessary measurements were made in experiments at a Mach number of the free stream M = 2.03 and an angle of attack α = 3.5o. It has been shown that the pressure distribution over the azimuth coordinate has a spatial character. Comparison of numerical calculations with the results of the experiment has shown their good agreement in the part of the profile where shock waves from the nose and from the forward and the rear wings are present. Because of the design features of the model–holder coupling, the trailing shock wave and the wake were not modeled in the experiment.

This paper is a review of the series of investigations carried out at the S. A. Khristianovich Institute of Theoretical and Applied Mechanics (ITAM, Siberian Branch of the Russian Academy of Sciences (RAS)) aimed at solving the sonic boom problem. It has been shown, with the example of flow over a body of revolution at Mach number M = 2, that a decrease in the flow temperature near the body leads to a change in the flow structure in the region of the formation of a hanging shock wave and a decrease (by 12%) in the sonic boom intensity on the ground. The results are presented of investigations of the influence of the relative position and geometrical shape of configuration elements of a supersonic passenger airplane on the sonic boom parameters at long distances from the disturbance source. The prospective aspect of the supersonic passenger airplane providing the minimum level of sonic boom in cruising flight has been determined. The scheme (IZS method) for measuring in the experiment disturbed pressure profiles at a given distance from the investigated model has been described.

The level of the sonic boom forming on the Earth surface which is negative for the environment is an important challenge related with the potential usage of the SCA (supersonic civil aircraft). Creation of the SCA configuration with the decreased level of the sonic boom involves the investigations of the influence of the aircraft elements and their position on the forming disturbed pressures in the near zone and their propagation over big distances from the source [1].

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.

The paper presents a description of the procedure used for determining the aerodynamic characteristics (forces and moments acting on a model of a flying vehicle) obtained from the results of pressure measurements on the surface of a model of a re-entry vehicle with operating retrofire brake rockets in the regime of hovering over a landing surface is given. The algorithm for constructing the interpolation polynomial over interpolation nodes in the radial and azimuthal directions using the assumption on the symmetry of pressure distribution over the surface is presented. The aerodynamic forces and moments at different tilts of the vehicle are obtained. It is shown that the aerodynamic force components acting on the vehicle in the regime of landing and caused by the action of the vertical velocity deceleration nozzle jets are negligibly small in comparison with the engine thrust.

The author gives results of parametric calculations of shock-boom levels in the case of flow with a free-stream Mach number of 2.03 past configurations of a supersonic aircraft. The calculations are aimed at investigating the influence of the relative position of basic elements and their geometric shape on the aerodynamic quality of the configuration and on the parameters of shock boom at great distances from the perturbation source. The geometric models of the configurations were formed by combining and joining component elements: the body, the front wing, and the rear tapered wing with root dogtooth extension. From an analysis of all the considered models of tandem configurations with account of the resolvability of shock waves in a perturbed profile compared to the monoplane configuration, the optimum configuration has been singled out that ensures a reduction of 24% in the intensity level of shock boom with an increase of 0.24% in its aerodynamic quality.

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.

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 δ cs = ±5° on the basis of the averaged Navier−Stokes equations and the SST k-ω turbulence model.

Results of an experimental study of the aerodynamic characteristics and flow structure in the vicinity of a re-entry vehicle (RV) with supersonic braking plumes impinging onto landing surface (LS) are reported. The experiments were carried out for two smallscale (1:15) RV models provided with pressure taps, with cold air being used as the test gas. The RV models differed from one another both in the number of braking plumes (16 and 8) and in the geometric arrangement of nozzles on the RV body. The influence of LS-reflected jets on the distribution of pressure over the RV surface was examined. Experimental data demonstrating effects due to the angle of inclination of the RV axis to the LS plane and due to the RV-to-LS distance are reported. Visualization of the flow structure due to supersonic braking plumes and visualization of limit streamlines on landing surface are presented. A brief description of designs of the used models, and also of the employed experimental equipment and procedure, is given. Numerical simulated data for the flow structure arising during the interaction of RV plumes with the landing surface are outlined.

The paper presents the results of numerical investigation of the effect of relative area of the leading wing and nose part shape on the forming middle zone of the sonic boom (minimization region) with the tandem-location of the wings of the fuselage. It is demonstrated that the modified power-law body used as the nose part provides the middle-zone length above the cruising flight altitude, as well as the distance between the head and intermediate shock waves, which permits significantly reducing the sonic boom affect.

Based on the analysis of various aspects of creating a supersonic transport aircraft of the second generation, the necessity of developing unconventional active methods of sonic boom level reduction is demonstrated. Surface cooling is shown to exert a significant effect on formation of the disturbed flow structure up to large distances from the body by an example of a supersonic flow around a body of revolution. A method of reducing the intensity of the intermediate shock wave and excess pressure momentum near the body is proposed. This method allows the length of the reduced (by 50%) sonic boom level to be increased and the bow shock wave intensity in the far zone to be reduced by 12%. A possibility of controlling the process of formation of wave structures, such as hanging pressure shocks arising near the aircraft surface, is demonstrated. The action of the cryogenic mechanism is explained.

Results of experimental and numerical investigations of the effect of gas injection through a permeable porous surface on the drag coefficient of a cone-cylinder body of revolution in a supersonic flow with the Mach number range M h = 3–6 are presented. It is demonstrated that gas injection through a porous nose cone with gas flow rates being 6–8% of the free-stream flow rate in the mid-section leads to a decrease in the drag coefficient approximately by 5–7%. The contributions of the decrease in the drag force acting on the model forebody and of the increase in the base pressure to the total drag reduction are approximately identical. Gas injection through a porous base surface with the flow rate approximately equal to 1% leads to a threefold increase in the base pressure and to a decrease in the drag coefficient. Gas injection through a porous base surface with the flow rate approximately equal to 5% gives rise to a supersonic flow zone in the base region.

The influence of the basic factors of cryogenic forcing on formation of the middle zone on the sonic boom and aerodynamic characteristics of the flying vehicle is studied by experimental and numerical methods. Experimental data obtained with alcohol or liquid nitrogen as an injected liquid are used for comparisons; as a result, the total effect of temperature and coolant evaporation can be determined. The influence of temperature is studied by means of numerical simulations of the cryogenic action of distributed injection of air. A comparison of numerical and experimental data reveals the effect of the coolant evaporation process on perturbed flow formation. It is demonstrated that evaporation of the coolant outgoing onto the vehicle surface should be intensified to increase the efficiency of cryogenic forcing (to decrease the coolant flow rate).