Annotation: The article is devoted to optimization of a trajectory of the antiaircraft guided missile performed in design phase. The review of existing solutions on this issue confirmed the topicality of the problem. The analytical solution cannot be obtained, therefore, according to modern tendencies, optimization by numerical method of original development was performed. The basis of the method is two-level optimization which is carried out, in turn, by two different numerical methods and for two different criteria functions. At the top level, by method of random search and as a variant, by method of coordinate descent, the search was carried out for a fixed set of intermediate for the specified flight range trajectory points which co-ordinates in aggregate provide the necessary optimum. At the bottom level, for each pair of consecutive intermediate points, the boundary problem of falling into distant point by one-dimensional optimization was solved. The coordinate descent method was used for search for the simplified flight program. As optimization criteria for top level, minimum flight time or maximum final speed, for bottom  terminal criterion were used. The control program selected the angle of attack  program. As a result, the optimum and suboptimum (additionally ensuring minimum calculation time) trajectories and flight programs to maximum range and different altitudes were obtained. The analysis of results showed practical proximity of trajectories of minimum flight time and maximum final speed.

Annotation: The process of creating modifications of aircraft in the transport category is a very relevant and widespread phenomenon in modern aircraft construction. A separate group of military transport aircraft has been distinguished in connection with the specific character of their mission: – the need to formulate the characteristics “cargo – range” for light, medium, operational tactical and strategic military transport aircraft, since it is precisely according to this characteristic that they are positioned by their purpose; –specific requirements are imposed on military transport aircraft cargo compartment not only with respect to its geometrical dimensions and usable volume, but also with respect to the possibility of simultaneous accommodation of military equipment and people, as well as the placement of a stretcher with t he wounded during their evacuation from the war zone; – the possibility of airborne landing of military equipment and paratroopers, which requires specific hatches and means of maintaining weight balance in flight; – the possibility of basing on poorly prepared sites with a runway length of less than 800 m in the short take-off and landing (STL) mode, especially for operational tactical military-technical vehicles, which significantly expands their use in combat zones; – the possibility of conversion into a civilian aircraft: for the delivery of goods to areas of the far north, when fighting fires, when evacuating victims from disaster zones, etc. The article shows that creation of modifications of expensive military transport aircraft is the main direction of their development. All leading aircraft manufacturing companies in the world use modification procedures as the way to most quickly meet constantly changing requirements for military transport aircraft. Along with the traditional methods of designing the modifications, the domestic school proposed a new methodology for determining the necessary parameters for “deep” modifications in wing geometry and propulsion system. The methodology is based on the use of three principles: – ensuring growth of carrying capacity and the required range of modifications of military transport aircraft of various purposes; – geometric re-arrangement of wing and system of carrying surfaces “wing + tail units” according to the criterion of minimum inductive resistance when lifting forces are equal to basic version; – coordination of modifications in wing with the required parameters of propulsion system as a condition for ensuring the required fuel efficiency.

Annotation: The main parameters are understood as: carrying capacity mг, range L and fuel efficiency qт, which largely determine the competitiveness of aircraft of this type, including military transport aircraft. The reason for the creation of modifications of transport category aircraft is the requirement for a constant increase in their flight performance by increasing the carrying capacity and range. Among the main goals of implementing such decisions there is an indispensable increase in the fuel efficiency of modifications, since the cost of fuel reaches 80 % of the cost of an airplane hour during operation. There are a number of models that make it possible to assess the influence of the weight and aerodynamic parameters of the airframe of the aircraft and the fuel performance of the power plant (specific engine consumption) on the integral indicator of the fuel efficiency of the modification at cruising mode and the average hourly fuel consumption at the certification stage, when all parameters of the airframe and engine are fixed and consideration of options is not possible. A new model is proposed for the stage of designing modifications, in which deep modification changes are made in the geometry of wing and in the power plant with various variants of their correlation and coordination. The parameters of new model: specific fuel efficiency – specific route productivity, in order to form the relative carrying capacity and relative range of action for the required specific fuel efficiency. An analysis of such dependencies showed: – with an increase in relative range L , fuel costs per flight also increase; – the adequacy of changes in route performance is observed only at L < 0.5. At L > 0.5 productivity is constantly decreasing, while the specific indicator of fuel consumption per unit of work increases exponentially; – if in the analysis we take into account the specific value of transport efficiency, that is, the characteristic “load – range” ( mп.н  f L ), it becomes obvious that the most favorable (from the point of view of fuel efficiency) are relative ranges of 0,3 < L < 0,5. In this range L , not only acceptable fuel efficiency values are realized, but also the maximum value of the route performance, that is the main parameters for which modifications are developed.

Annotation: High-precision guidance of supersonic flying vehicles maneuvering while descending in the atmosphere with high degree of thermal protection ablation is a well-known problem of space ballistics. The existing methods for calculating the ablation of thermal protection and the subsequent calculation of aerodynamic characteristics lead to scatter of the landing points of a flying vehicle reaching 5 km or more. The functional guidance method, in principle, allows achieving the required guidance accuracy (hundreds of meters), however, it requires a reserve of power of the controls at a level 50% to counter the influence of disturbing factors. The known terminal guidance method, which has recently become widespread, is based on a highly accurate prediction of motion parameters and, in this regard, has little promise. The method has been described in the article that allows 15-20-fold reducing the flight range scatters caused by lack of knowledge (including due to coating ablation) of its current aerodynamic characteristics and ensuring that the accumulated lateral deviation is counteracted in the limit to 1-1.5 km. The method is applicable to the flying vehicles with weight asymmetry (“transverse” displacement of the center of mass), performing maneuvering under conditions of aerodynamic balancing. The method is based on the solution to increase the accuracy of hits by spinning the shells around longitudinal axis. It is proposed that when a flying vehicle moves in the dive mode by means of the onboard CVC, it is regular (at intervals) to calculate its flight path in the (conditionally) autorotation mode. Based on the results of processing single calculations, the corresponding flight ranges of a flying vehicle and the lateral displacement of the touchdown points are determined, the point in time is predicted at which the flight range of the flying vehicle is equal to the specified one and the average lateral deviation is determined. At this moment the angular movement of the flying vehicle is transferred to the autorotation mode. Counteraction of the lateral displacement is introduced by adjusting the half-periods of flying vehicle movement along the angle of the precession. An example of pointing a flying vehicle at a given range, and bringing it to the touchdown point, shifted to the right relative to the original flight path by 1 km. The error of the terminal guidance of a maneuvering while reducing the aircraft using the proposed guidance method is determined.

Annotation: Basic dynamic properties of the reentry part of the aircraft-type first stage were examined when turbojet engine is used in the recovery phase. Such configuration can be of interest because turbojets have considerably smaller rate of flow in comparison to rocket engines. Moreover, they are launched in the lower stratosphere or in the troposphere so that there is no need to place oxidizer supply on board. This recovery plan differs from an alternative rocket recovery system and, from our point of view, provides more efficient usage of the fuel stores because it doesn’t require the main propulsion to be started in the recovery phase. Besides the analysis of qualitative characteristics of the descend phase for this stage, the efficiency of a wing with moderate values of maximum aerodynamic characteristics and a turbojet was studied. In this case three ways for stage recovery were investigated. The first one implied unguided descend with zero angle of attack assuming that the stage is statically stable. This descend trajectory was considered as standard and was used to evaluate the efficiency of the wing and turbojet with relatively small propulsion. The second and the third design cases offered the gliding guided descend with turbojet being started only in the lower stratosphere. The last two cases used the same program for the angle of attack. The possibility to ensure permissible overload values at the critical points of the descend trajectory and acceptable values of kinematic characteristics at the earth surface tangency point are also of great interest. Thereby the program for the angle of attack was developed in a way that allowed kinematic characteristics on touchdown be as close as possible to the corresponding values, shown by civil and/or military-transport heavy aircraft. Simulation was conducted on Microsoft Visual Studio 2010.

Annotation: The paper considers the project of prospective integrated gas purification device for large-sized LRE test stand. The prognostic mathematical models are presented for evaluation of ecological indices of the integrated gas purification equipment.

Annotation: An alternative method of kerosene dehydration is proposed, which is based on application of cyclic technology of supersaturation decrease using dry nitrogen. A comparison of nitrogen and time specific consumption in dehydration operations is done and recommendations are given for their use in the cosmodromes’ launch complexes fuel storage and preparation facilities.

Annotation: Mission control of the orbital space plane is one of the actual and complicated applied problems of the theory of mobile objects control. Dynamic configuration of this plane as an object of control is described by the system of non-linear differential equations of higher order. Research of stability of such system is a difficult problem. However, thanks to known theorems of Lyapunov, often stability of the real system can be estimated by the roots of the characteristic equation of the linearized system. Thereupon the stability analysis in the linear setting is the necessary link in the process of orbital space plane control system development. Among the methods of synthesis of the automatic control linear systems developed to date one can emphasize the trend, which has become widely-spread in the engineering area. According to this trend the issues of synthesis of the dynamic regulator, observability and controllability for the orbital space plane are considered. Procedure of selection of the dynamic regulator parameters at the early phase of development of the control system for the orbital space plane motion about the center of mass is suggested. Observability and controllability of the orbital space plane are considered. It is shown that the considered control system of the orbital space plane is observable and controllable, i.e. it is possible to develop the stable dynamic regulator, which provides the required speed and accuracy of the angular position of the orbital space plane during the orbital flight. Factors selection procedure is offered for the factors being the part of the control laws for the control system actuators.

Annotation: This paper suggests method for analysis of the dynamics of the aircraft with weight asymmetry (transverse displacement of the center of mass) maneuvering in the atmosphere under the impact of the short-time alternating moment of engine thrust, spread out over a period. The engines are installed on the bottom of the aircraft at the maximum distance from its longitudinal axis. Angular motion with nominal and perturbed performances of the aircraft and flight conditions has been consistently considered. Before maneuvering, the aircraft is set at the trimming angle of attack, determined by the magnitude of transverse displacement of the center of mass and aerodynamic characteristics. The direction of the aircraft maneuvering in the atmosphere depends on the acting moments of forces and time diversity of the engine firings to speed up and shutdown the angular motion. In the absence of disturbances, the angular motion of the aircraft shows in part signs of regular precession (almost constant precession velocity and nutation angle) and autorotation (close to zero self-rotation angle). Under the influence of disturbances, the spread of the aircraft angular motion parameters increases, mainly at the angle of precession, which characterizes changes in the direction of maneuvering. Composition of disturbances includes the spread of the aircraft technical characteristics (position of the center of mass, moments of inertia, aerodynamic coefficients, velocity head, etc.), errors associated with the operation of the engines (thrust spread, time of ignition and shutdown, angular alignment of their longitudinal axes). Terminal control was introduced to realize the given final state and to reduce the disturbances impact on the maneuvering parameters based on the registered deviations of the angular motion from the nominal one after the first shutdown of the attitude maneuver engine. Monte Carlo method (1000 variations of random realizations of the acting perturbations) confirmed the effectiveness of the proposed terminal control of the angular motion of the aircraft to provide the specified maneuvering parameters.

Annotation: The main task of the radiotransparent radome design is to develop a structure that reliably protects the aircraft components, sensitive instrumentation of the guidance system from incoming air flow and heating in all phases of operation, meeting all the requirements of minimum weight, allowable temperature under the radome, strength, centre of mass and radio characteristics. Development of the radome solves the complex task of coupling and optimization of the geometric aspects of design with physical-mechanical, radio and thermal properties of materials. The article dwells on the aspects of development of the radiotransparent radomes depending on the aircraft flight speed; basic requirements for the selection of outside perimeter configuration; structural materials, providing the required radio properties in combination with minimum mass and necessary strength, acceptable temperature inside the fairing. Yuzhnoye-developed radome for the up to 5 M aircraft was considered, consisting of ogive shell made of АФ-10ПО fiberglass, heat-resistant tip and structural ring made of aluminum alloy. Methods of thermal and strength analyses of the shell are considered, results of calculations confirming the fitness for work of the structure are presented. Options of configuration with ceramic shells and methods of their coupling with rings are presented for the aircraft that reaches up to 5 M speed. Basic principles of radiotransparent radomes development and aspects of ceramic shells coupling with structural rings are considered, as well as the results of strength analyses and thermal calculations that ensure the performance of the structure and equipment in the area under the radome.