Annotation: The paper presents the experiment-calculated data on hydraulic parameters of ring intake device in different operation modes. The method of their calculation is proposed taking into account limit deviations of influencing factors. Satisfactory convergence of calculated and experiment data is shown.

Annotation: The scientific and methodological propositions for the designing single-stage guided missiles with the solid rocket motors for advanced multiple launch rocket systems are defined. The guided missiles of multiple launch rocket system are intended for delivering munitions to the given spatial point with required and specified kinematic motion parameters at the end of flight. The aim of the article is an analysis of the development trends of the guided missiles with the solid rocket motors for the multiple launch rocket systems, identifying the characteristics and requirements for the flight trajectories, design parameters, control programs, overall dimensions and mass characteristics, structural layout and aerodynamic schemes of missiles. The formalization of the complex task to optimize design parameters, trajectory parameters and motion control programs for the guided missiles capable of flying along the ballistic, aeroballistic or combined trajectories is given. The complex task belongs to a problem of the optimal control theory with limitations in form of equa lity, inequality and differential constraints. To simplify the problem, an approach to program forming is proposed for motion control in the form of polynomial that brings the problem of the optimal control theory to a simpler problem of nonlinear mathematical programming. When trajectory parameters were calculated the missile was regarded as a material point of variable mass and the combined equations for center-of-mass motion of the guided missile with projections on axes of the terrestrial reference system were used. The structure of the mathematical model was given along with the calculation sequence of the criterion function that was used for determination of the optimal parameters, programs and characteristics. The mathematical model of the guided missile provides adequate accuracy for design study to determine depending on the main design parameters: overall dimensions and mass characteristics of the guided missile in general and its structural comp onents and subsystems; power, thrust and consumption characteristics of the rocket motor; aerodynamic and ballistic characteristics of the guided missile. The developed methodology was tested by determining design and trajectory parameters, overall dimensions and mass characteristics, power and ballistic characteristics of two guided missiles with wings for advanced multiple launch rocket systems produced by the People’s Republic of China, using the limited amount of information available in the product catalog.

Annotation: The high demands for the flow rate and thrust characteristics specified for the modern solid-propellant rocket motors (SRM) under the strict mass and overall dimensions constraints require high level of mass fraction of propellant. And in the process of propellant grain combustion, erosive burning often takes place (increase of propellant burning rate depending on combustion products flow rate along the grain channel). This may play both negative (off-design increase of chamber pressure) and positive role (for example, increasing the launch thrust-to-weight ratio of the rocket). It is typical of the main SRMs of various rocket systems (multiple launch rocket systems, anti-aircraft guided missiles, tactical missiles, boosters). This paper proposes a methodology for calculating the internal ballistic characteristics of a solid propellant rocket motor under erosive burning, which is relatively time and resource consuming. The methodology is based on equidistant model of propellant grain combustion, where grain is divided lengthwise into a number of intervals. For any point of time during the engine operation, burning area and port area of each interval are calculated, taking into account erosive impact on each interval; total burning area is the sum of all intervals burning areas. Gas flow rate in each interval of the grain channel is calculated using gas-dynamic equations. The motor mass flow rate is a mass input sum of all the intervals; and the burning rate in each interval is estimated with proper erosion factor. The combustion chamber pressure had been calculated for four erosive burning models proposed by different authors. All the models showed convergence with the experimental SRM test data sufficient for engineering estimate (in particular, for maximum chamber pressure and combustion time). Selected as a result erosive burning model may be used to design new motors with solid propellants similar in chemical composition, and the model parameters are to be further customized using the test specimens.

Annotation: Despite stringent environmental requirements, modern launch vehicles/integrated launch vehicles (LV/ILV) burn toxic propellants such as NTO and UDMH. Typically, such propellants are used in the LV/ILV upper stages, where a small amount of propellant is contained; however, some LV/ILV still use such fuel in all sustainer propulsion stages. For launch vehicles containing toxic rocket propellants, flight accidents may result in the failed launch vehicle falling to the Earth’s surface, forming large zones of chemical damage to people (the zones may exceed blast and fire zones). This is typical for accidents occurring in the first stage flight segment, when an intact launch vehicle or its components (usually individual stages) with rocket propellants will reach the Earth’s surface. An explosion and fire following such an impact will most likely lead to a massive release of toxicant and contamination of the surface air. An accident during the flight segment of the LV/ILV first stage with toxic rocket propellants, equipped with a flight termination system that implements emergency engine shutdown in case of detection of an emergency situation, has been considered. To assess the risk of toxic damage to a person located at a certain point, it is necessary to mathematically describe the zone within which a potential impact of the failed LV/ILV will entail toxic damage to the person (the so-called zone of dangerous impact of the failed LV/ILV). The complexity of this lies in the need to take into account the characteristics of the atmosphere, primarily the wind. Using the zone of toxic damage to people during the fall of the failed launch vehicle, which is proposed to be represented by a combination of two figures: a semicircle and a half-ellipse, the corresponding zone of dangerous impact of the failed LV/ILV is constructed. Taking into account the difficulties of writing the analytical expressions for these figures during the transition to the launch coordinate system and further integration when identifying the risk, in practical calculations we propose to approximate the zone of dangerous impact of the failed LV/ILV using a polygon. This allows using a known procedure to identify risks. A generalization of the developed model for identifying the risk of toxic damage to people involves taking into account various types of critical failures that can lead to the fall of the failed LV/ILV, and blocking emergency engine shutdown during the initial flight phase. A zone dangerous for people was constructed using the proposed model for the case of the failure of the Dnepr launch vehicle, where the risks of toxic damage exceed the permissible level (10–6). The resulting danger zone significantly exceeds the danger zone caused by the damaging effect of the blast wave. Directions for further improvement of the model are shown, related to taking into account the real distribution of the toxicant in the atmosphere and a person’s exposure to a certain toxic dose.

Annotation: Specialized design office for liquid engines was established on July 22, 1958 to develop engines and propulsion systems, powered by liquid propellants to be installed on the combat missile systems and integrated launch vehicles (LV), developed by Yuzhnoye SDO. Moreover, liquid engines design office was assigned with manufacturing and testing of the main rocket engines, developed by NPO Energomash and to be installed on Yuzhnoye-developed launch vehicles. Over the past 66 years Yuzhnoye SDO has developed more than 40 liquid rocket engines (LRE) of various purpose, designed both to gas-generator cycle and to staged combustion cycle. Seventeen of them were commercially produced by Yuzhmash PA and installed on launch vehicles. Nowadays Yuzhnoye propulsion experts keep working on development of the advanced liquid rocket engines powered both by cryogenic and hypergolic propellants, which satisfy the majority of launch service market demands. Within the framework of extensive cooperation with foreign space companies, on a contract basis, Yuzhnoye propulsion experts are working on the design and development testing of the liquid rocket engines, as well as their components. The accumulated vast experience in the development of liquid rocket engines nowadays enables high scientific and technical level in the creation of up-to-date engines, demanded in the world market. Significant steps in this area have been made by the experts from the Yuzhnoye propulsion division and then subsequent manufacture and delivery by Yuzhmash PA of the engine intended for the European rocket Vega Stage 4; and designing the individual components for the engines with thrusts ranging from 500 kgf to 200 tf ordered by foreign customers. This article provides the review of current and scheduled activities of the Yuzhnoye SDO to develop the liquid rocket engines within the thrust ranges from ~ 40 kgf to ~ 500 tf.

Annotation: The main solid rocket motors of surface-to-air missiles and some short-range missiles have, as a rule, two operation modes – starting (augmented rating) and cruise (with decreased propellant consumption level). The methods to calculate intraballistic characteristics of such motors have a number of peculiarities, which set them apart from the methods of determining the characteristics of motors with constant propellant consumption level. The purpose of this article is to analyze such peculiarities, design methods, to find interrelation between the parameters of propellant consumption diagram, to determine the impact on the latter of motor design features and propellant characteristics. To achieve this goal, the method of analytical dependencies was developed. The equations obtained show that the required parameters of diagrams (including consumption-thrust characteristics difference between the starting and cruise modes) can be ensured due to varying either case diameter or propellant combustion rate or due to combined variation of these values. In practice, the cases are possible when for some reasons it does not seem possible to vary the case diameter or propellant combustion rate and the requirements to consumption diagram cannot be satisfied to the full extent. The task of motor developer in that case consists in determination of acceptable (alternative) propellant consumption diagrams that would be closest to required. The proposed method is based on calculation and construction of nomograms of dependencies of relative propellant consumption in cruse mode on relative time of starting leg at different propellant combustion rates and constant (required) case diameter and vice versa, at different values of case diameter and constant (available) propellant combustion rate. Using these nomograms, the rocket developer can determine the propellant consumption diagram acceptable for the rocket. In a number of cases, design limitations for separate main motor assemblies are imposed on consumption characteristic diagram that have an impact on its required parameters. The presented materials allow evaluating that impact and contain the proposals to remove it. The presented method allows quickly determining the conditions needed to fulfill required propellant combustion products consumption diagrams and in case of nonfulfillment of these conditions – allow presenting alternative options for selection of most acceptable one.

Annotation: The basic mathematical relations are considered to construct static characteristics of nozzle-shutter twostage piston pneumatic drive with the working medium – powder combustion products.

Annotation: One of the main types of arming of modern tactical and short-range missiles are cassette warheads based on nonguided blast-fragmentation submunitions that are widely used to kill group targets. The fullscale testing (flight tests) is an integral part of their creation. In the process of flight tests of tactical and shortrange missiles with cassette warheads, the safety issues are topical. Based on the capabilities of existing test ranges, it is planned to conduct such tests for the tactical and short-range missiles, being under development in Ukraine, in the Black sea water area where the cassette warheads with nonguided blastfragmentation submunitions (or their equivalents) will pose major hazard for ships. In the paper, two mathematical models are proposed to assess probability of killing (risk) a ship that may be present in the impact area of submunitions (submunitions equivalents) of cassette warhead. The first model was constructed in the assumption that the coverage area of cassette warhead and group dispersion of submunitions are known. Such model may be used to determine safety in the initial phases of cassette warheads development. The second model assumes that the configuration of cassette warhead and the scheme of submunitions firing were finalized, and accordingly, the nominal impact points of submunitions and their group and ind ividual dispersion are considered to be known. Practical application of both models requires the use of numerical procedures.

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: Being a final executive element of rocket control systems, a hydraulic actuator is at the same time the main source of various non-linear dependencies in rocket dynamic design whose availability dramatically com plicates theoretical analysis of their dynamics and control systems synthesis. The required accuracy and complexity of mathematical models of hydraulic servo mechanisms are different for different design phases of guided rockets. The paper deals with the simplest models of hydraulic servo actuators intended to calculate rocket controllability and to define requirements to response and power characteristics of the actuators. To calculate the rocket stability regions and to evaluate own stability of servo actuators, a linearized mathematical model of hydraulic servo actuator is used that takes into account the most important parameters having impact on stability of the servo actuator itself and on that of the rocket: hardness of working fluid, stiffness of elastic suspension of the actuator and control element, slope of mechanical characteristic of the actuator in the area of small control signals, which, as full mathematical model analysis showed, is conditioned only by dimensions of initial axial clearances of slide’s throats. The full mathematical model constructed based on accurate calculations of the balance of fluid flow rate through the slide’s throats allows, as early as at designing phase, determining the values of most important static and dynamic characteristics of a future hydraulic actuator, selecting optimal characteristics of slides based on specified degree of stability and response of servo actuator and conducting final modeling of rocket flight on the integrated control system test benches without using real actuators and loading stands. It is correct and universal for all phases of rockets and their control systems designing and testing. Using this mathematical model, the powerful actuators of a line of intercontinental ballistic missiles with swinging reentry vehicle and the main engines actuators of Zenit launch vehicle first stage were developed. The results of their testing separately and in rockets practically fully comply with the data of theoretical calculations.