Mikhail Konstantinov
Doctor of Technical Sciences
Professor,
Slow and steady wins the race!
1962
Graduated from Moscow Aviation Institute (MAI). Specialty “Aircraft”. Qualification of mechanical engineer for aircraft was awarded.
1960 - 1965
Studied and graduated from Lomonosov Moscow State University. Specialty “Mathematics”. Qualified as a mathematician.
1962 - 1967
Engineer, senior engineer, senior researcher at the Department of Aircraft Design of Moscow Aviation Institute.
1962 - 1966
Correspondence postgraduate course at Moscow Aviation Institute.
1966
Candidate thesis for the degree of Candidate of Technical Sciences in the specialty “Aircraft design and construction” was defended.
1967 - 1976
Assistant, senior lecturer, associate professor of the Department №601 “Aircraft Design” of Moscow Aviation Institute.
1972
Academic title of Associate Professor at the Department of “Aircraft Design” was awarded.
1975
Doctoral thesis for the degree of Doctor of Technical Sciences in the specialty 05.07.02 “Aircraft design and construction” was defended.
1976 - present
Professor of the Department of Space systems and rocket engineering at Moscow Aviation Institute.
1978
Academic title of Professor at the Department of “Aircraft Design” was awarded.
1996 - present
Senior researcher, leading researcher, chief researcher of the Research Institute of Applied Mechanics and Electrodynamics of MAI.
2003 - 2009
Chief specialist and leading specialist of the Federal State company “Lavochkin Research and Production Association”
2006 - present
Academician of International Academy of Astronautics (IAA).
2016 - 2017
Professor of the Department of “Space Flight Mechanics” of the Institute of Applied Technical and Economic Research and Expertise of RUDN University.
2018 - present
Professor of the Department of Mechanics and Mechatronics of the Institute of Space Technologies of RUDN University.
Teaching
Gives a course of lectures to RUDN University students: “Interplanetary Flight Path design”.
Science
- Analyzed the required perfection of a nuclear electric rocket propulsion system (specific mass of the installation) for the implementation of a manned Mars expedition. Analyzed this required perfection as a function of the time of the expedition and the mass of the space complex being put into the base earth orbit.
- Analyzed the influence of the characteristics of the power plant when using an electric propulsion system in the mercury research project.
- Analyzed the rational characteristics of the solar power plant of a spacecraft (SC) with an electric rocket propulsion system for the Solar research project. Considered the direct (without gravitational maneuvers) launch of the SC into a low heliocentric orbit with a large inclination to the plane of the solar equator.
- Showed that at the beginning of an energetically complex interplanetary flight, it is advisable to use a heliocentric Earth - to - Earth flight with a gravitational maneuver near the Earth. The trajectory of the heliocentric flight is implemented using an electric propulsion system. This maneuver allows us to significantly increase the hyperbolic excess speed and expands the transport capabilities of the SC. It was shown how the transport capabilities of space systems based on medium-sized (Soyuz-2) and heavy-class (Soyuz-2) launch vehicles expand when using such a flight scheme and a solar electric propulsion system with an electric power of 5 kW.
- Analyzed the change in the optimal thrust profile of an electric rocket propulsion system (the law of engine on - off) as a function of the characteristics of the transport system for space transport tasks.
- Developed a method for optimizing complex interplanetary flight schemes (flights with a chain of gravitational maneuvers) of the SC with an electric propulsion system. The method uses three stages. At the first stage, the problem of optimizing the flight path to the destination planet using gravitational maneuvers and additional speed pulses in deep space was analyzed. Used the method of evolutionary strategy with adaptation of the covariance matrix. At the second stage, each of the heliocentric sections (planet-planet) of the route under consideration was optimized separately. At the third stage, the multipoint boundary value problem of end-to-end optimization was solved. The entire set of necessary optimality conditions for gravity maneuvers were met.
- Analyzed several schemes for launching the SC into heliocentric orbits to study the Sun (project “INTERHELIO-PROBE”). Showed that the use of an electric rocket propulsion system at the initial stage of a heliocentric flight and a system of gravitational maneuvers makes it possible to bring a sufficiently large mass to the final working orbit of the SC in a relatively short time (for example, 5 years). Analyzed a number of chains of gravitational maneuvers that ensure the launch of the SC into working orbits and identified flight patterns that can be recommended for use.
- Studied the problem of parrying trajectory disturbances that may occur during an interplanetary flight of the SC with an electric rocket propulsion system due to the temporary impossibility of regular use of the engine. Showed that an abnormal engine shutdown should be provided for when designing the interplanetary trajectory of the SC. Proposed a new approach, taking into account the need to parry the trajectory perturbation associated with the emergency shutdown of the electric propulsion system. Concluded that it is advisable to adjust the nominal trajectories to increase the maximum permissible time for an abnormal engine shutdown. It was shown that optimization of the characteristics of additional passive sections (their position on the trajectory and duration) leads to an increase in the maximum permissible time of emergency engine shutdown to a level that can satisfy the transport system designer.
Scientific interests
- Design and ballistic analysis of space transportation systems;
- Space flight mechanics of a spacecraft with low-thrust engines;
- Design of trajectories in the implementation of complex schemes of interorbital and interplanetary flights.
The possibility of implementing the solar probe project by launching a research satellite into the system of heliocentric orbits with a relatively small perihelion radius and a sufficiently large inclination to the solar equator (the inclination of the last heliocentric orbit to the plane of the solar equator should be at least 30°) is analyzed. A comparative design and ballistic analysis of the possibility of using chemical and electrojet propulsion systems (EPSs) when launching a spacecraft into the considered system of heliocentric orbits is conducted. The transport systems being analyzed assume the use of the Soyuz-2-1b launch vehicle and the chemical Fregat booster when launching the spacecraft from Earth. The propulsion systems of the spacecraft itself are different. In one case, a chemical propulsion system is used, in the other, a solar EPS based on one stationary plasma engine of the SPD-140 type. The time of launching the spacecraft into the last heliocentric orbit of the considered orbit system is limited to 5 years from above. It was shown that using EPSs can significantly increase the spacecraft mass in operational orbits (from 910 to 1600 kg).
The results of the design and ballistic analysis of a 2-year manned mission to Mars are presented. Dependences of the maximum permissible specific mass of the electrically powered propulsion system on the spacecraft (SC) mass in the initial earth orbit are obtained. It is shown that, for the analyzed set of characteristics of SC systems with the SC mass in the initial orbit of 200 t, to implement a Mars mission with a 2-year duration, the required perfection of the electrically powered propulsion system should be ideally high (the specific mass of the electrically powered propulsion system should be no more than 3.26 kg/kW). Increasing the mass in the initial orbit leads to a relaxation of the requirements for the perfection of a transport system. If the mass in the initial orbit is increased to 475 t, then the maximum possible specific mass of the SC electrically powered propulsion system grows to 11 kg/kW. The optimal electric power of the nuclear power plant with a change in the initial mass within the mentioned range (from 200 to 475 t) increases from 7.7 to 11.7 MW. The optimal specific impulse of the electric propulsion system falls from 9000 s (this value is accepted as the maximum permissible) to 6880 s. It is shown that, if the specific mass of the electrically powered propulsion system is 5, 7.5, and 10 kg/kW, then, to implement the mission, the SC mass in the initial orbit should be no less than 234.1, 305.4, and 415.5 t, respectively.
The possibility of the spacecraft insertion into the system of operational heliocentric orbits is analyzed. It is proposed to use a system of several operational heliocentric orbits. On each orbit the spacecraft makes one or more revolutions around the Sun. These orbits are characterized by a relatively small perihelion radius and relatively high inclination, which allows investigating the polar regions of the Sun. The spacecraft transition from one orbit to another is performed using an unpowered gravity assist maneuver near Venus and does not require the cruise propulsion operation. Each maneuver transfers the spacecraft into the sequence of operational heliocentric orbits. We have analyzed several systems of operational heliocentric orbits into which the spacecraft can be inserted by means of the considered transportation system with electric propulsion (EP). The mass of the spacecraft delivered to these systems of operational orbits is estimated.
The possibility of the spacecraft insertion into the system of operational heliocentric orbits is analyzed. It is proposed to use a system of several operational heliocentric orbits. On each orbit the spacecraft makes one or more revolutions around the Sun. These orbits are characterized by a relatively small perihelion radius and relatively high inclination, which allows investigating the polar regions of the Sun. The spacecraft transition from one orbit to another is performed using an unpowered gravity assist maneuver near Venus and does not require the cruise propulsion operation. Each maneuver transfers the spacecraft into the sequence of operational heliocentric orbits. We have analyzed several systems of operational heliocentric orbits into which the spacecraft can be inserted by means of the considered transportation system with electric propulsion (EP). The mass of the spacecraft delivered to these systems of operational orbits is estimated.
In the mid-90s of the past century Arnon Spitzer has offered a flight scheme comprising an elliptical synchronous equatorial orbit for the SC injection into geostationary orbit (GEO). The advantages of such scheme are related to the SC control simplicity during its transfer from the elliptical synchronous orbit into GEO. The main advantage of Spitzer's scheme is the possibility of observing spacecraft from Earth using a limited number of ground stations along the entire trajectory of the SC flight from an intermediate orbit into GEO. Some flight schemes, which preserve the main advantage of the Spitzer's scheme, are analyzed. The inclination and the eccentricity of the intermediate synchronous orbit are optimized. Several variants of the yaw angle control are analyzed. Numerical analysis was carried out for space transportation system based on the launch vehicle "Angara-A5", chemical upper stage “KVTK”, and electric propulsion system comprising four thrusters SPD-140 with the specific impulse of 1700 s.
The complicated interplanetary flight path of the spacecraft with electric propulsion that uses multiple gravitational maneuvers is examined. The developed method bases on the using of solution of the auxiliary problem for SC with the chemical propulsion system. The results of solution of this auxiliary problem offered to use to identify rational flight paths of the spacecraft with electric propulsion and to optimize these flight paths. The auxiliary problem of the analyzed particular flight path is formulated as a problem of mathematical programming. A large number of equalities and inequalities constraints must be satisfied in solving this problem. The complicated flight path of the spacecraft to Jupiter is discussed as a typical example. The Earth, Venus and Mars are considered as the intermediate planets where the gravitational maneuvers will be accomplished.
The article provides the analysis of the ballistic capabilities to reject trajectory disturbances during interplanetary flight of a spacecraft with the electric propulsion system connected with temporary inability to ensure normal use of the electric propulsion on heliocentric segments of the flight trajectory. The main result of this analysis is the method for determination of a new nominal trajectory for spacecraft injection which allows having the longest contingency cutoff of the electric propulsion at any point of the injection trajectory. Numerical analysis is provided for one of the possible scenarios of spacecraft injection into operational heliocentric orbit for solar studies.
The paper is devoted to the design features of the prospective Russian INTERGELIO-ZOND spacecraft using, depending on the configuration version, an electric or chemical propulsion system as a sustainer. The scientific goal of the mission is the study of near-solar space from close distances (60–70 solar radii). The paper presents the description of several versions of the spacecraft options depending on the installed propulsion system, as well as the main characteristics of the flight profile depending on the engine type.