UPMSat-2 Systems Engineering

Transcription

1 UPMSat-2 Systems Engineering Prof. Gustavo Alonso Prof. Sebastian Franchini Ali Ravanbakhsh Pablo Arriazu UPMSat-1 Launched in July 7th 1995 May 2011

2 Although UPMSat-2 is a University Class Microsatellite with special challenges and limitations, let's make it a TRUE SUCCESS, TOGETHER Systems engineering group objectives Adapting established practices of SE to this University Class Microsatellite (based on ECSS). Conducting necessary trade off studies in design driver subsystems. Decision making for design baselines. Controlling the design requirements. Controlling the interfaces between different subsystems. Necessary global document release for each phase of the project (MDD, PDD,SRD, )

3 Systems engineering for a satellite project is needed because: Structural design Strength Stiffness Orbital parameters Mission objectives Piggy back launch Life time On orbit performance Software design (C&DH, TT&C) Hardware electronics design Electrical power design Solar array Battery sizing Thermal control design Passive Active Attitude determination and control design Strategy Sensors Actuators

4 System engineering Near-term steps Deal with subsystems to investigate their requirements from other subsystems. Finalizing satellite dimension. Finalizing performance concept of different subsystems. Setting a power consumption profile according to subsystems data. Conducting a preliminary on orbit task schedule. Investigate the possible payloads for the satellite.

5 Geometry envelope UPMSat-2 geometry envelope should be adaptable to different auxiliary payload launch services (Versatile design). Survey on different lauchers such as: Ariane 5, Dnepr, PSLV and H-IIA Survey on microsatellites launched as an auxiliary payloads. Microsatellite Launcher Dimensions [mm] UPMSat-1 [2] Ariane ANUSAT [3] PSLV QSAT [4] H-IIA

6 Geometry envelope 500 mm 500 mm 600 mm {X,Y,Z} Architecture Cubic based parallelized sides 4 trays A Link to separation system Batteries Power control unit electronics B Electronic components (C&DH and TT&C boards) C Payload(s) D External sensors Comm. antennas

7 Subsystems mass budgets Satellite total mass: 50 Kg Subsystem Portion of Sat total mass Position Electrical Power 20 % 10 Kg Tray A Attitude Determination and Control Sys. 10 % 5 Kg Distributed Comm. boards 4 % 2 Kg Tray B Comm. antennas 1 % 0.5 Kg Tray D Onboard Data Handling 5 % 2.5 Kg Tray B Thermal Control 2 % 1 Kg Distributed Harness 3 % 1.5 Kg Distributed Payload min 30% 15 Kg Tray C Structure and payload Max 25% 12.5 Kg Distributed

8 Mission design (1) Sun-synchronous orbit near 600 km An earth orbit with curios property of providing a constant sun angle for the observation of Earth. Orbital inclination is nearly polar ( degrees) Global coverage Relative position of the SS-O remains almost fixed WRT the sun s direction. Lightning conditions will be the same during the mission along the foot print areas which must be observed during day light. SS-O could provide a continual dark-side for satellite in orbit (Thermal Control). Discrete altitudes can be selected to provide SS-Os ground tracks which repeat after a fixed interval of days.

9 Repeat Cycle (days) Inclination (deg) E.T.S. DE INGENIEROS AERONÁUTICOS Mission design (2) Sun-synchronous orbit near 600 km 99.5 A wide range of scenarios Altitude (km) A wide range of altitudes Altitude (km)

10 Mission design (3) Near-term steps Selecting a range of altitudes near 600 km for calculating parameters like: Eclipse time Revisit time Orbital Field of View (FOV) Distance between ground tracks Deciding about number and position of ground station(s). According to orbital characteristics and power-volume budgets verifying the possible payloads.