Minimize SLATS

SLATS (Super Low Altitude Test Satellite)

SLATS is a technology demonstration minisatellite of JAXA to orbit the Earth at an altitudes in the range of 180-250 km. Its main objectives are to understand the effects of high-density atomic oxygen on the satellite and to verify the possibility of orbit control using an ion engine system. In this mission, an ion engine system is used to compensate for air drag, which cannot be neglected in such a low Earth orbit. 1) 2) 3) 4) 5) 6)

Specific goals of the mission are:

1) Verification of super low altitude satellite system

2) Measurement of atmospheric density in super low altitude

3) On-orbit monitoring of atomic oxygen to understand the effects of high-density atomic oxygen on the satellite.

SLATS is planned to be launched as a secondary payload on the ALOS-2 mission in 2013/14 on the H-IIA vehicle. Hence, the microsatellite will be deployed into a sun-synchronous orbit of ~630 km altitude, inclination = 97.9º, the LSDN (Local Sun time on Descending Node) is 12:00 hours ± 15 min.

In the IOC (Initial Orbit Control) operation period, SLATS will reach the 250 km altitude and LSDN of 16:00 hours orbit using its own chemical propulsion system (RCS). At this level and below, the ion engine is used to compensate for the atmospheric drag. A long duration operation of the ion engine is needed.

Orbit

Initial: 628 km, circular orbit, LSDN = 12:00 hours
Mission: 250-180 km, circular orbit, LSDN = 16:00 hours

Minisatellite mass, size

~400 kg, 2.5 m x 5.2 m x 0.9 m (deployed)

Mission life

~1.5 years

Electric power consumption

~700 W (max)

Sensor complement

AOFS (Atomic Oxygen Fluence Sensor)
MDM (Material Degradation Monitor)
OPS (Optical Sensor) for Earth observation

Ion engine

Thrust: 10-28 mN (1 engine)
Isp: ~2000 s
Electric power: 370 W @ 10 mN
Fuel mass: 10 kg (Xenon)

Chemical propulsion engine

Thrust: 1 N x 4 engines
Isp: ~ 200 s
Fuel mass: 34 kg (N2H4)

Table 1: Overview of the SLATS performance parameters

In mid-2012, SLATS is in the critical design phase. The project is developing a new ion engine system for the SLATS and the practical satellites to follow based on the ETS-VIII/Kiku-8 ion engine system.

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Figure 1: Illustration of the super low altitude satellite concept (image credit: JAXA)

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Figure 2: Orbit profile of the SLATS mission (image credit: JAXA)

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Figure 3: Artist's rendition of the deployed SLATS minisatellite (image credit: JAXA)

Verification of super low altitude satellite system:

In the super low altitude orbit, ~200 km high, atmospheric density is too high to keep satellite’s altitude by high air drag. Figure 4 shows the relation between the altitude and the air density. By the study of the air drag for the super low altitude satellite, it is found that the satellite's aerodynamic shape and the particular attitude direction to the velocity vector is an important factor for drag compensation. In order to observe the Earth continuously at super low altitude orbit, it is necessary to cancel the air drag to maintain the satellite’s altitude.

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Figure 4: Atmospheric density as a function of altitude (image credit: JAXA)

The SLATS project will conduct on-orbit engineering tests regarding the orbit and altitude control of the spacecraft with the ion engine system as shown in Figure 5.

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Figure 5: SLATS operational profile (image credit: JAXA)

Measurement of atmospheric density in super low altitude:

There is a need for air density measurements in the altitude range of ~200 km. Currently, applicable data is estimated by using an air density model based on actual measurement in nominal LEO (Low Earth Orbit), higher than about 400 km and on the ground. In the aerodynamic design of SLATS, the project takes account of the worst case in estimated data. Actual measurements of atmospheric density data will be acquired to be correlated with AOCS and GPS receiver data to study the effect caused by atmospheric density.

Monitoring on-orbit atomic oxygen data:

Since the atmosphere in the altitude range of 180-200 km contains a lot of atomic oxygen (AO), it can deteriorate the materials of a satellite and may cause some unexpected issues. Hence, actual AO density data measurements in the super low altitude range is also a very desirable parameter next to that of atmospheric density data. The project requirements call for the degradation monitoring of material samples (surface change) in the super low altitude environment.

Spacecraft:

The SLATS microsatellite is intended to be a precursor mission to follow-on operational super low altitude satellites. 7)

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Figure 6: Functional block diagram of SLATS (image credit: JAXA)

NSTT (Next-generation Star Tracker): JAXA in collaboration with NEC Toshiba Space Systems is developing NSTT to provide high-accuracy attitude determination results: the random error is < 4 arcsec (3σ), while the bias error is < 6 and 4 arcsec (3σ), respectively, for wide and narrow temperature ranges. NSTT is able to track and acquire stars under high attitude-rate of 2º/s with 99.9% probability. A demonstration/qualification flight of NSTT is scheduled for SLATS. 8)

Parameter

Performance specification

Attitude accuracy
(random)

Cross boresight axis

< 4 arcsec (3σ, angular rate is < 0.3º/s)

Boresight axis

< 40 arcsec (3σ, angular rate is < 0.3º/s)

Attitude accuracy

(bias)

Cross boresight axis

< 6 arcsec (3σ, accuracy is defined within operation temperature)
< 4 arcsec (accuracy is defined within the range of a reference temperature ±5ºC

Boresight axis

< 20 arcsec (3σ, accuracy is defined within operation temperature)

Max acquisition rate

> 3º/s (99.9% of the whole celestial sphere)

FOV (Field of View)

16º x 16º

Output rate

4 Hz

Sun avoidance angle

25º

Maximum power consumption

< 20 W

Instrument mass

< 6.2 kg

Vibration tolerance

> 20 G rms

Shock tolerance

> 1000 G srs

Operation temperature

-25º to 55ºC

Table 2: NSTT performance specification

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Figure 7: Block diagram of NSTT (image credit: JAXA)

 

IES (Ion Engine System) : The IES consists of a PMU (Propellant Management Unit), an ion thruster and a PPCU (Power Processing Control Unit), including an Ion thruster controller. The PMU supplies xenon gas to the ion thruster. The PPCU supplies electrical power to the ion thruster according to on-board control sequence. The ion thruster generates ion beam and thrust. 9)

The PMU is almost the same unit as used in the ETS-VIII (Kiku-8) propellant management system. Xenon is used as propellant and is stored in three tanks at a pressure of approximately 7MPa. The xenon mass flow rate into the thruster is controlled properly through a high-pressure latching valve, a regulator, a low-pressure latching valve and mass flow control devices. The mass flow control device is a fixed orifice and the mass flow rate is determined only based on the upstream pressure. The mass flow rate is adjusted at 10.5 cm3/s (including 2 cm3/s for the neutralizer).

The PPCU, developed at MELCO and JAXA, is a new 20 mN class ion engine system used for air drag compensation on the spacecraft. The PPCU consists of seven power supplies (PS-1~PS-7) for an ion thruster, an auxiliary power converter, a primary bus interface, and a signal interface circuit. The Kiku-8 ion thruster has the lifetime of 16,000 hours at 20mN; hence, the lifetime satisfies the SLATS requirement. 10) 11) 12) 13)

The ion engine (Figure 8) produces thrust through the reaction force obtained when ions accelerated by an electrical force are expelled. It consists of an "ion generator" (generates ions from a propellant such as xenon), an "ion acceleration system" (accelerates those ions), a "neutralizer" (emits electrons for neutralizing the ions expelled into space), seven "power sources" (operate the above three sections), and a "controller" (controls the power sources).

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Figure 8: Schematic view of the ion engine concept and the electrical interface between ITR and PPCU (image credit: JAXA, MELCO)

The electrical interface with the satellite system has been changed from the Kiku-8 system. The primary input voltage of the Kiku-8 PPU is 100 V. On the other hand, the SLATS PPCU is designed to work under an input voltage between 24 V and 32 V. The control algorithm is different from that of the Kiku-8 IES. The SLATS keeps its altitude by thrust on/off control. The on/off command must be given to the PPCU autonomously as the contact time from a ground station is very short. The thrust is automatically generated or stopped using orbit data gained by an on-board GPS receiver.

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Figure 9: Functional block diagram of the IES (image credit: JAXA, MELCO)

The signal interface circuit is connected to the satellite data bus and contains the operation modes. The ORBIT mode is used in the altitude keeping operation of SLATS. The control logic of the operation modes are installed in FPGAs. PS-1 and PS-2 use the same high-voltage transformer to synchronize their output voltage. Their power efficiency of 90 % is required.

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Figure 10: Block diagram of the PPCU (image credit: JAXA, MELCO)

Mode

Operation

IDLG

Neutralizer and main hollow-cathode heating (PS-4 and PS-6 on, low level)

ACTV_N

Neutralizer hollow-cathode activation (PS-6 on, high level)

ACTV_M

Main hollow-cathode activation (PS-4 on, high level)

NEUT

Neutralizer discharge on (PS-7 on)

DISC

Main hollow-cathode discharge and main discharge on (PS-5 and PS-3 on)

ORBIT

All discharge on and beam on/off by command from orbit control system (PS-3, PS-5 and PS-7 on and PS-1 on/off by command)

CM

Grid cleaning, short-circuit opening (function for failure)

Table 3: PPCU operational modes

Power Supply

Name

Voltage range (V)

Current range (A)

Ripple (%) P.P.

Regulation (%)
C.V (constant voltage)

C.V (constant current)

Maximum Power (W)

PS-1

Beam PS

800~1,100

0.2~0.6

5

C.V ±3

660

PS-2

Accelerator PS

400~500

0.001~0.1

5

C.V ±5

5.5

PS-3

Discharge PS

45
100

1.5~3.5
0.1

5
10

C.V ±5
C.V ±5

140
12.5

PS-4

Main hollow cathode heater PS

15

1.4~4.0

5

C.C ±3

60

PS-5

Main hollow cathode keeper PS

15
150

0.5
0.01

5
20

C.C ±5
C.C ±5

7.5
12.5

PS-6

Neutralization hollow cathode heater PS

15

1.4~4.0

5

C.C ±3

60

PS-7

Neutralization hollow cathode keeper PS

30
150

0.5
0.01

5
20

C.C ±5
C.C ±5

15
1.5

Table 4: PPCU output characteristics

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Figure 11: Illustration of the PPCU ion thruster (image credit: JAXA, MELCO)

Engine type

Kaufman

Propellant

Xenon

Thrust

11.5~17mN

Mass flow rate

10.5 cm3/s

Instrument mass, primary voltage

43 kg, 24-32 V

Power consumption

<403 W @ 11.5 mN, <580W @ 17 mN

Signal interface

RS422

Propellant mass

12 kg

MEOP

15 MPa (LBB designed tank)

Proof pressure

>1.5 x MEOP (LBB designed tank)

Burst pressure

2.0(tank), 2.5(valve), 4.0(tube etc.) x MEOP (LBB designed tank)

Leakage

internal: <0.5 Pa m3/h(Ghe), external: <1x10-7 Pa m3/s(Ghe)

Table 5: Main performance parameters of the SLATS IES

 

Launch: A demonstration/qualification flight of SLATS as a secondary payload is planned for 2016. 14)

Orbit: Initial sun-synchronous orbit, altitude 500-600 km.

After deployment and on-orbit checkout, the SLATS orbit altitude is reduced to 250 km using chemical thrusters and atmospheric drag. After reaching the circular altitude of 250 km, the experiment of altitude control begins by using the IES. The IES generates thrust autonomously when the satellite altitude is lower than the target altitude. The SLATS has GPS receivers provides the altitude data. An experiment period of over 90 days is planned to get the attitude disturbance data due to air drag and the atomic oxygen data at the surface and inside of the satellite body. Those data will be used in the design of practical super low altitude satellites.

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Figure 12: Launch configuration of SLATS (image credit: JAXA)

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Figure 13: Configuration of SLATS at launch and on orbit (image credit: JAXA)

 


 

Sensor complement (AOFS, MDM, OPS)

The super-low altitude of SLATS requires to observe the material response in the surrounding environment of simultaneous hyperthermal collisions both of AO (Atomic Oxygen) and of N2.

Note: a description of the sensor complement will be provided when available.

AOFS (Atomic Oxygen Fluence Sensor)

The AO (Atomic Oxygen) effects are measured by AOFS (Atomic Oxygen Fluence Sensor) and MDM (Material Degradation Monitor). AOFS consists of eight TQCMs (Thermoelectric Quartz Crystal Microbalances ) with polyimide film and its controller. The TQCMs are located inside/outside of the SLATS structure, they measure the mass decrease of polyimide film which reacts with AO to become gas. The amount of AO is calculated on the basis of the mass decrease data.

MDM (Material Degradation Monitor)

MDM consists of the material samples and the optical camera. The samples consist of the conventional outside satellite materials and the next generation AO-resistance materials and so on. They are located on the under surface of the SLATS structure facing the moving direction. The optical camera takes imagery of the samples for studying the mechanism of degradation due to AO.

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Figure 14: Photo of the MDM EM (Engineering Model), image credit: JAXA

OPS (Optical Sensor) for Earth observation

The small OPS system will observe Earth's surface to verify the merit in spatial resolution. The spatial resolution of optical cameras is proportional to the altitude. The project will examine the change of spatial resolution following orbit transfer.

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Figure 15: Photo of the OPS FM (image credit: JAXA)


1) Hiroshi Nagano, Kenichi Kajiwara, Yukio Hayakawa, Toshiyuki Ozaki:, Hiroyuki Osuga, “Development of the ion engine system for SLATS,“ Proceedings of the 61st IAC (International Astronautical Congress), Prague, Czech Republic, Sept. 27-Oct. 1, 2010, IAC-10.C4.4.1

2) Hiroshi Nagano, Kenichi Kajiwara, Hiroyuki Osuga, Toshiyuki Ozaki, Takafumi Nakagawa, Kazuo Shuto, “Development Status of a New Power Processing Unit of Ion Engine System for the Super Low Altitude Test Satellite,” 31st International Electric Propulsion Conference, University of Michigan, Ann Arbor, Michigan, USA, Sept. 20–24, 2009, paper: IEPC-2009-058, URL: http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2009index/IEPC-2009-058.pdf

3) Shunsuke Imamura, Masayoshi Utashima, “Initial Orbit Control of SLATS using Atmospheric Drag,” Proceedings of the 28th ISTS (International Symposium on Space Technology and Science), Okinawa, Japan, June 5-12, 2011, paper: 2011-d-48

4) Akio Tsujihata, “Strategy and R&D for Space Applications Mission,” Microelectronics Workshop,Tsukuba International Congress Center, Japan, Nov. 10, 2010, URL: https://eeepitnl.tksc.jaxa.jp/mews/en/23rd/data/10-03.pdf

5) Kazuya Konoue, S. Yamakawa, S. Imamura, H. Kohata, “Development of Super Low Altitude Test Satellite (SLATS),” Proceedings of the 63rd IAC (International Astronautical Congress), Naples, Italy, Oct. 1-5, 2012, paper: IAC-12-B1.2.18

6) Haruo Kawasaki, Hiroki Kohata, Kazuya Konoue, Yohei Satoh, Shunsuke Imamura, “Development of the Super Low Altitude Test Satellite and Thermal Control,” Proceedings of the 29th ISTS (International Symposium on Space Technology and Science), Nagoya-Aichi, Japan, June 2-8, 2013, paper: 2013-r-27

7) Kazuya Konoue, Nobuaki Igarashi, Shunsuke Imamura, Shiro Yamakawa, Hiroki Kohata, “Development of Super Low Altitude Test Satellite (SLATS),” Proceedings of the 28th ISTS (International Symposium on Space Technology and Science), Okinawa, Japan, June 5-12, 2011, paper: 2011-n-06

8) Shuichi Matsumoto, Takanori Iwata, Hiroshi Kawai, Isamu Higashino, Kazuhide Noguchi, Koshi Sato, “Precision Autonomous Star Tracker for Agile Spacecraft,” Proceedings of the GNC 2011, 8th International ESA Conference on Guidance, Navigation & Control Systems, Carlsbad (Karlovy Vary), Czech Republic, June 5-10, 2011

9) Hiroshi Nagano, Yukio Hayakawa, Hiroyuki Osuga, Toshiyuki Ozaki, Kazuo Shuto,“A New Orbit Control Algorithm for the 20 mN Class Ion Engine System,” Proceedings of the 33rd International Electric Propulsion Conference (IEPC), Washington D.C., USA, Oct. 6-10, 2013, paper: IEPC-2013-064, URL: http://www.iepc2013.org/get?id=064

10) Hiroshi Nagano, Yukio Hayakawa, Kenichi Kajiwara, Hiroyuki Osuga, Isao Terukina, Kentaro Suzuki, Kazuo Shuto, “A New Power Processing Control Unit for a 20 mN Class Ion Engine System,” Proceedings of the 63rd IAC (International Astronautical Congress), Naples, Italy, Oct. 1-5, 2012, paper: IAC-12.C4.4.26

11) Hiroshi Nagano, Kenichi Kajiwara, Yukio Hayakawa, Toshiyuki Ozaki, Yukikazu Kasai, “Optimization of the Operating Parameters for a 20 mN Class Ion Thruster,” 32nd International Electric Propulsion Conference, Wiesbaden, Germany, September 11 – 15, 2011, paper: IEPC-2011-032, URL: http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2011index/IEPC-2011-032.pdf

12) Katsuhiro Miyazaki, Hiroshi Nagano, Kenichi Kajiwara, Yasushi Okawa, “Research and development of an ion engine for a super-low-altitude satellite,” 2011, JAXA, URL: http://www.ard.jaxa.jp/old/eng/publication/sorasora/2011_no42/ss2011no42_02.html

13) Hiroshi Nagano, Kenichi Kajiwara, Hiroyuki Osuga, Toshiyuki Ozaki, Kazuo Shuto, “Design of a new discharge power supply for ion engine,” 14th International Power Electronics and Motion Control Conference (EPE/PEMC), Ohrid, Macedonia, Sept. 6-10, 2010

14) Keizo Nakagawa, “R&D of JAXA Satellite Application Mission,” MEWS26 (26th Microelectronics Workshop), Tsukuba, Japan, Oct. 24-25, 2013, URL: https://eeepitnl.tksc.jaxa.jp/mews/jp/26th/data/1_1.pdf


The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.