PRECURSOR MICROSATELLITE

PROJECT DESCRIPTION

LANDSPACE

LAUNCH VEHICLE

SATELLITE DESCRIPTION

ON-BOARD COMPUTER

OPERATING SYSTEM

POWER MODULE

COMMUNICATION MODULES

THERMAL CONTROL

ATTITUDE SYSTEM

PAYLOAD

STRUCTURE

DEPLOYER

ENVIRONMENTAL TESTS

GROUND SEGMENT

MISSION OPERATIONS

TEAM

PARTNERS

SPONSORS

 

 

 

PROJECT DESCRIPTION

The PRECURSOR Microsatellite is a private non-profit mission. Private people, research institutes and universities support this project. The PRECURSOR is built with both commercial and open source resources. This means hardware as well as software.
The mission of the PRECURSOR belongs to IOD/IOV (In Orbit Demonstration/In Orbit Validation) activities.

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LANDSPACE

Land Space Technology Corporation Ltd. is a Chinese private aerospace enterprise engaged in the R&D and operations of launch vehicles. Focusing on small and medium scale commercial aerospace application market, Land Space is devoted to the development of Liquid-fuel Rocket Engines (LREs) and low-cost commercial launch vehicles with independent intellectual property rights. Land Space could complete the system and unit design, manufacture, test and delivery with highly-integrated design and innovation capability by a first-class technical team, to provide the global market with standardized launch service solutions.

Land Space always considers the technical innovation and market-orientation as the core development task, and is confident to become a beneficial supplement of China Aerospace, to continuously boost its future development.

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LAUNCH VEHICLE

The PRECURSOR Microsatellite will be launched from China on June 2022 from the Jiuquan Satellite Launch Center (JSLC) in China. The launch vehicle operator is the Chinese Start-Up LandSpace and the PRECURSOR flight will be the second flight of the ZQ-2 launch vehicle. The ZQ-2 is a two-stage liquid propellant (LOX+LCH4) launch vehicle with a capability of 1.8 tonnes payload into 500 km SSO (Sun-Synchronous Orbit).

Figure 1: ZQ-2 Launcher

 

Figure 2: ZQ-2 Launcher Series

 

 

 

Figure 3: Mission Profile

 

Table 3: typical mission profile (ZQ-2 Block 1, 500km SSO)

Event

Time (s)

Altitude (km)

Velocity (m/s)

1st stage Ignition & Lift-off

0

0

0

1st stage cut-off (MECO)

151

69

1,960

2nd stage ignition

155

74

1,928

Fairing separation

212

140

2,733

2nd stage main engine cut-off (SECO)

290

239

4,896

2nd stage Vernier cut-off (VECO)

760

500

7,613

Payload deployment

770

500

7,617

 

 

Table 4: Typical Separation Accuracy

Parameter

Angle

Rate

Roll deviation

≤0.7°

≤0.3°/s

Pitch deviation

≤1.5°

≤1.0°/S

Yaw deviation

≤1.5°

≤1.0°/s

 

Table 5: Injection Accuracy

Symbol

Parameters

Deviation

a

Semi major Axis

±5km

e

≤1.5°

0.003

i

≤1.5°

±0.08

 

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SATELLITE DESCRIPTION

The PRECURSOR Microsatellite doesn’t follow any kind of standards like the CubeSat standard. The CubeSat standard has limitations and for the mission objectives a customized solution is the best option. The PRECURSOR satellite has the sizes 15x30x60 cm3. The weight is approximately 17kg. The main experiment is a new generation propulsion system for Orbital Plane Change (inclination) and/or Orbital Altitude Change and uses green propellant. The PRECURSOR Microsatellite will validate several technologies. The satellite is designed using Commercial electronics and materials, Open Source Hardware and Open Source Software. This strategy allow the quick implementation of this new satellite platform.  Other experiments inside the PRECURSOR is a multi-band GNSS receiver with a newly High-Performance GNSS Antenna for small satellites.

Figure 4: Representation PRECURSOR Satellite in orbit

 

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ON-BOARD COMPUTER (OBC)

The OBC is based on the EFR32FG12 processor from Silicon Labs. The OBC has four processors, one is the main and the other one the backup. The processor was irradiated with Gamma rays and it was demonstrated that this processor can be used in LEO orbit for several years.
Following are the most important characteristics of the OBC:

  • Four ARM Cortex M4 with 40MHz.
  • Operating supply voltage between 3.7 – 4.0 V
  • Internal FLASH memory of 1MB
  • RAM size of 256 kB
  • Interfaces available are I2C, SPI and UART
  • 46 GPIOs
  • Operating temperature in a range between +20°C to +85°C
  • Data Bus Width of 32 bit
  • ADC with a resolution of 12 bit

Figure 5: Small size computer

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OPERATING SYSTEM

The PRECURSOR Microsatellite uses RODOS (Realtime Onboard Dependable Operating System) as operating system.
RODOS was developed at the German Space Agency (DLR) and is a derivate from the BOSS operation system, which is used in the micro-satellite program DLR's. Satellites using this operation system are BIRD and TET-1 and BiROS. RODOS is ideal for Telemetry and Telecommands tasks. It is light to migrate to other systems.
RODOS is an open source software and is implemented in C++ and has also following features:

  • object-oriented C++ interfaces,
  • ultra-fast booting
  • real time priority controlled pre-emptive multithreading,
  • time management (as a central point),
  • thread safe communication and synchronisation,
  • event propagation

Figure 6: RODOS OS Logo

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POWER MODULE

The power module consists of a set of solar cells, Lithium batteries and the electronic to regulate and distribute the power in the system.

The Power Conditioning Distribution Unit (PCDU) distributes it internally to the energy consumers or to the batteries.

The Maximum Power Tracking (MPPT) charge controller allow to maximize the energy extraction generated by the solar panels.


Solar Cells
The solar cells are space grade and has the size of 65x40mm. Each cell is equipped with 12.5µm Ag interconnector welded to the solar cell n-contact. The solar cell is covered with 100µm thick cover glass type CMX with the size of 65.15x40.15mm bonded by use of silicone adhesive DC93500. The cells have an efficiency of around 30%.

Figure 7: PRECURSOR Solar Cells

Figure 8: Solar Cells Characteristics


Batteries
In order to increase the lifetime of the batteries a properly reload cycles is essential. The power electronic will configure automatically the best reload cycles and power consumption, monitoring constantly the status level of the batteries.
Each cell has following characteristics:

  • Nominal voltage: 3.6 V/cell
  • Nominal capacity: 1600 mAh /cell
  • Operating temperature: -25°C to 55°C

The charge and discharge process will be controlled with a charge controller IC, which is equipped with a MTTP (Maximum Power Point Tracking) control unit.
The power module has a power capacity of 10Ah (11V).

Figure 9: Lithium Ion Cell High Energy (Source EAS)

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COMMUNICATION MODULE

It consists in a modem and transceivers which converts the digital data in an analogue signal. The data to be transmitted will be provided by the OBC, which administrates the data distribution and encapsulation. The standard AX.25 will be used for encapsulation purposes. The communication module transmits and receives in UHF and S band.
Following transceiver will be used:
UHF Transceiver The selected transceiver will be tested in order to see the radiation robustness. Thermal-Vacuum and Outgassing tests will be also included.
Characteristics of the transceiver:

  • Data rate: 256 kbps
  • Power transmission:0,1-2W
  • Frequency: 405-480 MHz

S-Band Transceiver

Having a camera as payload is important to transfer data in a higher data rate as is possible by the UHF transceivers. For this reason for the remote sensing experiment S-Band transceiver is recommended. Following modules characteristics are considered:

  • Data rate: 1,06 Mbit/s
  • Power transmission: 2 W
  • Error correction: 32 bits CRC
  • Frequency: 2,2-2,3 GHz

 

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THERMAL CONTROL

The thermal behaviour of some sensible components like the batteries and solar cells will be monitored by using temperature sensors. The thermal sensors and the heaters will be controlled by the OBC.
Temperature Sensors The temperature sensor PT1000 will be used in the PRECURSOR. Such sensor is commonly used as thermometer in microcontrollers but also often in satellites
Heaters. The Heaters will be also used specially for batteries conditioning. The selected heaters have high quality, low power consumption and small sizes. In that way the temperature can be regulated in a range for save operation of sensible electronic, typically between -20°C and +60°C.
Heat Sink For thermal control one side (30x60 cm2) of the satellite structure will point constantly to the Earth. This side will be used as heat sink and it will dissipate the excess warmth.

 

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ATTITUDE SYSTEM

The Attitude subsystem will consist of two different parts. The first one will be a sensor set of Sun sensors, Magnetometers, Accelerometers and Gyroscopes for orbit determination.
The second part will be built with a set of actuators, consisting of magnetic coils and micro-reaction wheels.
With the mentioned configuration the Attitude system fulfils a required accuracy of +/- 5°. Further, the Attitude System will have its own microcontroller to manage the complexity of the required calculations for attitude control.
The attitude system will have following characteristics:

  • Size: 9,5x9,5x10cm
  • Mass:1000g
  • Power consumption: 0,5-1,8W
  • Attitude Knowledge: ±
  • Pointing accuracy: ±
  • Sensors: Accelerometers (3-axis); Gyroscope (3-axis); Magnetometer (3-axis)
  • Actuators: 3 reactions wheels (torque 0,087mNm / 1,5mNms)

Figure 10: Multi-sensor Board

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PAYLOAD

In the PRECURSOR mission many technologies will be tested but following main experiments are the principal payloads:

Propulsion System

The main payload consist in a new generation propulsion system. This propulsion system was designed for:

  • Orbital Plane Change (inclination)
  • Orbital Altitude Change

This is the first time that such engine will be tested in space. The propulsion system uses green fuel and has a maximal thrust of 2.5mN. The fuel volume is 1700 ml.

The thrust engine is based on the Magnetic Field Oscillating Amplified Thruster (MOA). This thruster is able to accelerated charged gases to extremely high velocities generating a high energetic plasma jet.

During the mission several manoeuvres will be achieved with the propulsion system. The exact positions during the manoeuvres will be tracked and registered on-board with the GNSS receivers and the raw data will be processed on the ground.

Figure 11: PRECURSOR Thrusters activation

Figure 12: Fuel tanks

Figure 13: Thruster magnetic field intensity

Figure 14: Plasma Test

Figure 15: Magnetic Momentum Test

Figure 16: Thrust Test in the TV Chamber

Figure 17: TV Chamber and Test Setup

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GNSS Receiver

A LimeSDR receiver will be used as GNSS receiver. The application software is based on open-source code.
The receivers have a high precision clock with following characteristics:

  • Frequency 10.000000MHz
  • Stability 0.1 - 0.5 ppm

Figure 18: LimeSDR Radio Receiver (Source: limemicro.com)

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GNSS Antenna

The GNSS antenna is a new generation antenna designed for small satellites. The antenna has following characteristics:

  • Size: 10x8.3x1 cm
  • Mass: 20g
  • Passband: 1160 - 1300 MHz and 1525 - 1610 MHz
  • Polarization: RHCP
  • Passive zenith gain: L1, E1, G1 >1.5 dBic and L5, L2, E6 >0 dBic
  • Passive horizon gain: >-7 dBic

Figure 19: PRECURSOR GNSS Antenna

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RADIATION SENSORS

In this experiment commercial transistors will be used as radiation sensors. The selected transistors are manufactured with standard low-power enhancement mode lateral pMOS technology. Designing a low-cost hand-held measurement system using this pMOS as the sensor would have a clear advantage due to the lower cost incurred by a standard technological process.

Figure 20: Transistor 3N163 as radiation sensor

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STRUCTURE

To protect the electronic inside against radiation a robust structure helps to reduce the radiation dose and extend the lifetime of the electronic parts.
The design of the structure is customized and compliant with mission requirements, where following types of duralumin will be used for the walls: 5052 H32 and 6061 T6. The solar cells will be attached to the structure sides. Small pieces will be printed and with PEEK material manufactured.

Figure 21: PRECURSOR Structure

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PAYLOAD CONTAINER DEPLOYER

For the satellite a customized deployer is designed. The deployer will be manufactured with Duralumin 7075 T6, and 3D printed parts.

Figure 22: PRECURSOR Deployer

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ENVIRONMENTAL TESTS

The payload has to absolve an acceptance qualification. In this process will be proofed that the payload can withstand the flight on the ZQ-2 without damaging other payloads.

Following environmental tests will be achieved prior the launch:

  • Vibration Sinusoidal
  • Vibration Random
  • Shock Test
  • Outgassing
  • Thermal Vacuum

Some tests will be achieved at System level or at module level.

 

·         VIBRATION

The major stress during the flight will be generated from the launch vehicle vibrations. It must be guaranteed that the payload will not be destroyed during the launch.

 

o   SINE VIBRATION

The table 6 lists the experimental conditions for low frequencies sine vibration from 

5Hz to 100Hz, which are used to simulate the transient and stationary random vibration of a launcher.

 

Table 6: Low Frequency Sine Vibration Scan Test Condition

Position

Frequency Range(Hz)

Acceptance Level

Identification Level

Payload Bracket

5~10

2mm

3mm

10~100

0.8g

1.2g

Scanning Rate R

4.0Oct/min

2.0Oct/min

 

o   RANDOM VIBRATION

The table 7 gives the test conditions of random vibration at high frequencies ranging from 20 Hz to 2000 Hz, which can be used to simulate the transient and stable random vibration of a rocket.

 

Table 7: Payload Interface Random Vibration Levels

Position

Frequency

Range

(Hz)

Acceptance Level

Identification Level

Power Spectral Density

(g2/Hz)

Duration

of Test Runs

(min)

Total Mean Square Root(g)

Power Spectral Density

(g2/Hz)

Duration

of Test Runs

(min)

Total Mean Square Root

(g)

Payload Bracket

20~150

3dB/Oct

1

6.94

3dB/Oct

2

10.41

150~800

0.04

0.09

800~2000

-6dB/Oct

-6dB/Oct

 

·         SCHOCK

The impact experiment condition (shock response spectrum Q=10) is shown as the table 8. The response acceleration time history in three vertical directions at the bottom of the specimen must be measured during the test.

 

Table 8: Shock Response Spectrum (Q=10)

Position

Frequency Range(Hz)

Acceptance Level

Identification Level

Satellite

50~1000

9dB/oct

9dB/oct

1000~8000

3000g

4500g

Semi-sine vibration test: peak value of the accelerated velocity is 200g;PW is  6ms with X direction; Repeat the test 3 times for each direction.

 

 

·         ACOUSTIC

Acoustic condition is detail in table 9.

 

Table 9: Payload Acoustic Environment

1/3 band

Central frequency (Hz)

1/3Sound Pressure Level / dB in band Width

Inside the Fairing

25

111

31.5

113

40

115

50

120

63

125

80

127

100

129

125

130

160

131

200

131

250

131

315

132

400

132

500

132

630

131

800

131

1000

129

1250

126

1600

123

2000

121

2500

120

3150

119

4000

117

5000

115

6300

114

8000

112

SPL

142

 

·         EMC

The Electromagnetic radiation for the radio equipment on the launch vehicle and also the launch site cannot exceed the requirement shown in the table10.

 

Table 10: EM radiation of the rocket and launch site

Frequency(MHz)

Intensity(dBμV/m)

0.01~0.05

80

0.05~3

90

3~30

70

30~550

65

550~750

103

750~2200

90

2200~2300

138

2300~1000

90

10000~

90

 

 

 

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GROUND SEGMENT

The ground segment consists in radio-amateur ground stations, which allow the effective operation of the mission. For commanding two stations will be used. For downlink the telemetry if open and free for the radio-amateur community. For the communication issues the project has the support from AMSAT-DL.
Principal capabilities and features of the ground segment:

  • All stations have the similar hardware and software configuration
  • Remote or autonomous (without staff) control
  • Collected data can be transferred between stations
  • Uplink with two station possible
  • Downlink for all support stations
  • The duration of one pass takes around 10 min
  • Number of contacts in average per day and per station is 4 in average
  • Data download in average of 703 KB (UHF) per contact

The core components of the PRECURSOR ground station are:

  • Transceiver ICOM IC-9100 (VHF and UHF)
  • Power supply for the transceiver PS-125.
  • Two pre amplifiers (one for VHF, other for UHF)
  • Antenna rotor Big-RAS
  • Two Yagi antennas, one for VHF (2 m), other for UHF (70 cm)
  • Cables and connectors
  • Personal computers with network adapter

Figure 23: AMSAT-DL 20m antenna in Bochum

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MISSION OPERATIONS

A group of 4 Operators (licenced radio amateurs) will be available for this project.
The mission operations team will manage the schedule of the satellite contacts, depending on the payload activities.
The mission operation prepared is based on the ESA standards for flight control organisation and will have the following phases:

Pre-Launch phase

Before satellite will be launched the team will be trained for all phases of the mission using a duplicate of the most important satellite parts. This twilling system will build a simulator helping to restage all the cases of the mission to operate the satellite. The team will be also trained to manage all routine activities and emergency cases. In this phase all necessary procedures and documents will be written and validated.

LEOP phase

During this critical phase the satellite will be in contact with all possible ground stations in order to initialise all the subsystems. Test transmissions with the ground stations and each function of the satellite will be checked. All the subsystems will be configured for the routine operation. In this phase the operations team will work with the bus only but not with the payloads. The operation service of 24 hours begins on shifts. The organization of the flight control team will be redundant following the ESOC typical organisation for a LEOP.

Commissioning phase

In this phase the operations team will work with the payloads. The payloads will be switched, checked and configured. After that the calibration of the instruments will be achieved. During this phase is necessary to be in contact with the instruments responsible in case of malfunctions.

Routine phase

During this phase the team will keep the 24 hours shift work but the activities will be reduced and for any anomaly an On-call service will be available. The satellite will keep the nominal mode and the instruments (payload) will collect science data. In the scheduled contacts the science data will be down-linked, stored and forwarded.

Recovery

In case of an emergency situation and the satellite is not able to go back to the nominal mode due to several errors the team has to execute emergency procedures and have to prepare the commands to recover the spacecraft.

 

For the PRECURSOR mission operations following activities are planned:

  • Daily downlink of the housekeeping telemetry
  • Daily downlink of the science data
  • Manoeuvres with the propulsion system
  • Experimental measurements with the GNSS receivers
  • Downlink of the GNSS measurements raw data
  • Post-processing of the raw data

 

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TEAM

Jaime            

 

Jaime Estela is an electronic engineer born in Lima-Peru. He worked at the German Space Operations Center of DLR in Oberpfaffenhofen for 11 years. In this period he gathered experience in satellite operations and systems engineering and supported several LEO orbit satellite missions like Terrasar-X, Tandem-X, Prisma A & B, Grace 1 & 2, CHAMP, BIRD, TET and was also involved, as Ground Segment Engineer, in the ESA project Columbus, space laboratory onboard the ISS,.

Furthermore he has supported CubeSat projects developed by Universities. He supported the project QB-50, an international constellation of 50 CubeSats which will study the higher ionosphere in a low Earth orbit and during its re-entry as suborbital research platform.

Jaime is a supporter of the NewSpace initiative and contributes to this space movement actively with new concepts, ideas, products and services.

 

 

Alex               

Dr. Alexander Popugaev received the M.S. degree (with honours) in radio engineering from the Vladimir State University, Russia, in 2004, and the Dr.-Ing. (Ph.D.) degree (summa cum laude) from the Technical University of Ilmenau, Germany, in 2013.

Since 2004, he has been with the Fraunhofer Institute for Integrated Circuits IIS, Germany.

During his over 15 years long tenure at the RF and Microwave Design Department, later RF and SatCom Systems Department, Dr. Popugaev worked his way up from Research Associate to Chief Scientist.

Due to his successful long-time R&D activities in the field of customer-specific GNSS antennas, Dr. Popugaev moved in 2020 into the Satellite Based Positioning Systems Department. His current position is Business Development Manager for GNSS Antennas.

Dr. Popugaev has authored and co-authored several scientific papers and book chapters and has received several patents.

 

 

Tobias                      

Tobias Bartusch is an electrical engineer since 2000 (FH Augsburg) and physicist (Univ. Augsburg) since 2006. He gained theoretical and practical experience in many space projects in different companies since 2006 (neutralizers and RF-multi staged magneto-plasma thrusters, boxes / modules for various customers, atomic clocks, timing systems and analog space design, AIT and Systems Engineering). In addition, he has developed high vacuum systems (tubes, electron emitters, plasma sources and measurement equipment) and vacuum chambers. He is experienced in the development of vacuum COTS equipment and space materials / high vacuum materials such as coatings for thermal radiators, AtOx resist, PTFE, and other. In the Precursor project, he is the responsible engineer for AIT, thermal design, deployment box, solar panels, charger, electrical thruster, some power converters, parts of the antenna integration and test. He is also a radio amateur (DH2MBT) and member of the DARC and AMSAT-DL. The satellite will be integrated in his workshop and laboratory which has the capability to perform all necessary steps.

 

Miguel Ángel Carvajal Rodríguez: Tenured Professor of the Department of Electronics and Computer Technology, CITIC-UGR, IMUDS-UGR, University of Granada, Granada, SPAIN

1637248571340

 

Alberto José Palma López: Full Professor of the Department of Electronics and Computer Technology, CITIC-UGR, University of Granada, Granada, SPAIN

Foto alberto-1

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PARTNERS

                                                                   

 

                                                                                                  

 

 

 

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SPONSORS

                Bartusch Plasma and Vacuum   

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