An Analysis of the Efficiency of Electric Propulsion Engines for Maintaining a Low Orbit of Small Spacecraft

2020 ◽  
pp. 65-74
Author(s):  
V.V. Volotsuev ◽  
V.V. Salmin
Keyword(s):  
Electric Propulsion ◽  
Aerodynamic Drag ◽  
Specific Impulse ◽  
Active Movement ◽  
Engine Fuel ◽  
Low Thrust ◽  
Atmospheric Density ◽  
Small Spacecraft ◽  
Drag Forces ◽  
Electric Engine

This paper examines the problem of maintaining the plane parameters of the working orbit of a small spacecraft using an electric propulsion engine. In low working orbits, due to the Earth’s atmosphere, a spacecraft is subjected to aerodynamic drag forces, which results in a decrease in the radius of the orbit and a potential termination of the useful target functioning. The time parameters of the cyclogram for maintaining the working orbit of a small spacecraft with an electric low thrust engine are analyzed taking into account the variability of the atmospheric density. The cyclogram consists of sections of the passive and active movement under the action of the low thrust engine. For the satellite under study, suitable thrust parameters of the electric engine are selected, which allow the correction of the plane parameters of the low orbit. Using the characteristics of the thrust and specific impulse of the electric jet engine, fuel reserves for correction over a long period of time are calculated. The results of the analysis confirm the effectiveness of the electric propulsion engine in terms of fuel consumption for correction.

scholarly journals Analysis of characteristics of electric propulsion systems intended for carrying out maneuvers of maintenance of low Earth working orbit of small satellites

2021 ◽  
Vol 20 (3) ◽  
pp. 65-76
Author(s):  
V. V. Salmin ◽  
V. V. Volotsuev ◽  
A. V. Nikitin
Keyword(s):  
Electric Propulsion ◽  
Aerodynamic Drag ◽  
Specific Impulse ◽  
Operating Time ◽  
Propulsion System ◽  
Working Fluid ◽  
Propulsion Systems ◽  
Small Satellites ◽  
Small Spacecraft ◽  
Design Characteristics

An analysis of the mass of the working fluid and motor operating time of electric propulsion systems applied as a part of small spacecraft to carry out maneuvers of maintenance of the low Earth working orbit is carried out. The analysis is carried out for the small spacecraft with the weight in the range from 300 to 1000 kg functioning in working orbits with the height in the range from 400 to 600 km. When carrying out the analysis the values of the specific impulse of the propulsion system in the range from 800 to 1600 sec were accepted. Procedural guidelines for assessing the value of the required characteristic speed depending on the aerodynamic drag force, as well as for assessing the value of mass of the working fluid with account for the value of the specific impulse and defining the motor operating time of the propulsion system depending on the exhaust speed of the working fluid were used. The results of calculations given in the article show that the mass of the working fluid and the motor operating time vary depending on the height of the orbit and the mass of the small spacecraft and allow making quick preliminary assessment of the main design characteristics of the electric propulsion engines used to carry out maneuvers of maintenance of the low Earth working orbit of small spacecraft with different weight dimension characteristics during the prescribed term of active existence.

scholarly journals Minimum-Fuel Planar Earth-to-Mars Low-Thrust Trajectories Using Bang-Bang Control

2021 ◽  
Author(s):  
Daero Lee
Keyword(s):  
Optimal Control ◽  
Boundary Value Problems ◽  
Electric Propulsion ◽  
Direct Method ◽  
Boundary Value ◽  
Specific Impulse ◽  
Direct Methods ◽  
Optimization Methods ◽  
Low Thrust ◽  
Primer Vector Theory

Recent advance in electric propulsion systems have demonstrated that these engines can be used for for long-duration interplanetary voyages. Constant specific impulse engine described as a thrust-limited engine is an example of this type of engine, processing the ability to operate at a constant level of impulse. The determination of minimum-fuel, planar heliocentric Earth-to-Mars low-thrust trajectories of spacecraft using a constant specific impulse is discussed considering the first-order necessary conditions derived from Lawden’s primer vector theory. The minimum-fuel low-thrust Earth-to-Mars optimization problem is then solved in two-dimensional, heliocentric frame using both indirect and direct methods. In the indirect method, two-point-boundary-value problems are derived to solve boundary value problems for ordinary differential equations. In the direct method, a general-purpose optimal control software called GPOPS-II is adopted to solve these optimal control problems. Numerical examples using two different optimization methods are presented to demonstrate the characteristics of minimum-fuel planar low-thrust trajectories with on-off-on thrust sequences at three chosen flight times and available maximum powers. The results are useful for broad trajectory search in the preliminary phase of mission designs.

scholarly journals Increasing the specific impulse of the ion engine by zone engineering of the solid-state field cathode

2021 ◽  
pp. 41-50
Author(s):  
Nikolay Petrov ◽  
◽  
Tamara Antonova ◽  
◽  
Keyword(s):  
Electric Propulsion ◽  
Open Space ◽  
Technology Development ◽  
Rapid Development ◽  
Specific Impulse ◽  
Solid Material ◽  
Space Technology ◽  
Electric Motors ◽  
Low Thrust ◽  
Deep Space

With the rapid development of space technology, the scale of human space exploration is expanding significantly. However, the growing demand for deep space travel cannot be met with conventional chemical engines. Thus, the need for new mechanisms for providing jet thrust, including electric motors, becomes clear. Electric propulsion technology has significant advantages over traditional chemical engines in deep space flight due to its characteristics such as high specific impulse, small size, long service life. A negative feature of electric motors can be called low thrust, however, firstly, in open space this is insignificant and, secondly, the thrust of electric motors can be significantly increased, and for this, there are reserves available at the current level of technology development. Ways to increase the thrust of electric ion thrusters will be detailed and discussed in this work. The increase in the power of ion engines is limited to a large extent by the erosion of the control grids; the ion flow hits the surface of the solid material of the control grid electrode with energetic ions and gradually leads to the failure of this electrode. In this work, the authors will show that the use of field emission as a source of electron beams ionizing the working medium can solve the problem of erosion of control electrodes, due to which it will be possible to significantly increase the strength of the working fields for ion engines, which in turn will increase the specific impulse, efficiency, flow rate and power of the ion engine as a whole.

Conceptual design studies of an electrically propelled upper stage for deployment of a multi-plane orbital constellation of small spacecraft

2021 ◽  
pp. 76-87
Author(s):  
Aleksandr LEVANDOVICH ◽  
◽  
Dmitry MOSIN ◽  
Aleksandr TYUTYUKIN ◽  
Igor URTMINTSEV ◽  
...  
Keyword(s):  
Conceptual Design ◽  
Electric Propulsion ◽  
Aerodynamic Drag ◽  
Orbital Plane ◽  
Design Studies ◽  
Small Spacecraft ◽  
Solar Arrays ◽  
Upper Stage ◽  
The Difference ◽  
Orbital Constellation

The paper presents results of conceptual design studies to determine configuration of an electrically propelled upper (a space tug) with main engines arrays. It addresses the problem stage (EPUS) — a space transportation stage based on electric propulsion powered by solar of deploying a multi-plane orbital constellation of to in of small spacecraft (SSC) using an electrically propelled upper stage. It proposes change the SSC operational orbital planes based on the effect of precession rates between the parking and the working orbits eccentricity in the Earth gravitational field. Requirements owing have the difference to the effect been defined for the EPUS electrical propulsion system that take into account the need to operate it to offset the aerodynamic drag while waiting in the parking orbit for the SSC operational orbital plane to turn. It demonstrates the feasibility of employing four EPUS that use Stationary Plasma Thruster-type electric propulsion as their main engines and gallium arsenide solar arrays for deployment in a 600 km orbit in four planes an orbital constellation of 24 small spacecraft with a mass of ~250 kg each using one launch of a medium capacity launch vehicle of Soyuz-2.1b type.

COMPARATIVE ANALYSIS OF THE EFFICIENCY OF CONSTANT POWER THROTTLED ROCKET ENGINES FOR INTERORBITAL FLIGHTS TO GEOSTATIONAR

2021 ◽  
Vol 6 ◽  
pp. 66-77
Author(s):  
Igor Vasiliev ◽  
◽  
Boris Kiforenko ◽  
Yaroslav Tkachenko ◽  
◽  
...  
Keyword(s):  
Mathematical Models ◽  
Electric Propulsion ◽  
Rocket Engine ◽  
Power Level ◽  
Specific Impulse ◽  
Rocket Engines ◽  
Low Thrust ◽  
Constant Power ◽  
Control Functions ◽  
Thrust Control

Carrying out low-thrust transfers of spacecrafts in the near-earth space from intermediate elliptic to the geostationary orbit using electric rocket engines seems to be one of the most important tasks of modern cosmonautics. Electric rocket engines, whose specific impulse of the reactive jet is an order of magnitude more than in chemical RD, are preferable for interorbit flights with a maximum payload in the case when a significant increase in the duration of the maneuver is permissible. Ability to throttling the rocket engine thrust is traditionally considered as one of the ways to reduce both the engine mass and the required fuel assumptions for performing the specified maneuver. Using the concept of an ideal-rocket engine provides the upper estimates of the payload mass of interborbital flights for the given power level. Accounting for the properties of real engines leads to the need of considering the mathematical models with more strict limits on control functions. A study of the efficiency of three modes of thrust control of an electric propulsion rocket engine was carried out when performing practically interesting spacecraft flights from highly elliptical intermediate near-earth orbits to geostationary orbits. A mathematical model of constant power relay rocket engine has been built. The formulation of the variational problem of the Maer type is given about the execution of a given dynamic maneuver for the throttled and unregulated electric rocket engines of constant power. Using the Pontryagin maximum principle, an analysis of the optimal control functions was carried out, for which the final relations were written out, which allowed to write down the system of differential equations of the optimal movement of the spacecraft, equipped with relay electric rocket engine. The obtained numerical and quality results of the study of the effectiveness of various modes of thrust control of an electric propulsion engine to increase the payload of a given orbital maneuver confirmed the correctness of mathematical models of throttled and relay engines and, in general, the efficiency of using solutions of the averaged equations of optimal motion of a spacecraft for numerical solution of the corresponding boundary value problems in an exact formulation.

Indirect low-thrust trajectory optimization with gridded ion thruster model

2021 ◽  
pp. 095441002110438
Author(s):  
Zhemin Chi ◽  
Yang Wang ◽  
Lin Cheng
Keyword(s):  
Trajectory Optimization ◽  
Electric Propulsion ◽  
Specific Impulse ◽  
Input Power ◽  
Minimum Time ◽  
Mission Design ◽  
Low Thrust ◽  
Ion Thruster ◽  
Solar Electric Propulsion ◽  
Operational Constraints

The work deals with indirect optimization of minimum-time and minimum-fuel interplanetary trajectories when gridded ion thruster models are considered. Using an accurate model of solar electric propulsion is beneficial in preliminary mission design, and allows including operational constraints. The maximum thrust and the specific impulse are expressed as a function of thruster input power, which is achieved by means of point-fitting lines that match the performance capabilities of the thrusters. Minimum-time and minimum-fuel problems are formulated to be solved by indirect optimization. In order to increase the accuracy and robustness of the shooting procedure, analytic Jacobians are derived, and a hybrid switching detection technique is used to improve the integration accuracy for minimum-fuel problems. Two examples of Earth-to-Mars transfer and Near-Earth rendezvous mission using the realistic NASA’s Evolutionary Xenon Thruster (NEXT) are given to substantiate the feasibility and effectiveness of the proposed method.

Conceptual design studies of an electrically propelled upper stage for deployment of a multi-plane orbital constellation of small spacecraft

2021 ◽  
pp. 97-108
Author(s):  
Aleksandr V. LEVANDOVICH ◽  
Dmitry A. MOSIN ◽  
Viktor V. SINYAVSKIY ◽  
Aleksandr Ye. TYUTYUKIN ◽  
Igor A. UPTMINTSEV
Keyword(s):  
Conceptual Design ◽  
Electric Propulsion ◽  
Aerodynamic Drag ◽  
Orbital Plane ◽  
Design Studies ◽  
Small Spacecraft ◽  
Solar Arrays ◽  
Upper Stage ◽  
The Difference ◽  
Orbital Constellation

The paper presents results of conceptual design studies to determine configuration of an electrically propelled upper stage (EPUS) – a space transportation stage (a space tug) with main engines based on electric propulsion powered by solar arrays. It addresses the problem of deploying a multi-plane orbital constellation of small spacecraft (SSC) using an electrically propelled upper stage. It proposes to change the SSC operational orbital planes based on the effect of the difference in precession rates between the parking and the working orbits owing to the effect of eccentricity in the Earth gravitational field. Requirements have been defined for the EPUS electrical propulsion system that take into account the need to operate it to offset the aerodynamic drag while waiting in the parking orbit for the SSC operational orbital plane to turn. It demonstrates the feasibility of employing four EPUS that use Stationary Plasma Thruster-type electric propulsion as their main engines and gallium arsenide solar arrays for deployment in a 600 km orbit in four planes an orbital constellation of 24 small spacecraft with a mass of ~250 kg each using one launch of a medium capacity launch vehicle of Soyuz-2.1b type. Key words: Electrically propelled upper stage, electric propulsion, small spacecraft, orbital constellation.

Rendez vous with Saturn's rings

1984 ◽  
Vol 75 ◽  
pp. 743-759 ◽  
Author(s):  
Kerry T. Nock
Keyword(s):  
Solar Energy ◽  
Electric Propulsion ◽  
Power Source ◽  
Propulsion System ◽  
Low Thrust ◽  
Ring Plane ◽  
The Earth ◽  
Total Impulse ◽  
Saturn's Rings

ABSTRACTA mission to rendezvous with the rings of Saturn is studied with regard to science rationale and instrumentation and engineering feasibility and design. Future detailedin situexploration of the rings of Saturn will require spacecraft systems with enormous propulsive capability. NASA is currently studying the critical technologies for just such a system, called Nuclear Electric Propulsion (NEP). Electric propulsion is the only technology which can effectively provide the required total impulse for this demanding mission. Furthermore, the power source must be nuclear because the solar energy reaching Saturn is only 1% of that at the Earth. An important aspect of this mission is the ability of the low thrust propulsion system to continuously boost the spacecraft above the ring plane as it spirals in toward Saturn, thus enabling scientific measurements of ring particles from only a few kilometers.

scholarly journals Design and In-orbit Demonstration of REGULUS, an Iodine electric propulsion system

2021 ◽  
Author(s):  
Nicolas Bellomo ◽  
Mirko Magarotto ◽  
Marco Manente ◽  
Fabio Trezzolani ◽  
Riccardo Mantellato ◽  
...  
Keyword(s):  
Electric Propulsion ◽  
Specific Impulse ◽  
Input Power ◽  
Propulsion System ◽  
Electric Propulsion System ◽  
Propulsive Performance ◽  
Plasma Thruster ◽  
Total Impulse ◽  
Acceptance Tests

AbstractREGULUS is an Iodine-based electric propulsion system. It has been designed and manufactured at the Italian company Technology for Propulsion and Innovation SpA (T4i). REGULUS integrates the Magnetically Enhanced Plasma Thruster (MEPT) and its subsystems, namely electronics, fluidic, and thermo-structural in a volume of 1.5 U. The mass envelope is 2.5 kg, including propellant. REGULUS targets CubeSat platforms larger than 6 U and CubeSat carriers. A thrust T = 0.60 mN and a specific impulse Isp = 600 s are achieved with an input power of P = 50 W; the nominal total impulse is Itot = 3000 Ns. REGULUS has been integrated on-board of the UniSat-7 satellite and its In-orbit Demonstration (IoD) is currently ongoing. The principal topics addressed in this work are: (i) design of REGULUS, (ii) comparison of the propulsive performance obtained operating the MEPT with different propellants, namely Xenon and Iodine, (iii) qualification and acceptance tests, (iv) plume analysis, (v) the IoD.

The Aerodynamic Drag of Perforated Plates at Zero Incidence

1958 ◽  
Vol 62 (568) ◽  
pp. 301-303 ◽  
Author(s):  
P. Minton ◽  
J. R. D. Francis
Keyword(s):  
Pipe Flow ◽  
Aerodynamic Drag ◽  
Perforated Plate ◽  
Perforated Plates ◽  
Drag Forces ◽  
Flow Experiments

Perforated Plates have been used at large angles of incidence to produce drag forces and evidence on their properties has been published by de Bray. Less appears to be known about the drag forces on such surfaces at zero incidence, although they are usually considered to be aerodynamically rough. This has been confirmed by Ambrose, who carried out pipe flow experiments using perforated liners which fitted tightly in the bore of a pipe. Perforated plates used in this way do not allow flow completely through them and give “pitted” surfaces. If a perforated plate is mounted so that it is possible for cross flows to occur between the main flows on both sides of the plate the drag forces on it may be affected, and in this case the perforations will be referred to as “holes.”

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