# Abstracts

## The Space Engineering and Technology magazine

¹ 3 (22), 2018

### Contents

#### INNOVATIVE TECHNOLOGIES IN AEROSPACE ACTIVITIES

**Chvanov V.K., Sudakov V.S., Levochkin P.S. **

*Modern liquid-propellant rocket engines made by NPO Energomash named after academician V.P. Glushko. Current status of programs and future prospects (to mark the 110th anniversary of academician V.P. Glushko)
*

The paper reviews the current status of development programs for high-power Liquid-Propellant Rocket Engines (LPRE) at NPO Energomash. The history of NPO Energomash goes back to May 1929, when, under the direction of V.P. Glushko a division was set up within the Gasdynamics Lab in Leningrad to develop rocket engines and rockets. For many years academician V.P. Glushko was the head of this company, he is the farther of our country’s liquid rocket propulsion engineering, outstanding scientist and designer, one of pioneers of rocket and space technology, general designer of Energia-Buran system. It was for the superheavy launch vehicle Energia that the world’s most powerful LPRE RD-170 were developed and became the basis for developing at NPO Energomash a large family of oxygen-kerosene LPRE. The paper presents a summary of the results of the program to develop and subsequently operate a family of oxygen-kerosene LPRE with afterburning of the oxidizing generator gas designed on the basis of LPRE RD-170/171 originally developed for launch vehicles Energia and Zenit. The paper provides key performance data for engines RD-170/171, RD-180, RD-191 and RD-181. It discusses the new challenge of developing the engine RD-171MV for LV Soyuz-5 and future prospects for further operation and upgrading of the family of oxygen-kerosene LPRE of NPO Energomash.

**Key words: ** Liquid-propellant rocket engines, NPO Energomash, academician V.P. Glushko, development of rocket engines.

**Bezyaev I.V., Stoyko S.F. **

*A review of projects for manned missions to Mars
*

The paper presents a historical survey of some of the projects for manned missions to Mars in US and in our country, describes design features of interplanetary vehicles and reconstructs their configurations. It also presents the current concepts, projects, plans, prospects for study and exploration of Mars. It proposes a strategy and objectives for the first manned missions to Mars, a possible configuration and features of a reusable interplanetary expedition complex (including an interplanetary orbital vehicle, power-generation and propulsion system, landing and ascent vehicle, crew delivery and return vehicle), that is capable to support manned exploration of interplanetary space from Venus to Mars, including the Moon, moons of Mars and asteroids, taking into account current safety requirements and a relatively low cost of the interplanetary complex. It proposes an alternative design of the reusable interplanetary complex which uses liquid-propellant propulsion system and aerocapture in planetary atmospheres.

**Key words: ** Manned mission to Mars, history of projects for manned interplanetary vehicles, interplanetary expedition vehicle, aerocapture in planetary atmospheres.

#### FLYING VEHICLES AERODYNAMICS AND HEAT EXCHANGE PROCESSES

**Salosina M.O.**

*Optimization of multi-layer heat shield of solar probe
*

The paper presents a methodology for optimal design of multilayer thermal protection coating, ensuring required operational temperature on the boundaries of layers and minimum of total mass of system. The optimization problem is solved using a computational scheme, which combines the projected Lagrangian method with the quadratic subproblem and the penalty function method. The penalty function method is characterized by a large region of convergence and provides a good initial estimate of the optimal parameters’ values for the projected Lagrangian method with excellent local convergence properties. To illustrate the implementation of the developed algorithm and the corresponding software, the paper considers the problem of choosing of the optimal layer thicknesses for the multilayer heat shield of a solar probe, exposed to extreme heat loads during operation. The main properties of the potential high-temperature materials for solar probe shield are discussed, and the results of thickness calculation for the two-layer heat shield are presented, taking into account the dependence of the thermal properties of the layers' materials on the temperature and radiation from the heated surface of the shield.

**Key words: ** solar probe, thermal protection, multi-layer heat shield, optimal design.

#### FLYING VEHICLES ENGINEERING, DESIGN AND MANUFACTURING

The paper discusses a concept for optimizing design parameters of electromagnetic shields to protect sensitive equipment onboard a spacecraft against powerful electromagnetic interference in the near zone of radiation. The objective of the optimization is to select geometric parameters of the shield and construction materials in order to achieve the required shielding efficiency at lowest economic cost. The paper describes a method of simulating sources of electromagnetic interference in the form of equivalent radiators and their emitted fields. This method is based on using measurements of field strength from actual onboard hardware, which significantly improves the accuracy of the models. The paper presents a method for calculating the electrical component of the electromagnetic field generated by the source in the near zone. It discusses a method for calculating electrical field penetrating through the shield. It also presents an algorithm for obtaining initial data needed to optimize shield parameters for powerful sources of electromagnetic interference onboard a spacecraft.

**Key words: ** spacecraft, electromagnetic compatibility, shielding, simulation, electric field, field strength.

#### FLYING VEHICLES STRENGTH AND THERMAL ENVIRONMENTS

**Avershyeva À.V., Bobylev S.S., Mezhin V.S. **

*Verification experience of spacecraft load-bearing structure and equipment units loading parameters results
*

One of the most important issues which should be resolved during the space vehicles development phase is dynamic coupled loads analysis. The results of such analysis for spacecrafts is the determination of loading parameters (including internal forces, moments and accelerations, as well) for load-bearing body structure and separate secondary structure items, such as solar arrays antennae reflectors, fuel tanks and others. The RSC Energia’s experience under loading parameter determination results is generalized in this article. The object is to be investigated was conditionally divided by 7 parts, (named as partial units) under finite-element models (FEM) development. The partial FEM was developed for each mentioned item. Partial FEM’s verification was implemented by modal survey or/and vibrostrength tests (on the frequency response functions determination phase). The verified partial FEMs were assembled into “generalized” FEM of the Spacecraft by using «synthesis of partial dynamic parameters» methodology. The RSC Energia’s experience regarding determination of the loads for principal interfaces of spacecraft body and limit values of load factors harmonic spectra for attachment zones of the SC secondary structure items verification by using coupled loads analysis results is presented in the article.

**Key words: ** spacecraft, launch vehicle, coupled loads analysis, finite-element mathematic model, modal survey test, vibrostrength test, loads verification, loading parameters, load-bearing structure, synthesis of dynamic characteristics.

**Topilskaya S.V., Borodulin D.S., Kornyukhin A.V. **

*Making a compact gyroscopic angular rate vector meter resistant to mechanical forces
*

The paper discusses a Compact Gyroscopic Angular Rate Vector Meter (CGARVM) developed and built by Kuznetsov research institute of applied mechanics based on a dynamically adjustable gyroscope with a rotor gas-lubricated spin-axis bearing system.. The paper provides general technical data on the gyroscopic device (dimensions and mass parameters, measurement accuracy, the number of measuring axes, operational life, power consumption, etc.). The paper presents advantages and disadvantages of using the selected gyroscope (dynamically adjustable) as the sensing element. The advantages of this choice include the mean accuracy of measurements of this device while it has small mass and dimensions and long operational life. A disadvantage is the need to use special systems within gyroscopic devices (of the CGARVM type) designed to protect sensing elements against external mechanical loads, occurring when a spacecraft is being put into orbit by a launch vehicle. The paper sets forth key principles of assuring resistance to such loads. It provides a universal analytical model of the shock absorption system of the instrument. It presents theoretical simulation results and actual results of mechanical field tests of the CGARVM device. The practical result of the work consists in the development of a multi-purpose analytical model of the shock absorption system for CGARVM based on theoretical studies and results of practical development. The universal analytical model makes it possible to select design parameters of the instrument shock absorption system at the design stage without having to rebuild its prototypes.

**Key words: ** gyroscopic meter, instrument resistance, mechanical loads.

#### FLYING VEHICLES THERMAL, ELECTRIC PROPULSION ENGINES AND POWER GENERATING SYSTEMS

**Akhmedov M.R., Bideev A.G., Makarova E.Yu., Sazonov V.V., Khamits I.I. **

*Comparative analysis of calculated and experimentally measured output capacity of the orbital space vehicle solar batteries on the example of the Service Module of the International Space Station Russian Segment
*

The comparative analysis of calculated and experimentally measured output capacity of the service module solar batteries on the Russian Segment of the International Space Station (ISS) is presented. Special software is used for calculation; it considers the effect of partial shading of solar batteries upon the ISS structural units. The algorithm specifies shadowed cells definition by means of ray tracing and the electric current calculation based on the voltage balance equation. Input data includes the ISS three-dimensional computer model, orbit parameters, orientation parameters of the station and its movable units, current-voltage characteristics of photovoltaic cells and blocking diodes. The analysis assumed comparing of diagrams of calculated and measured electric current of the service module solar batteries; telemetric sensors are the source of experimental data. The analysis is executed for instantaneous and integrated values of current. The following factors of calculation accuracy are determined in the research resume: the necessity of simplifying of a geometrical model, the light reflection by the Earth and inadequate prediction of orientation of some ISS movable units (solar batteries and radiators of the ISS American Segment) cause of its depending upon the current need of electric power. The calculation accuracy is estimated, recommendations for software users are given and ways for its improvement are offered.

**Key words: ** space vehicle solar cell calculation, ISS 3D-model, current-voltage characteristics of solar cells, telemetry data from Russian Segment of ISS.

**Årmilov V.À., Kazankin F.À., Potabachny L.À., Emlin R.V., Morozov P.À. **

*Investigation of inluence of magnetic ield on thrust of high-voltage nanosecond pulsed plasma thruster
*

The paper presents thrust measurement results for a mockup of pulsed plasma thruster when the discharge area is exposed to an external magnetic field. A pulsed generator with 1 J inductive energy storage and semiconductor opening switch serves as a voltage source for the thruster mockup. The voltage pulse duration is 60 ns at amplitude of 250 kV. The applied magnetic field coming up to 8 mT in the discharge field is co-directional with the electric field of the discharger. The thrust pulse measurements show that with the magnetic field available the pulse increases from 1,7 to 2,3 µN•s. The estimates show that under these experimental conditions electrons in the discharge are fixed in a magnetic field. This results in increase of plasma temperature and ionization degree, which causes a pulse increase. Therefore, the magnetic field parallel to the discharge is a factor enabling to improve the performance of pulsed plasma thrusters.

**Key words: ** pulsed plasma thruster, nanosecond pulsed discharge, magnetic field.

#### ROBOTS, MECHATRONICS AND ROBOTIC SYSTEMS

**Yaskevich A.V. **

*Algorithms for contact parameters determination during math simulation of spacecraft docking and berthing
*

Spacecraft docking and berthing are realized using active and passive docking units, whose guiding surfaces provide decreasing of relative lateral and angular misalignments during approach. Complex shapes of these surfaces decrease the number of degrees of freedom and a relative motion volume of joining units during approach. Theses surfaces are not arbitrary, they are presented as a combination of geometrical elements described by first up to fourth order equations. This feature is taken into account when developing methods for describing surfaces and algorithms for calculation of contact parameters. Elements of guiding surfaces are presented by geometrical primitives or theirs ordered sets. An algorithm for contact possibility estimation and for calculation of contact parameters is developed for each pair of such primitives on the base of simple analytical formulas. Such surface elements as a truncated cone, cylinder, torus or theirs fragment are presented by ordered sets of geometrical primitives of a lower order, and a dichotomy algorithm, with these analytical formulas at each iteration step, is used for theirs contact possibility estimation. All above mentioned provide a real time simulation of docking and berthing.

**Key words: ** spacecraft, docking, berthing, contact interaction, contact parameters.

#### SYSTEMS ANALYSIS, CONTROL AND DATA PROCESSING

It is shown on the example of ExoMars that problem of combining mathematical thermal models of separate device in the unified instruments assembly model can be solved with help of nodal method, based on the graph algorithm. Despite the nodal models simplicity they have benefits which do nodal models extremely useful for space devices modeling.

This method has been applied to nesting of Russian ACS instrument thermal model in ExoMars European instruments assembly model. Comparison of calculation results and telemetric temperatures of ACS shows that nodal modeling errors can be lowered to the level at which the nodal models can be used as the main simulation instruments for space devices.

**Key words: ** mathematical thermal modeling, thermal mode of space devices, nodal thermal model, inverse thermal problem, ExoMars mission.

**Glebov I.V., Kogan I.L. **

*On the adequacy of the simulation model of functioning carbon dioxide recycling system
of habitable space objects
*

The paper concentrates on the issues confirming the adequacy of the simulation model of a newly developed carbon dioxide recycling system (ÑDRS). The justification of the criterion of adequacy of the simulation model is provided. To verify the adequacy of the simulation model, the results of computational experiments are compared with the data obtained during testing of design and technological model of ÑDRS. Criteria for testing hypotheses about the homogeneity of samples of different series of measurements are problems of statistical analysis. To evaluate the results of computational experiments, mean values of samples and dispersions (variances) were compared with the results obtained during testing of the design and technological model of ÑDRS. A comparative analysis of the experimentally obtained values of the main indicator of the system function, i. e. a «water yield» with the results of computational experiments is presented

**Key words: ** simulation model, criterion of adequacy, carbon dioxide recycling system,
statistical analysis, indicators of applicability.

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**¹ 3 (22) july - september 2018**