Improving the Efficiency of Earth Monitoring Missions ... - ScienceDirect

37 downloads 18 Views 1MB Size Report
2. Estimation of orbital period evolution intensity under the influence of .... The next stage of the study was to modeling the processes of station keeping and ...
Available online at www.sciencedirect.com

ScienceDirect Procedia Engineering 185 (2017) 198 – 204

6th Russian-German Conference on Electric Propulsion and Their Application Improving the efficiency of Earth monitoring missions by equipping small spacecraft AIST-2 with electric propulsion

V. V. Salmina, S. I. Tkachenkoa, V. V. Volotsueva,*, I. V. Kaurova a

Samara University, 34 Moskovskoe shosse, Samara 443086, Russia

Abstract The paper considers the problem of equipping a small spacecraft (SS) of the AIST-2 series with electric propulsion for orbital plane maneuvers, such as station keeping and attitude control. Solution of this problem will increase the satellite's operational service life on a low-Earth orbit and, in future, allow attitude control of a satellite constellation. Published byAuthors. ElsevierPublished Ltd. ThisbyisElsevier an openB.V. access article under the CC BY-NC-ND license © 2017 The (http://creativecommons.org/licenses/by-nc-nd/4.0/). Peer-review under responsibility of the organizing committee of RGCEP – 2016. Peer-review under responsibility of the scientific committee of the 6th Russian-German Conference on Electric Propulsion and Their Application Keywords: small spacecraft; electric propulsion; aerodynamic drag; design parameters.

1. Introduction A promising area of space monitoring development is distant Earth probing by small spacecraft. The low mass of such spacecraft places additional constraints on the acceptable design parameters of the main and supporting on-board systems, which in turn determine the efficiency values of an observation spacecraft, such as spatial resolution. For some on-board systems it is possible to maintain functional efficiency, while meeting the mass constraints by miniaturization. Other types of on-board instruments, due to mass and size constraints, will substantially limit the spacecraft's function. An example of this is the solar electric power system: it is difficult to place large solar panels on a small spacecraft, which limits available electric power. Spacecraft of limited size, as a rule, are equipped with an optical system with shorter focal distances as compared to "heavy" satellites. To improve the spatial resolution of the images of Earth's surface, it is desirable to position the SS in low orbit. However, in low orbit, atmospheric drag contributes to faster evolution of the orbital parameters (such as orbital period), and a station keeping propulsion system becomes necessary. If a chemical propulsion system is used, the propellant supply necessary for a long service life (five years and up) will be too great to place this on board due to mass constraints.

* Corresponding author. Tel.: +79276875450; E-mail address: [email protected]

1877-7058 Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license

(http://creativecommons.org/licenses/by-nc-nd/4.0/). Peer-review under responsibility of the scientific committee of the 6th Russian-German Conference on Electric Propulsion and Their Application

doi:10.1016/j.proeng.2017.03.300

V.V. Salmin et al. / Procedia Engineering 185 (2017) 198 – 204

An alternative option for an orbit correction system, is an electric propulsion thrust system (EPTS). Electric propulsion systems, with a specific impulse significantly higher than that of chemical propulsion systems, open the road for a further increase in spacecraft efficiency, as they ensure better load ratios. Decrease of the propellant mass required for the spacecraft's maneuvers and orientation is provided at the expense of a substantial decrease in mass efficiency of the on-board electric power system, because for every gram of thrust developed by, for instance, stationary plasma thrusters, they consume from 150 to 200 Wt of on-board electric power. The electric power systems of modern spacecraft feature specific power of ~0,1 kWt/kg [1]. An important characteristic of EP is that its firing consumes electric power, which is in limited supply on board. Therefore, the problem of equipping a SS with correctional EPTS requires an analysis of the constraints associated with providing an on-board electric power supply for the firing of the thrusters. 2. A study into the applicability of equipping SS AIST-2 with a correcting EPTS A study into the applicability of equipping a small spacecraft AIST-2 with an EP unit for flight correction must solve the following problems: 1. Calculation of aerodynamic drag forces that act on the small spacecraft under different conditions of the upper layers of the Earth's atmosphere. 2. Estimation of orbital period evolution intensity under the influence of aerodynamic drag. 3. Selection of design parameters for EP cruising thrusters (thrust value, power consumption, type of EP) and estimation of EP efficiency range. 4. Modelling station keeping processes and EP maneuvers for long periods of time. 5. Selection of design parameters for the propellant storage and supply system of the correcting thruster unit and development of EP-powered SS design layout. Before examining these problems, it is necessary to consider the spacecraft under study. The AIST-2 small spacecraft was designed jointly by the Rocket and Space Center "Progress", Samara, and Samara University, for the following missions: - development of a small spacecraft platform (under 250 kg) design and on-board equipment for distant probing of the Earth; - development of software for a small spacecraft platform; - development of instruments, ground control devices, receipt and processing of data, and methods of processing the data from high fidelity and capture Earth probing sensors; - development of radiolocation methods of the Earth surface and sub-surface in ultra short wave length; - development of real-time follow-through technologies for scientific experiments in outer space using Internet-based information and communication satellite technologies on the "Globalstar" platform [2]. The main technical characteristics of the AIST-2 SS are represented in Table 1 [2]. AIST-2 exterior design and its inner arrangement are shown in Fig. 1 and 2.

Fig. 1. AIST-2 exterior view

199

200

V.V. Salmin et al. / Procedia Engineering 185 (2017) 198 – 204 Table 1. Technical characteristics of AIST-2 SS.

Characteristic

Value.

Insertion method Operational orbit parameters for launch with Volga IM – near circular with mean altitude, km. – inclination, degrees Capture range for visible specter equipment, km Capture range for IR specter equipment, km Rate of target data transmission, Mbit/Sec Data storage capacity, Gbites

By Soyuz-2-1a launch vehicle with Wolga insertion module

510 97,4 41,3 48,9 64-175 No less than 32

Operational lifetime, years

No less than 3-х

Average daily solar panel power, Wt

No less than 285

Spacecraft mass, kg

470

Mass of on-board equipment and scientific instruments, kg Orientation by all channels (3σ) – in orbital system of coordinates – solar orientation

No less than 125 By angle: 10ˊ by yaw, 30′ by roll and pitch / by velocity ≤ 0,005 °/с By angle ≤ 30 /by velocity ≤ 0,5 °/с

Transmission of data to Earth

By radio (bandwidth 8,025 GHz to 8,4 GHz)

Fig. 2. AIST-2 inner layout.

201

V.V. Salmin et al. / Procedia Engineering 185 (2017) 198 – 204

Let us now return to the primary problems of applicability of EPTS for small spacecraft and consider them in detail. The aerodynamic drag force is calculated by the following equation:

Fаэр

mКА ˜ aаэр

mКА ˜ с х ˜

where aаэр is aerodynamic acceleration;

mКА

G S мид ˜ U h, t ˜ V ˜ V 2 ˜ mКА

is mass of spacecraft;

сх ˜

сх

G S мид ˜ U h, t ˜ V ˜ V 2

(1)

is aerodynamic coefficient; S мид is mid-

G

section of the spacecraft; U h, t is the density of the upper layers of the Earth's atmosphere; V is the velocity vector of incoming air flow. From the expression (1) it can be seen that the aerodynamic drag force depends on mid-section, aerodynamic coefficient, atmospheric density and parameters of operational orbit. The density of the upper layers of Earth's atmosphere is a complex function of spacecraft altitude and current flight time, which is why the aerodynamic drag value changes exponentially with orbit altitude and "complex-periodically" with flight time. Estimation of the orbit evolution intensity for the operational orbit of the AIST-2 spacecraft shows that with low solar activity (minimal atmospheric density) the altitude will decrease, on average, by about 10 meters in 3.5 days (Fig. 3a). With high solar activity (very dense atmosphere) the 10 meter decrease takes place in only about one day (Fig. 3b).

a)

b)

a) with minimal atmospheric density; b) with maximum atmospheric density Fig. 3. Change in the altitude of SS AIST-2 orbit:

Consequently, in first approximation it is possible to state that after a year in orbit the mean altitude for SS AIST-2 may decrease by 1.04 to 3.65 km. This is a significant deviation from the planned parameters for the spacecraft. SS AIST-2 was launched into space on April 28, 2016 and has been since operating successfully. In the six months in orbit its mean altitude decreased by 2.5 km. Selection of the design parameters for a cruising EP unit has to satisfy the condition:

FЭРДУ t Fаэр Where

FЭРДУ

n ˜ FЭРД FЭРД

is the thrust power of the electric propulsion unit;

(2)

n is the number of thrusters firing at a

given moment; is the thrust power of one thruster (determined by its type). At the SS AIST-2 operational orbit (altitude 510 km) the maximum power of aerodynamic drag is about 2mN (Fig. 3a). Consequently, to guarantee that the condition (2) is satisfied it is sufficient to have an EP system with one EP thruster constantly firing to provides at least 2 mN of thrust. The authors of this study stipulated the following conditions: a) the desirability to use existing EP thrusters made by domestic producers (such as Fakel Research Bureau or the Keldysh Center); b) the low amounts of electric power available on board of a small spacecraft which places limits on the specific firing time of EP thrusters on active revolutions (the algorithm implies periodic engagement and shutdown of thrusters). The thruster that was considered the most satisfactory in view of the above constraints was the SPD-50 made by Fakel (Kaliningrad). It can provide a maximum thrust of about 20 mN with electric power consumption of about 300 Wt.

202

V.V. Salmin et al. / Procedia Engineering 185 (2017) 198 – 204

For attitude control of a spacecraft transversal thrust must be applied both in positive and negative directions. Therefore the EPTS must contain at least two SPD-50 thrusters fixed to opposing panels of the spacecraft. In addition, the system must have backup options in case the main thruster fails. Consequently, it is assumed that the EPT consists of four SPD-50 thrusters, of which only one is firing at any given moment. The next stage of the study was to modeling the processes of station keeping and spacecraft maneuvering by means of EPTU over a long period of time. The result of this stage was to build cyclograms of EPTU operation and estimate average propellant consumption. The design parameters selected for the SS AIST-2 EPTS are shown in Table 2. Table 2. Main design parameters of electric propulsion thrusting system.

Parameter Torque of a cruising thruster SPD-50, mN Specific impulse of SPD-50 cruising thrusters, m/s Electric power consumption at transfer stage, kWt Electric power consumption at GEO, kWt Propellant Time of obtaining 90% of nominal charge from appearance of the charge, s Minimal interval between shutoff and re-firing of a thruster, min.: • on the same cathode • on different cathodes Operation life of any of the four thrusters: • by thrust time • by number of firings Spatial angle of the plasma stream spread from geometrical axis (allowing for error of measurement) for complete service life, degrees, no more than Deviation of the thrust vector from its geometrical axis (allowing for error of measurement) for complete service life, degrees, no more than

Nominal value 20 1750 0,2…0,6 0,2…0,6 Xenon No more than 60 15 10 No less than 4500 hours No less than 5000 ±450 (95% of the ion flow generated by the thruster) ±1

3. Development of layout for an improved SS AIST-2 with EPTS The mass distribution of the EPTS is provided in Table 3. The exterior view is shown on Fig. 4 and 5. Table 3. Mass distribution of the EPTS.



Element

Mass, kg.

Notes

1

SPD-50 thruster.

5,6

4 units.

2 3

Fixing and drives. Conversion and control system.

10 20

4 units.

4 5 6

On-board instruments (receiver etc) Propellant tanks On-board cables

6 4 3

2 units. К = 0,13

7 8 9

Main frame Xenon Reserve

20 10 5

Total:

83,6

V.V. Salmin et al. / Procedia Engineering 185 (2017) 198 – 204

Fig. 4. Exterior view of thruster unit.

Fig. 5. Exterior view of xenon storage unit.

When equipping a spacecraft with electric propulsion thruster system, one has to consider the increase in electric power consumption, as well as the increase in total spacecraft mass due to the mass of the EPTS. Due to this, some scientific instruments of the AIST-2 had to be discarded, and the solar panel area increased to compensate for additional energy required. Therefore, the external layout and dimensions of the spacecraft remained practically unchanged. The inner layout, naturally, suffered more considerable changes. The external view of the AIST-2 with EPTS and the location of its on-board instruments is shown on Fig. 6 and 7.

Fig. 6. Exterior view of SS AIST-2 with EPTS

203

204

V.V. Salmin et al. / Procedia Engineering 185 (2017) 198 – 204 V. V. Salmin / Procedia Engineering 00 (2017) 000 000

Fig. 7. Location of on-board instruments of SS AIST-2 with EPTS

Table 4 shows the main parameters that were changed due to equipment of the SS with EPTS. Table 4. Main parameters that were changed due to equipment of the SS with EPTS

Parameter

Value AIST-2 with EPTS No less than 5

Operational service life, years.

AIST-2 No less than 3-х

Average daily electric power consumption, Wt.

No less than 285

No less than 400

Mass of the spacecraft, kg.

470

470

Payload mass, kg.

No less than 125

No less than 80

5. Conclusions The present study on equipping a small spacecraft of the AIST-2 series with an electric propulsion thruster system leads to the conclusion that application of an EP system allows the spacecraft's operation lifetime to be significantly increased in a low Earth orbit. The system can be also used, if need arises, for final insertion, orbital formation keeping, emergency orientation and a decrease of the load on the rotor orientation system. In future, this concept will enable attitude control over a satellite constellation. The algorithm used to select the design parameters of a low orbit SS with EPTS was the search for the optimal design solution for minimum SS mass. The proposal for equipment of AIST-2 small spacecraft with electric propulsion is illustrated by a 3D model created in Creo Elements/Pro software. Comparative analysis of the parameters of SS AIST-2 with and without EPTS demonstrates the advantages of the EP-powered spacecraft. References [1] Malyshev G. V., Kulkov V. M., Egorov Y. G. "Application of electric propulsion for satellite insertion, orbit correction and formation keeping". In Russian. Polyot, 2006, № 7, pp.34-40 [2] Kirillin A. N., Shakhmatov, E. V. et. al. "Scientific and technological experiments of a university AIST series small spacecraft constellation." In Russian. "Scientific and technological experiments on automatic spacecraft and small satellites: Theses of the 3rd international conference Samara, September 9-11 2014г., pp.149-154.