HADES: H1 Absorption in the Dark Ages
HADES mission aims to position a small satellite in advantageous positions in lunar orbit free from significant terrestrial RFI disturbances. Such positions present unique opportunities for neutral hydrogen measurements of the early universe. Collected data is downlinked to Earth at near side orbital positions.
2021 Participants
Sponsors + Partner Organizations
We are very thankful for the following organizations for making this summer’s program possible!
Proposal
HADES shall probe the structure of the Universe between the Dark Ages and the Cosmic Dawn. Lunar instrumentation for ultra-low radio frequency astronomy reduces terrestrial interference and facilitates observations of HI absorption between redshifts z~1000 to 10. This is the period between recombination and reionization for which cosmological models remain largely untested.
Measurements of the cosmological HI (21-cm) line require subtracting foregrounds that are up to ~6 orders of magnitude brighter.
Our scientific objectives target the following foregrounds:
- Galactic synchrotron emission
- Dust absorption
- Terrestrial/solar radio frequency interference (RFI)
The pointing requirements are driven by target threshold for Galactic foreground (~5000K @ 50 MHz). Best attenuation of terrestrial RFI at <100km above lunar surface.
The science instrument consists of a 3.3m dipole antenna, and LNA, Low Pass Filter (LPF), ADC, and FPGAs. The LPF shall filter out < 1 MHz. Lunar ionosphere plasma frequencies are 0.2 – 1 MHz. The ADC shall have a sampling rate of ~200 Mbps, with FPGAs for time-averaging, channeling data, and creation of the frequency spectrum.
Science Orbit
- 28.5 – 91.5 km altitude
- 0° – 2° inclination
- 0- 0.035 eccentricity
- Duration = 5 months
Spacecraft Design
Trajectory:
Continuous Low Thrust from GTO to Moon Sphere of Influence. Delta V = 3.5 km/s, Transfer Time = 600days- 4 flybys for Lunar Capture
- Continuous Low Thrust Burn for Lunar Descent into Science Orbit. Transfer Time = 135 days.
- Intentional crash land onto the surface of the moon for Disposal. Continuous Low Thrust trajectory. Delta v = 20 m/s, Duration = 3.7 days
Propulsion:
Electric propulsion closes the loop for first ever GTO to lunar transfer for a small satellite. Will set new delta V record for small satellites: 4.35km/s.
Two Enpulsion Micro R3 units were selected to meet the delta V and power requirements. 0.635 mN of thrust each.
Small form-factor cold gas thruster unit (VACCI CuSP MiSP) for attitude momentum dumping. 0.3U volume. Four 25 mN thrusters. 69.4 N-s total impulse.
Attitude Determination and Control System:
With a pointing requirement of 0.715° (driven by trajectory control), a robust ADCS design is proposed:
- Actuators: 4 Reaction Wheel Assembly for non-singular 3-axis active control
- Sensors: Star tracker, 3 gyros, Sun/horizon sensor
- EKF for attitude determination (all sensors on near-side; gyros + star tracker on far side)
- EKF for orbit determination (2-way doppler ranging using DSN + NEN on near side; propogated on far side)
Thermal:
MATLAB modelling & ANSYS simulation conducted on spacecraft. Spacecraft body temperatures range from -5° to 35° C. Passive thermal control via Alodine coating (9°C increase of minimum boy temperature). Active control through battery heater and 1 extra safety heater (autonomously controlled via thermistor inputs from each subsystem).
Communications:
Science data shall be downlinked using Near Earth Network (NEN). 1.6MB of data downlinked per day. Navigation data downlinked using DSN, transmitted 5 times per orbit on near-side at a rate of 128 kbps.
Selected hardware:
- JPL/Iris Transponder
- X-Band (8000-8500 MHz)
- 4W Output Power
- BPSK Modulation
- 12.6W/35W consumption (receive only/full transponder)
- AntDevCorp Patch Array
- 9B gain, 20° beamwidth
Power:
12 solar panels (NanoPower TSP) divided into two deployable arrays.
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- 2DOF Sun Tracking
- Solar Input (180W max / 150W avg)
Total consumption: 30.78W orbit-average.
Energy storage: NanoPower BP9 Li-ion Battery (86Wh)
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- Sized for avg. Depth of Discharge of 20%
BeCu Gaskets + Al Fraraday Cage selected for EMI shielding to meet the 80dB surpression required by the science instrument.
Command & Data Handling:
For mission assurance, redundant flight hardware is used for the on-board flight computer, along with a rad-hardened external clock. The SatBus 3C2 is selected, with storage capabilities up to 2 GB.
Structures:
The science instrument primarily consists of a 3.3m dipole antenna. This antenna along with all spacecraft subsystems can be arranged within the volume and mass requirements of a 6U CubeSat when stowed. The dipole antenna and solar panels are both deployable.
Software:
Mission operations are divided between the Science Phase on the dark side of the moon, and the Downlink Phase on the Earth-facing side. The flight software is divided into mission modes from initialization phases through fault detection and handling in safehold modes.
