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The main purpose of this study was to determine the basic aerodynamic characteristics of the airliner Tu-154M at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°.

Wind tunnel tests of the Tu-154M aircraft model at the scale 1:20 were performed in a low-speed wind tunnel T-3 by using a six-component internal aerodynamic balance. Several model configurations were also investigated.

The results of the presented studies showed that at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°, the Tu-154M aircraft flap deflection affected the values of the drag and lift coefficients and generally had no major effect on the values of the side force and pitching moment coefficients.

The model vibration which was the result of flow separation at high angles of attack was the wind tunnel test limitation.

Studies of the airliner aerodynamic characteristics at the wide range of the overcritical angles of attack and sideslip angles allow assessment of the aircraft aerodynamic properties during possible unexpected situations when the passenger aircraft is found to have gone beyond the conventional flight envelope.

There are no social implications of this study to report.

The presented wind tunnel test results of the airliner aerodynamic characteristics at overcritical angles of attack and sideslip angles is an original contribution to the existing not-too-extensive database available in the literature.

The purpose of this paper is to present the result of simulations that were performed to assess the uncontrolled motion of the passenger aircraft following its wing tip was suddenly cut. Such a simulation can help to understand the tendencies of aircraft behaviour after wing tip cut, assess how fast this phenomenon is going on and estimate the values of angles of attack, sideslip and pitch angle basing on given aerodynamic characteristics. Also, answer the question if pilot can counteract high deviations from flight path initially planned during the final phase of approach to landing.

Simulation is based on the full non-linear equations of motion derived from generalised equations of change of momentum and moment of momentum of rigid body. Dynamic equations of motion in the so-called normal mode are solved in the so-called stability frame of reference.

It was found that asymmetric rolling moment must be compensated by essential increase of pitching moment. Moreover, it appeared that aircraft goes into high angles of attack and high pitch angle and, therefore, for reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs.

For accurate simulation, it is strongly recommended to perform to perform first the wind tunnel testing in the range of +20o ÷ 120o and use it in flight simulation.

The presented methodology is an original for numerical simulation of flight trajectory during the final phase of approach to landing in a hazardous state of flight. For reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs, whereas usually maximum angles of attack used in wind tunnel experiments for passenger aircraft are not higher than 25 degrees. The influence of limited range of experimental data on results of simulation is another value which can be adopted in the future investigations of hazardous states of flight.

The purpose of this study was to analyze the possibility of using combat aircraft including decommissioned as a platform for launching and carrying space rockets with satellites (nano and microsatellites). Thus, an airborne-launcher-to-space-system may be attractive to countries without ground-based space rocket launch sites.

For considered launch-to-orbit system configurations, simulations of space rocket effects on aerodynamic characteristics were performed using computational fluid dynamics (CFD ANSYS Fluent) methods. In addition, experimental studies were performed in a wind tunnel to verify the numerical simulations. Discrete models of the aircraft structure were developed for analysis using finite element method (FEM). The analysis of simulated structural properties of the models was carried out to test its stiffness and mass characteristics important for solving the static and dynamic problems of the structure. The validation analyses of aircraft models were based on mass distribution estimation and matching the stiffness properties of the individual airframe structural assemblies.

The results of numerical analyses and tunnel tests indicate that the influence of carrier rockets on the change of aerodynamic and strength characteristics of the airframe is rather negligible. The aircraft can be used as launching platforms for space rockets. Simulations have indicated that the aircraft will successfully perform a mission of taking away and launching a rocket of at least about 1,000 kg total weight with a 10 kg space payload included.

The combat aircraft can be used as launch platforms for space rockets, and the air/rocket set can become the equivalent of responsive space assets for countries with small space budgets.

The work presents original results obtained by the authors during a preliminary design of a low-cost satellite launch system consisting of a carrier aircraft and a space rocket orbiter. The possibility of using decommissioned combat aircraft as air-launch-to-orbit platforms was taken into consideration. In the absence of aircraft design documentation, reverse engineering methods and techniques were used to develop aircraft geometry and airframe strength structure. Use of CFD, FEM and simulation methods to evaluate system capabilities was demonstrated. Numerical results from CFD simulations were finally verified in experimental tests.

To reduce the heat load of a gas turbine blade, its surface is covered with an outer layer of ceramics with high thermal resistance. The purpose of this paper is the selection of ceramics with such a low heat conduction coefficient and thickness, so that the permissible metal temperature is not exceeded on the metal-ceramics interface due to the loss ofmechanical properties.

Therefore, for given temperature changes over time on the metal-ceramics interface, temperature changes over time on the inner side of the blade and the assumed initial temperature, the temperature change over time on the outer surface of the ceramics should be determined. The problem presented in this way is a Cauchy type problem. When analyzing the problem, it is taken into account that thermophysical properties of metal and ceramics may depend on temperature. Due to the thin layer of ceramics in relation to the wall thickness, the problem is considered in the area in the flat layer. Thus, a one-dimensional non-stationary heat flow is considered.

The range of stability of the Cauchy problem as a function of time step, thickness of ceramics and thermophysical properties of metal and ceramics are examined. The numerical computations also involved the influence of disturbances in the temperature on metal-ceramics interface on the solution to the inverse problem.

The computational model can be used to analyze the heat flow in gas turbine blades with thermal barrier.

A number of inverse problems of the type considered in the paper are presented in the literature. Inverse problems, especially those Cauchy-type, are ill-conditioned numerically, which means that a small change in the inputs may result in significant errors of the solution. In such a case, regularization of the inverse problem is needed. However, the Cauchy problem presented in the paper does not require regularization.

The purpose of this paper is to simulate with in-depth reconstruction of existing geometry a process of cooling of the aircraft engine in pusher configuration, which is more problematic than usually used, tractor configuration. Moreover, a complex thermal and fluid flow analysis is necessary to verify that an adequate cooling is ensured and that temperatures in the engine nacelle are maintained within the operating limits.

Methodology used in this research is based on computational fluid dynamics tools to model adequately the internal and the external flow, to find the state of cooling system and research the results of baffles modification inside the engine cover. Additionally, two types of the cover with different sizes of inlets and outlets are tested.

The results showed the influence of baffles modifications and changes in inlets and outlet sizes on the mass flow rate and temperature distributions inside the engine nacelle. The best configuration of air inlets and outlets was determined.

The method used in the research is the safest method in testing of such cases, provided the proper approach in modeling is taken. The collaboration of internal and external flow is crucial and should not be replaced with assumed flow rate through inlet and outlet area. The obtained results will help in future studies on cooling systems of engines in pusher configuration.

The work presents original results obtained by the authors during a complex fluid flow and heat transmission analysis and is a part of the design project of the OSA patrol aircraft.

The finite element model developed for a new-designed aircraft was used to solve some problems of structural dynamics. The key purpose of the task was to estimate the critical flutter velocities of the light airplane by performing numerical analysis with application of MSC Software.

Flutter analyses processed by Nastran require application of some complex aeroelastic model integrating two separate components – structural model and aerodynamic model. These sub-models are necessary for determining stiffness, mass and aerodynamic matrices, which are involved in the flutter equation. The aircraft structural model with its non-structural masses was developed in Patran. To determine the aerodynamic coefficient matrix, some simplified aerodynamic body-panel geometries were developed. The flutter equation was solved with the PK method.

The verified aircraft model was used to determine its normal modes in the range of 0-30 Hz. Then, some critical velocities of flutter were calculated within the range of operational velocities. As there is no certainty that the computed modes are in accordance with the natural ones, some parametric calculations are recommended. Modal frequencies depend on structural parameters that are quite difficult to identify. Adopting their values from the reasonable range, it is possible to assign the range of possible frequencies. The frequencies of rudder or elevator modes are dependent on their mass moments of inertia and rigidity of controls. The critical speeds of tail flutter were calculated for various combinations of stiffness or mass values.

The task described here is a preliminary calculational study of normal modes and flutter vibrations. It is necessary to prove the new airplane is free from flutter to fulfil the requirement considered in the type certification process.

The described approach takes into account the uncertainty of results caused by the indeterminacy of selected constructional parameters.

This study aims to create 6-degree of freedom (SDOF) for computational fluid dynamics (CFD) simulations of body movement, and to validate using the experimental data for empty tank separation from I-22 Iryda jet trainer. The procedure has an ability to be modified or extended, to simulate, for example, a sequential release from the joints.

A set of CFD simulations are calculated. Both the SDOF procedure and the CFD simulation settings are validated using the wind tunnel data available for the aircraft.

The simulation using designed procedure gives predictable results, but offers availability to be modified to represent external forces, i.e. from body interaction or control system without necessity to model the control surfaces.

The procedure could be used to model the separation of external stores and design the deployment of anti-radar chaff, flares or ejection seats.

The work presents original work, caused by insufficient abilities of original SDOF procedure in ANSYS code. Additional value is the ability of the procedure to be easily modified.

The paper describes the application of two different vibration measurement methods for the identification of natural modes of the miniature unmanned aerial vehicle (UAV). The purpose of this study is to determine resonant frequencies and modes of mini-airplane within the specified range of frequency values.

Special measuring equipment was used including both contact and non-contact techniques. The measuring systems on equipment of the Institute of Aviation Technology in the Faculty of Mechatronics, Armament and Aerospace of Military University of Technology (Warsaw, PL) were used to conduct measurements. In traditional ground vibration testing (GVT) methods a large number of sensors should be attached to the aircraft. The weight of sensors and cables is negligible in relation to the mass of the large aircraft. However, for small and lightweight unmanned aerial vehicles, this could bring a significant mass component in relation to the total mass of the tested object.

The real mini-UAV construction was used to investigate its resonant modes in the range of frequencies between 0 and 50 Hz. After receiving the output values it is possible to perform some flutter calculations within the range of operational velocities. As there is no certainty that the computed modes are in accordance with those natural ones some parametric calculations are recommended. Modal frequencies depend on structural parameters which are quite difficult to identify. Adopting their values from the reasonable range it is possible to assign the range of possible frequencies. The frequencies of rudder or elevator modes are dependent on their mass moments of inertia and rigidity of controls. The critical speeds of tail flutter were calculated for various combinations of stiffness or mass values.

In this paper, some specific techniques of performing the GVT test were presented. Two different measuring methods were applied, i.e. the contact method and the non-contact method. Using the dedicated apparatus in relation to the mini-airplane, properly prepared in terms of mass distribution, rudders deflection stiffness and proper support, some resonant characteristics can be determined. The contact measuring system consists of a multi-channel analyzer, piezoelectric accelerometers, electrodynamic exciters, amplifiers, impedance heads and a computer with the Test.Lab Software. As the non-contact method, a laser scanning vibrometer was used. The principle of its operation is based on the separation of the emitted laser beam. The returning beam reflected from a vibrating object is captured by the camera and compared to the reference beam. Dedicated software analyzes collected data and on the basis of it creates animations of structural vibrational shapes and spectral plots within the investigated frequency range.

The object used for research is the mini-UAV Rybitwa – composite mini-plane with a classic aerodynamic layout manufactured in Institute of Aviation Technology Military University of Technology. In the work, both measurement methods and some sample results were presented. Results referenced to dynamic properties of the mini-UAV can be applied in the future for its finite element model tuning, what would be useful for the needs of some parametric analyzes in case of some UAV modifications because of its structural or equipment modifications.